# Soft Inflatable Robotic Systems for Space Applications: A Survey

## Abstract

Soft inflatable robotic systems and structures are emerging as transformative technologies for space applications, offering compelling advantages in mass efficiency, compact stowage, compliance, and adaptability over traditional rigid-body systems. This survey provides a comprehensive review of the intersection of soft robotics, inflatable structures, and space engineering, organised around a unifying thesis: the same high-strength fabric technologies (Vectran, Kevlar, Nextel) that enable inflatable habitats also enable compliant debris capture mechanisms and large deployable shields. We examine two primary application domains---active debris removal, where soft compliant systems address the fragmentation paradox inherent in rigid capture, and space exploration, where inflatable habitats offer order-of-magnitude mass efficiency improvements over metallic modules. Eight enabling technology areas are reviewed: materials and structures, deployment mechanics, actuation, sensing and structural health monitoring, power systems, thermal management, attitude and orbit control, and robotic in-orbit assembly. We identify five critical research gaps, including the absence of quantitative soft-versus-rigid fragmentation comparisons, the lack of flight heritage for soft robotic capture, and the unexplored rigid-to-flexible assembly interface. A research roadmap spanning 5-year and 15-year horizons is proposed, with the most flight-ready near-term demonstrator identified as a gecko-adhesive gripper on an inflatable arm with fibre Bragg grating structural health monitoring. This survey differentiates itself from prior reviews in Progress in Aerospace Sciences by focusing specifically on soft and inflatable systems---a technology class not covered by existing reviews of rigid space robotics or contact/contactless debris removal.

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## Full Text

Soft Inatable Robotic Systems for Space Applications:
1

A Survey
2

3

Abstract
4

Soft inatable robotic systems and structures are emerging as transformative tech-
5

nologies for space applications, oering compelling advantages in mass eciency, com-
6

pact stowage, compliance, and adaptability over traditional rigid-body systems. This
7

survey provides a comprehensive review of the intersection of soft robotics, inatable
8

structures, and space engineering, organised around a unifying thesis: the same high-
9

strength fabric technologies (Vectran, Kevlar, Nextel) that enable inatable habitats
10

also enable compliant debris capture mechanisms and large deployable shields. We ex-
11

amine two primary application domainsactive debris removal, where soft compliant
12

systems address the fragmentation paradox inherent in rigid capture, and space explo-
13

ration, where inatable habitats oer order-of-magnitude mass eciency improvements
14

over metallic modules. Eight enabling technology areas are reviewed: materials and
15

structures, deployment mechanics, actuation, sensing and structural health monitoring,
16

power systems, thermal management, attitude and orbit control, and robotic in-orbit
17

assembly. We identify ve critical research gaps, including the absence of quantitative
18

soft-versus-rigid fragmentation comparisons, the lack of ight heritage for soft robotic
19

capture, and the unexplored rigid-to-exible assembly interface. A research roadmap
20

spanning 5-year and 15-year horizons is proposed, with the most ight-ready near-term
21

demonstrator identied as a gecko-adhesive gripper on an inatable arm with bre
22

Bragg grating structural health monitoring. This survey dierentiates itself from prior
23

reviews in Progress in Aerospace Sciences by focusing specically on soft and inatable
24

systemsa technology class not covered by existing reviews of rigid space robotics or
25

contact/contactless debris removal.
26

Contents
27

1
Introduction
4
28

2
The Case for Soft Inatables in Space
8
29

2.1
Space Debris Crisis and the Need for Active Removal . . . . . . . . . . . . .
8
30

2.2
Human Exploration Beyond LEO: The Habitat Challenge . . . . . . . . . . .
11
31

2.3
Unifying Thesis: Shared Fabric Technology Across Applications
. . . . . . .
12
32

3
Use Cases: Active Debris Removal
15
33

3.1
Rigid Capture Approaches and Fragmentation Risk . . . . . . . . . . . . . .
15
34

3.1.1
The Fragmentation Paradox . . . . . . . . . . . . . . . . . . . . . . .
16
35

3.2
Soft and Compliant Capture Mechanisms . . . . . . . . . . . . . . . . . . . .
17
36

3.2.1
Gecko-Inspired Dry Adhesive Grippers . . . . . . . . . . . . . . . . .
17
37

3.2.2
Dielectric Elastomer Minimum Energy Structure (DEMES) Grippers
19
38

3.2.3
Bistable and Passive Capture Grippers . . . . . . . . . . . . . . . . .
19
39

3.2.4
Thermally Qualied Soft Grippers . . . . . . . . . . . . . . . . . . . .
20
40

3.2.5
Inatable Robotic Arms for Capture
. . . . . . . . . . . . . . . . . .
20
41

3.2.6
INSIDeR: Net Capture with Inatable Deployment
. . . . . . . . . .
20
42

3.3
Inatable Debris Shields . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
23
43

4
Use Cases: Habitats and Exploration
23
44

4.1
Heritage Timeline: Echo to BEAM
. . . . . . . . . . . . . . . . . . . . . . .
24
45

4.1.1
Early Inatables: Echo and Volga (19601965) . . . . . . . . . . . . .
26
46

4.1.2
TransHab: Proving the Five-Layer Architecture (19972000) . . . . .
26
47

4.1.3
Genesis and BEAM: Orbital Validation (20062016+) . . . . . . . . .
27
48

4.2
Current Commercial Programs: LIFE, Orbital Reef, and Beyond . . . . . . .
27
49

4.2.1
Sierra Space LIFE
. . . . . . . . . . . . . . . . . . . . . . . . . . . .
27
50

4.2.2
Historical Context: B330 and Commercial Ecosystem Fragility . . . .
28
51

4.2.3
NextSTEP Competitive Landscape . . . . . . . . . . . . . . . . . . .
28
52

4.3
Future Concepts: Lunar Surface, Mars Transit, Planetary Entry . . . . . . .
28
53

4.3.1
Lunar Surface Habitats . . . . . . . . . . . . . . . . . . . . . . . . . .
28
54

4.3.2
Mars Transit and Surface Applications . . . . . . . . . . . . . . . . .
29
55

4.3.3
European Programmes . . . . . . . . . . . . . . . . . . . . . . . . . .
29
56

4.4
Radiation Shielding: The BEAM SPE Findings and Design Implications
. .
30
57

5
State of the Art: Materials and Structures
31
58

5.1
Space-Rated Fabrics: Vectran, Kevlar, Zylon, Nextel
. . . . . . . . . . . . .
31
59

5.2
Multi-Layer Shell Architecture . . . . . . . . . . . . . . . . . . . . . . . . . .
33
60

5.3
Rigidization Technologies . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
37
61

5.4
Environmental Degradation: AO, UV, Radiation, Creep . . . . . . . . . . . .
38
62

6
State of the Art: Deployment Mechanics
39
63

6.1
Fold Patterns and Packaging Eciency . . . . . . . . . . . . . . . . . . . . .
39
64

6.2
Ination Sequencing and Control
. . . . . . . . . . . . . . . . . . . . . . . .
39
65

6.3
Flight Heritage: InateSail, LOFTID, BEAM Deployment Lessons . . . . . .
41
66

6.4
Comparison with Rigid Deployable Alternatives . . . . . . . . . . . . . . . .
42
67

7
State of the Art: Actuation for Soft Space Systems
43
68

7.1
Dielectric Elastomer Actuators and DEMES . . . . . . . . . . . . . . . . . .
43
69

7.2
Vacuum-Gap Electrostatic Actuators: Vacuum as Enabler
. . . . . . . . . .
44
70

7.3
Ionic Electroactive Polymers: Space Tolerance Assessment
. . . . . . . . . .
45
71

7.4
Tendon-Driven Continuum Manipulators . . . . . . . . . . . . . . . . . . . .
45
72

7.5
Shape Memory Alloys for Deployment
. . . . . . . . . . . . . . . . . . . . .
46
73

7.6
Jamming in Vacuum: A Novel Opportunity
. . . . . . . . . . . . . . . . . .
46
74

7.7
Sealed Pneumatic Actuation in Space . . . . . . . . . . . . . . . . . . . . . .
48
75

7.8
Electroadhesion and Magnetic Actuation: Emerging Approaches . . . . . . .
48
76

8
State of the Art: Sensing and Structural Health Monitoring
49
77

8.1
Fibre Bragg Grating Sensors: From Proba-2 to Inatable Webbing . . . . . .
49
78

8.2
Multicore Fibre Optic Shape Sensing . . . . . . . . . . . . . . . . . . . . . .
52
79

8.3
Capacitive, Resistive, and Alternative Soft Sensors . . . . . . . . . . . . . . .
52
80

8.4
Distributed Fibre Optic Sensing: Rayleigh and Brillouin Scattering
. . . . .
53
81

8.5
Distributed Impact Detection
. . . . . . . . . . . . . . . . . . . . . . . . . .
53
82

9
State of the Art: Power Systems for Large Inatables
54
83

9.1
Flexible Solar Array Landscape: ROSA to Perovskite . . . . . . . . . . . . .
54
84

9.2
The Inatable-Power Integration Gap: PowerSphere and Beyond . . . . . . .
56
85

9.3
Energy Storage: Li-ion, RFC, and Mission-Dependent Selection
. . . . . . .
57
86

10 State of the Art: Thermal Management
58
87

10.1 Multi-Layer Insulation for Inatable Shells . . . . . . . . . . . . . . . . . . .
58
88

10.2 The JWST Sunshield as Deployable Thermal Barrier Precedent
. . . . . . .
59
89

10.3 Variable Emissivity Coatings and Smart Radiators . . . . . . . . . . . . . . .
60
90

10.4 Loop Heat Pipes for Deployed Structures . . . . . . . . . . . . . . . . . . . .
61
91

10.5 Phase Change Materials in Fabric Layers: The TRL 23 Gap . . . . . . . . .
61
92

11 State of the Art: Attitude and Orbit Control
62
93

11.1 Control-Structure Interaction for Flexible Spacecraft
. . . . . . . . . . . . .
63
94

11.2 Gyroelastic Body Theory and Distributed Momentum Management . . . . .
63
95

11.3 Drag Budget for 100 m-Class LEO Structures
. . . . . . . . . . . . . . . . .
64
96

11.4 The Missing Theory: AOCS for Pressure-Stabilised Membranes
. . . . . . .
66
97

12 State of the Art: Robotic In-Orbit Assembly
67
98

12.1 Assembly Robot Heritage and Current Programmes . . . . . . . . . . . . . .
68
99

12.2 Walking Robots for Large Structure Assembly: E-Walker . . . . . . . . . . .
68
100

12.3 The Rigid-to-Flexible Interface Gap . . . . . . . . . . . . . . . . . . . . . . .
69
101

12.4 Assembly-Enabled Inatable Platforms: Design Requirements
. . . . . . . .
70
102

13 Challenges, Open Questions, and Research Roadmap
71
103

13.1 Critical Research Gaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
71
104

13.2 Integration Challenges at System Level . . . . . . . . . . . . . . . . . . . . .
74
105

13.3 Proposed Research Roadmap: 5-Year and 15-Year Horizons . . . . . . . . . .
75
106

13.4 The Path to Flight Demonstration
. . . . . . . . . . . . . . . . . . . . . . .
79
107

14 Conclusions
80
108

1
Introduction
109

Two converging pressures threaten humanity's long-term access to and presence in space.
110

The rst is the accelerating degradation of the orbital environment: the low Earth orbit
111

(LEO) regime is increasingly populated with debris that endangers operational satellites,
112

whose services  from climate monitoring to navigation  underpin the global economy.
113

The second is the ambition for sustained human exploration beyond LEO, which demands
114

habitable volumes an order of magnitude larger than current metallic modules allow within
115

existing launch vehicle constraints. This survey argues that a single technology class 
116

soft inatable robotic systems based on high-strength technical fabrics  oers a coherent
117

engineering response to both challenges through a shared material and structural foundation.
118

The orbital debris environment has reached a critical threshold. The European Space
119

Agency's 2025 Space Environment Report records approximately 44,870 tracked objects,
120

with an estimated 54,000 objects larger than 10 cm, some 1.2 million objects between 1 and
121

10 cm, and an estimated 140 million fragments between 1 mm and 1 cm, totalling roughly
122

15,800 tonnes of mass in orbit ESA Space Debris Oce [2025]. The consequence is oper-
123

ational: SpaceX's Starlink constellation executed 144,404 collision avoidance manoeuvres
124

in the rst half of 2025 alone, a 65-fold increase relative to 2021 ESA Space Debris Oce
125

[2025]. Kessler and Cour-Palais identied in 1978 that mutual collision among catalogued
126

objects could generate a self-sustaining fragment cascade Kessler and Cour-Palais [1978],
127

and Liou and Johnson subsequently demonstrated with the LEGEND simulation suite that
128

the current LEO population is already gravitationally unstable: even with a complete halt to
129

new launches, the debris environment continues to grow through inter-object collisions Liou
130

and Johnson [2006, 2008]. Stabilising LEO requires the active removal of at least ve large,
131

rocket-body-class objects per year from the most critical orbital shells Liou et al. [2010].
132

Active Debris Removal (ADR) therefore transitions from a conceptual aspiration to an
133

operational necessity. Yet the dominant design paradigm  rigid robotic arms similar to
134

ClearSpace-1's four-arm capturing system  carries an ironic risk: forceful contact with a
135

tumbling, uncooperative object can fracture it, generating new fragments faster than they
136

are removed. Simulation studies and ground tests indicate that peak joint torques of order
137

195 Nm can arise during ENVISAT-class capture operations Ledkov and Aslanov [2022],
138

and the RemoveDebris harpoon demonstration saw a carbon-bre boom snap on contact at
139

20 m/s Aglietti et al. [2020]. The fragmentation paradox  rigid capture risks accelerating
140

the very cascade it aims to halt  provides the primary motivation for compliant, soft
141

capture architectures.
142

Simultaneously, the ambition to sustain human presence beyond LEO confronts a fun-
143

damental mass budget constraint. Metallic pressurised modules  Columbus (137 kg/m3)
144

and Tranquility (205 kg/m3)  are delivered at densities an order of magnitude higher than
145

fabric-based alternatives such as the TransHab concept (39 kg/m3) Valle et al. [2019a]. Vec-
146

tran high-tenacity yarn achieves a specic strength of 2,330 kN-m/kg, versus 220 kN-m/kg
147

for Ti-6Al-4V Valle et al. [2019a]  a 10× advantage that directly translates to launch mass
148

savings. The Bigelow Expandable Activity Module (BEAM), attached to the International
149

Space Station (ISS) since 2016, has accumulated more than eight years of continuous pres-
150

surised operation on the ISS, with periodic crew access for inspection and cargo storage, at
151

Technology Readiness Level (TRL) 9 NASA Johnson Space Center [2017].
152

The organising thesis of this survey is that the same high-strength fabric technology
153

 Vectran restraint layers, Kevlar/Nextel debris shielding, Kapton thermal insulation 
154

that enables BEAM's pressure vessel integrity also enables compliant robotic capture arms,
155

large deployable debris shields, and the next generation of deep-space habitats. Material
156

qualication campaigns, manufacturing processes, and design heritage are shared across
157

these application domains, providing an unusually coherent pathway from current ight-
158

proven technology to future operational systems.
159

Scope and Organisation
160

This survey reviews the intersection of three mature elds: soft robotics, inatable space
161

structures, and the enabling subsystem technologies (materials, power, thermal manage-
162

ment, attitude and orbit control, and robotic assembly) that together determine whether
163

soft inatable systems can be realised at mission-operational scale. The scope spans two
164

primary application domains:
165

1. Active Debris Removal  soft and compliant capture mechanisms (TRL 25) and
166

large inatable debris shields (design stage), examined against the rigid-capture base-
167

line.
168

2. Human Space Exploration  the heritage from Echo 1 (1960) through BEAM
169

(2016+) to current commercial programmes (Sierra Space LIFE, Orbital Reef), and
170

future concepts for lunar surface, Mars transit, and planetary entry decelerators.
171

Eight enabling technology areas are reviewed in depth: (1) materials and structures,
172

(2) deployment mechanics, (3) actuation, (4) sensing and structural health monitoring,
173

(5) power systems, (6) thermal management, (7) attitude and orbit control, and (8) robotic
174

in-orbit assembly. The survey concludes with a consolidated gap analysis and a research
175

roadmap spanning 5-year and 15-year horizons.
176

Relationship to Existing Reviews
177


![Table 1](paper-23-v1_images/table_1.png)
*Table 1*

Three prior surveys in Progress in Aerospace Sciences address adjacent territory, and this
178

survey is positioned explicitly as their complement (Table 1). Flores-Abad et al. reviewed the
179

state of space robotics for on-orbit servicing in 2014 Flores-Abad et al. [2014], establishing
180

the four-phase capture framework (approach, tracking, capture, post-capture stabilisation)
181

that remains the standard reference; however, that work predates the current wave of soft
182

robotics innovation and does not address inatable structures. Ledkov and Aslanov surveyed
183

contact and contactless ADR approaches in 2022 Ledkov and Aslanov [2022], providing com-
184

prehensive coverage of nets, harpoons, ion beam shepherds, and electrodynamic tethers, but
185

soft and compliant capture mechanisms receive minimal treatment and inatable structures
186

for ADR are absent. Zhao et al. reviewed rigid robotic manipulators for in-orbit servicing and
187

ADR in 2024 Zhao et al. [2024], covering Denavit-Hartenberg kinematics, impedance control,
188

and comparative arm performance; soft and inatable manipulators are outside scope.
189

The most relevant prior survey is Zhang et al. (2023), who examined soft robotics for
190

space across actuation, sensing, and manipulation Zhang et al. [2023a]. That work identies
191

vacuum as a challenge for pneumatic actuation and catalogues the soft gripper landscape;
192

however, it does not cover the inatable structure platform on which soft robots operate, nor
193

the enabling subsystems (power, thermal, AOCS, assembly) necessary for mission viability,
194

nor the dual ADR-and-exploration organising principle developed here.
195

The unique contribution of this survey is threefold. First, it covers eight enabling tech-
196

nology areas through a single integrative lens, rather than the one or two areas addressed
197

by prior reviews. Second, it presents the rst unied treatment of both ADR and explo-
198

ration applications as manifestations of the same fabric-based technology class. Third, it
199

maps cross-domain connections  between, for example, thermal management and actuator
200

design, or fold patterns and debris protection  that can only be identied from a broad
201


![Table 2](paper-23-v1_images/table_2.png)
*Table 2*

survey perspective.
202

Table 1: Comparison of this survey with prior reviews in Progress in Aerospace Sciences
covering adjacent domains. ✓= covered;  = not covered; ∼= partial coverage.

Topic
This survey
Zhao 2024
Ledkov 2022
Flores-Abad 2014

Soft/compliant capture
✓

∼

Inatable robotic arms
✓



Inatable debris shields
✓



Inatable habitats
✓



Rigid ADR approaches
∼
✓
✓
✓
Rigid manipulators
∼
✓
∼
✓
Materials & fabrics
✓



Power systems
✓



Thermal management
✓



AOCS for large structures
✓



Robotic in-orbit assembly
✓
∼

∼
Sensing & SHM
✓



Deployment mechanics
✓




Year
2026
2024
2022
2014
Soft/inatable focus
Primary
None
Minimal
None

The Paradigm Shift: Vacuum as Design Resource
203

A recurring theme throughout this survey is the inversion of the conventional assumption
204

that space vacuum is hostile to soft robotic systems. Three independent developments chal-
205

lenge this assumption. First, Sirbu et al. demonstrated vacuum-gap electrostatic multilayer
206

actuators in 2025 that require vacuum to function: thin-lm polymer multilayers with inter-
207

nal vacuum gaps zip closed on electrical activation, producing forces exceeding 4 N from a
208

0.7 g actuator at bandwidths above 100 Hz Sirbu et al. [2025]. On Earth, a vacuum pump
209

would be required to create this operating condition; in space, the environment provides it
210

at no mass or power cost. Second, the conning pressure for granular and layer jamming 
211

which terrestrially requires evacuating a sealed membrane with a pump  is provided for
212

free by the ambient vacuum dierential against a pressurised inatable interior Fitzgerald
213

et al. [2020]. Third, DEMES gripper geometry provides a passive negative feedback loop
214

in microgravity: grip force increases as a oating target drifts away from the actuator tip,
215

oering passive capture stability without active control  a property that is useful only in
216

the microgravity environment Araromi et al. [2015].
217

These developments suggest that soft inatable robotic systems are not merely terrestrial
218

technology adapted for space, but a distinct engineering discipline with unique environment-
219

enabled advantages.
220

Review Methodology
221

The literature for this survey was assembled through a structured search strategy span-
222

ning multiple databases and source types.
Primary databases searched include Scopus,
223

Web of Science, NASA Technical Reports Server (NTRS), ESA's publication repository, and
224

Google Scholar, using the following search term families: (i) inatable space structure
225

OR expandable habitat OR deployable membrane; (ii) soft robot* AND space OR
226

orbital; (iii) active debris removal AND (compliant OR soft OR inatable); and
227

(iv) technology-specic terms for each of the eight enabling areas (e.g., dielectric elastomer
228

actuator space, bre Bragg grating spacecraft, perovskite solar cell radiation). The tem-
229

poral scope spans 1960 (Project Echo) to early 2026, with no lower date restriction applied.
230

Inclusion criteria required that sources address at least one of the two application domains
231

(ADR or exploration) or one of the eight enabling technology areas in a space-relevant con-
232

text. Conference proceedings were included when they represented the primary publication
233

venue for mission results (e.g., AIAA, IAC, IEEE Aerospace). NASA technical memoranda,
234

ESA reports, and agency mission documentation were included for heritage programme data
235

not available in peer-reviewed form. Corporate press releases and datasheets were included
236

only when no peer-reviewed alternative existed for specic mission or material property
237

data. The eight technology areas were selected based on a preliminary scoping review that
238

identied all subsystem-level capabilities required for an operational soft inatable robotic
239

system at mission scale, following the principle that reviews in Progress in Aerospace Sci-
240

ences should enable the reader to assess system-level feasibility rather than component-level
241

performance alone. TRL assessments throughout the paper follow the NASA NPR 7123.1B
242

standard denitions NASA [2020].
243

Survey Statistics
244

This survey reviews approximately 120 primary sources spanning the period from 1960 to
245

2026.
Of these, approximately 74% are peer-reviewed journal papers or conference pro-
246

ceedings from indexed venues; the remainder comprises NASA technical memoranda, ESA
247

reports, and agency mission documentation. Coverage extends across eight technology areas
248

and two application domains, with the deepest literature pools in actuation (Zhang 2023 and
249

its references), inatable habitats (Litteken 2019 and the TransHab programme), and space
250

debris (Kessler 1978 through ESA 2025). The survey is organised with application use cases
251

preceding the technology state-of-the-art review, following the principle that applications
252

should motivate the technology landscape rather than the reverse.
253

2
The Case for Soft Inatables in Space
254

2.1
Space Debris Crisis and the Need for Active Removal
255

The accumulation of orbital debris is the dening environmental challenge of the space
256

age. Since Sputnik-1's launch in 1957, every mission has contributed to a growing cloud of
257

defunct satellites, spent rocket stages, and collision fragments. The debris environment is
258

now characterised not merely by nuisance but by irreversible instability.
259

Current Debris Environment
260

The ESA Space Environment Report for 2025 provides the most current comprehensive
261

characterisation ESA Space Debris Oce [2025]. As of early 2026, approximately 44,870
262

objects are tracked by ground-based surveillance networks, of which roughly one third are
263

operational satellites and two thirds are debris. The total catalogued population has grown
264

by more than 3,000 objects from fragmentation events in 2024 alone. At altitudes between
265

500 and 700 km  where ADR missions are most urgently needed  debris density is
266


![Table 3](paper-23-v1_images/table_3.png)
*Table 3*

comparable to or exceeds the density of active satellites.
267

Table 2: Current LEO debris population by size category (data from ESA Space Environment
Report 2025 ESA Space Debris Oce [2025]).

Size category
Estimated count
Trackable?
Primary threat

> 10 cm
∼54,000
Yes (radar)
Catastrophic collision
110 cm
∼1,200,000
No
Mission-ending damage
1 mm  1 cm
∼140,000,000
No
Surface/solar panel damage
< 1 mm
> 1012
No
Erosion/coating damage

Total mass
∼15,800 tonnes



More than 650 fragmentation events have occurred in orbit since 1961, with signicant
268

contributors including the 2007 Chinese ASAT test (Fengyun-1C), the 2009 Cosmos-Iridium
269

collision, and the 2021 Russian ASAT test (Cosmos-1408). These events collectively added
270

thousands of trackable fragments and orders of magnitude more sub-centimetre particles.
271

The Kessler Syndrome: From Prediction to Conrmation
272

Kessler and Cour-Palais (1978) predicted that beyond a critical debris density, mutual col-
273

lisions among catalogued objects would generate fragments faster than atmospheric drag
274

could remove them, leading to an exponential growth cascade now called the Kessler syn-
275

drome Kessler and Cour-Palais [1978]. For nearly three decades this remained a theoretical
276

concern. Liou and Johnson (2006) demonstrated with the LEGEND orbital debris evolution
277

model that the predicted threshold has already been crossed in the 8001000 km altitude
278

band: even if all future launches were halted immediately, the debris population in these
279

shells would continue to grow due to existing collision rates among currently catalogued
280

objects Liou and Johnson [2006]. Extended 200-year projections (Liou and Johnson 2008)
281


![Table 4](paper-23-v1_images/table_4.png)
*Table 4*

80,000

Projected (no ADR)

Projected (5 ADR/yr)

70,000

Number of catalogued objects in orbit

60,000

50,000

Mega-constellation

era begins
Kessler & Cour-Palais

(1978) prediction
ESA 2025:
44,870 tracked
(~54,000 est. >10 cm)

40,000

30,000

India ASAT
(Mission Shakti)

China ASAT
(Fengyun-1C)

Cosmos-Iridium

collision

20,000

10,000

0

1960
1970
1980
1990
2000
2010
2020
2030
2040
Year

Figure 1: Growth of the catalogued orbital debris population from 1960 to 2025, with projec-
tions to 2040. Discrete fragmentation events (Chinese ASAT 2007, Cosmos-Iridium collision
2009) are visible as step increases. Red dashed line: projected growth without active de-
bris removal. Green dashed line: projected stabilisation with ve large-object removals per
year Liou et al. [2010]. Data from ESA Space Environment Report 2025 ESA Space Debris
Oce [2025].

conrmed that the instability is neither transient nor recoverable without active interven-
282

tion Liou and Johnson [2008].
283

The required rate of removal has been quantied. Liou et al. (2010) showed that removing
284

at least ve large objects per year (primarily rocket bodies in the 8001000 km band) is nec-
285

essary and sucient to stabilise the LEO population over a 200-year projection horizon Liou
286

et al. [2010]. This represents an annual ADR cadence comparable to the total number of sig-
287

nicant deorbit missions conducted globally over the past decade  a formidable operational
288

challenge.
289

The Fragmentation Paradox
290

The dominant design approach to ADR  rigid robotic arms, exemplied by ESA's ClearSpace-
291

1 mission targeting the PROBA-1 satellite  faces a fundamental tension. Rigid contact
292

with a non-cooperative, tumbling debris object generates impulsive forces at the contact
293

interface. For an 8-tonne ENVISAT-class object rotating at 5 deg/s, e.deorbit trajectory
294

analyses reveal peak joint torques of 195 Nm at structural limits Ledkov and Aslanov [2022],
295

while experimental harpoon tests in the RemoveDebris mission saw a carbon-bre deploy-
296

able boom snap on contact with the capture target at 20 m/s Aglietti et al. [2020]. Wang
297

et al. (2023) note explicitly that rigid manipulation has the potential to generate fragments
298

during the capturing phase Wang et al. [2023], and Chen et al. (2024) characterise single
299

contact-based caging approaches as excessively risky for fast-tumbling targets Chen et al.
300

[2024a].
301

This fragmentation paradox is quantiable in energetic terms. The NASA/ESA IMPACT
302

model identies a catastrophic fragmentation threshold of 10 J/g of specic energy at the con-
303

tact interface Liou and Johnson [2006]. A 100-kg debris object rotating at ω = 5 deg/s and
304

2Iω2 ≈sev-
305

grasped rigidly at a moment arm of 0.5 m experiences a contact energy of order 1

eral hundred joules at the grasp point. If this energy is absorbed by the contact structure
306

rather than dissipated, the resulting specic energy may approach or exceed the fragmen-
307

tation threshold for lightweight aluminium honeycomb solar panel structures. No published
308

paper has conducted a systematic quantitative comparison of fragment generation probabil-
309

ity between rigid and compliant capture mechanisms  this gap is identied as a priority
310

experimental question in Section 13.
311

Compliant and soft capture systems address the paradox by absorbing and redistributing
312

contact energy rather than transmitting impulsive forces. Eight distinct soft and compliant
313

capture approaches are reviewed in Section 3, ranging from gecko-inspired dry adhesives
314

(microgravity-validated at TRL 45 Jiang et al. [2017]) to DEMES grippers with mission
315

heritage on CleanSpace One Araromi et al. [2015] and inatable robotic arms Palmieri et al.
316

[2023]. None has yet demonstrated in-ight capture, establishing a clear technology gap that
317

motivates the investment in ight demonstration infrastructure discussed in Section 13.
318

Operational Consequences
319

The operational burden of the debris environment is no longer theoretical. At 550 km altitude
320

 the operating shell of many Starlink satellites  the trackable debris density is sucient
321

to require avoidance manoeuvres at a rate that consumes propellant reserves and interrupts
322

normal operations. Starlink's 144,404 avoidance manoeuvres in H1 2025 (65-fold increase
323

from 2021 ESA Space Debris Oce [2025]) represent a structural operational cost that scales
324

with constellation size. ESA's own operational satellites execute hundreds of manoeuvres
325

annually, with collision avoidance emerging as a primary mission-operations driver. The
326

economic externality  uncontrolled debris imposes avoidance costs on all operators 
327

provides a market-failure argument for policy-mandated ADR that is increasingly reected
328

in international guidelines Liou et al. [2010].
329

2.2
Human Exploration Beyond LEO: The Habitat Challenge
330

The second driver for soft inatable systems is the ambition for sustained human presence
331

beyond the ISS. NASA's Artemis programme, ESA's Moon Village concept, and private
332

ventures such as Orbital Reef collectively assume that humans will occupy permanent or
333

semi-permanent outposts in cislunar space, on the lunar surface, in Mars transit, and even-
334

tually on the Martian surface. All of these scenarios require pressurised habitable volumes
335

substantially larger than any single rigid module that can be launched within existing fairing
336

constraints.
337

The Mass and Volume Eciency Argument
338

Valle et al. (2019) provide the denitive comparative analysis of inatable versus metallic
339

pressurised structures Valle et al. [2019a]. The key metric is areal density (mass per unit
340


![Table 5](paper-23-v1_images/table_5.png)
*Table 5*

oor area, or equivalently, mass per unit pressurised volume):
341

Table 3: Mass eciency comparison of representative pressurised space modules (adapted
from Valle et al. 2019 Valle et al. [2019a]).

Module
Type
Press. Vol. (m3)
Mass (kg)
Density (kg/m3)

TransHab concept
Inatable
339
13,200
39
BEAM (as-built)
Inatable
16
1,415
88
Columbus (ESA)
Metallic
75
10,300
137
Tranquility (Node 3)
Metallic
74
15,200
205

The mass eciency advantage derives directly from material specic strength. Vectran
342

HT, the primary restraint-layer fabric in BEAM and TransHab, has a tensile strength of
343

3.0 GPa at a density of 1.40 g/cm3, yielding a specic strength of 2,330 kN-m/kg Valle
344

et al. [2019a]. Kevlar 49, similarly used for restraint and micrometeoroid and orbital debris
345

(MMOD) protection, achieves approximately 2,080 kN-m/kg at the fabric level (3.0 GPa
346

UTS, 1.44 g/cm3 density) or 2,500 kN-m/kg at the lament level (3.6 GPa UTS) DuPont
347

[2019]. These compare to Ti-6Al-4V at 220 kN-m/kg and aluminium 7075-T6 at 204 kN-
348

m/kg: the fabric advantage is approximately one order of magnitude. This dierence directly
349

determines what pressurised volume can be delivered per kilogram of launch mass, and
350

therefore what human presence scenarios are economically feasible.
351

The volumetric launch eciency is equally compelling. A 300 m3 pressurised module at
352

metallic density would mass approximately 40,000 kg  exceeding the cargo capacity of any
353

current or planned launch vehicle for a single module. The Sierra Space LIFE 285 habitat,
354

targeting approximately 300 m3 of pressurised volume, folds into a fairing-compatible package
355

and deploys on orbit, representing a volume achievable in a single launch that has no metallic-
356

module equivalent Sierra Space Corporation [2024].
357

BEAM as Technology Proof
358

The BEAM module, delivered to the ISS by SpaceX CRS-8 in April 2016 and expanded
359

in May 2016, constitutes the highest-TRL demonstration of crewed inatable space struc-
360

tures NASA Johnson Space Center [2017]. BEAM provides 16 m3 of pressurised volume at
361

a deployed mass of 1,415 kg and has maintained pressure integrity for more than eight years
362

without rigidisation. Operational experience includes periodic crew access for inspection and
363

equipment storage, structural health monitoring via embedded accelerometers and impact
364

detection systems, and characterisation of the thermal, radiation, and MMOD environment.
365

BEAM's deployment was not without diculty: initial expansion attempts on 28 May
366

2016 required 25 pressurisation bursts over approximately seven hours to overcome friction
367

between compressed softgoods layers, compared to the planned single-burst expansion. This
368

experience provided critical engineering data on fold-compression set and deployment relia-
369

bility that directly informs the design of future autonomous deployment systems. Kennedy
370

(2002) documents the TransHab programme's prior exploration of this challenge, including
371

burst pressure tests to 4× operating pressure and the critical importance of restraint-layer
372

preloading for deployment force prediction Kennedy [2002].
373

Radiation: The Honest Assessment
374

BEAM data from the September 2017 solar particle event (SPE) revealed a critical nding
375

that must be stated clearly NASA Johnson Space Center [2017]. Absorbed dose measure-
376

ments in BEAM during the SPE were approximately 22.5 mGy, compared to approximately
377

0.25 mGy measured simultaneously in adjacent metallic ISS habitable volumes  an 810×
378

ratio. This nding demonstrates that fabric walls alone provide substantially less radiation
379

shielding than the aluminium walls of conventional modules.
380

This is not a disqualifying result, but it is a design constraint. The TransHab architecture
381

addressed this through a water-wall concept: a ∼10 cm thick water reservoir integrated into
382

the inner wall layers that provides both radiation shielding (hydrogen-rich material) and
383

useful crew water storage.
Norbury et al. (2025) review passive shielding materials for
384

space and conrm that polyethylene achieves a 27.8% mass saving relative to aluminium
385

for equivalent proton shielding at the same areal density Norbury et al. [2025]. The design
386

solution is established; its implementation requires deliberate integration rather than passive
387

reliance on wall thickness.
388

2.3
Unifying Thesis: Shared Fabric Technology Across Applications
389

The central organising principle of this survey is that the high-strength fabric technology
390

enabling inatable habitats is the same technology enabling compliant ADR capture arms,
391

large deployable debris shields, and the soft robotic systems operating within and around
392

both. This material unity has engineering consequences that extend beyond mere analogy.
393


![Table 6](paper-23-v1_images/table_6.png)
*Table 6*

Material Traceability Across Applications
394

Table 4 maps the four primary fabric families across their roles in dierent application do-
395

mains. The key observation is that the same material qualication data  creep behaviour,
396

AO erosion yield, UV degradation rate, thermal cycling tolerance  is relevant across all
397

applications. A Vectran creep characterisation campaign conducted for habitat restraint-
398

layer lifetime prediction Weadon [2013] is directly applicable to Vectran inatable robotic
399

arm links Palmieri et al. [2023].
A Nextel/Kevlar debris shield hypervelocity test cam-
400

paign Destefanis et al. [2003] produces data applicable to both habitat MMOD protection
401


![Table 7](paper-23-v1_images/table_7.png)
*Table 7*

and inatable debris shield design Cha et al. [2024].
402

Table 4: Shared fabric technology across application domains. The same material families
serve multiple functions, sharing qualication heritage and manufacturing processes.

Material
Habitat role
ADR role
Robotic
arm
role

Vectran HT
Restraint
layer
(primary load)
Inatable
arm
links
Inatable
ma-
nipulator links
Kevlar 49
MMOD
rear
wall;
restraint
co-layer

Net
tether;
shield backing
Arm outer jacket

Nextel 440
MMOD bumper
(ceramic)
Debris
shield
bumper layer


Kapton/Mylar
MLI outer lay-
ers; bladder liner
Shield
thermal
layer
Bladder
inner
liner
Beta cloth
AO-resistant
outer cover

AO-resistant
cover

The Mars Airbag Precedent
403

Vectran's role in the Mars Pathnder (1997), Mars Exploration Rover (2004), and subse-
404

quent airbag systems provides heritage that extends beyond Earth orbit. These missions
405

demonstrated that Vectran-based inatable structures can survive the combined stresses of
406

launch vibration, interplanetary cruise, hypervelocity atmospheric entry, and impact landing
407

on an extraterrestrial surface Litteken [2019]. The qualication data base thus spans not
408

merely LEO but the full range of conditions relevant to deep space exploration  a heritage
409

directly relevant to future Mars transit habitat designs.
410

Origami Geometry Unies Packaging and Protection
411

A particularly striking example of cross-domain material unication is the Inatable Modular
412

Space Shield (IMSS) proposed by Cha et al. (2024) Cha et al. [2024]. The IMSS uses a wa-
413

terbomb origami tessellation to fold a multi-layer ultra-high-molecular-weight polyethylene
414

(UHMWPE)/Kevlar/Nextel shield into a package achieving 90% volume reduction relative
415

to a rigid Whipple shield of equivalent protection.
The same Miura-ori and waterbomb
416

fold patterns Miura [1985] used in IMSS for debris shield deployment are the canonical fold
417

patterns for large membrane space structures generally Schenk et al. [2014]  packaging
418

eciency and multi-shock protection are simultaneously optimised by the same tessellation
419

geometry.
420

Scale-Dependent Challenges
421

While the material foundation is shared, the engineering challenges depend strongly on scale.
422

The scale-dependent challenge landscape can be summarised as follows: at centimetre scale
423

(soft gripper ngers), actuation force and contact compliance dominate the design; at metre
424

scale (inatable arms, BEAM-class habitats), deployment mechanics and pressure-retention
425

integrity dominate; at 10-metre scale (large solar concentrators, small debris shields), control-
426

structure interaction begins to matter; at 100-metre scale (large debris shields, solar power
427

collectors), attitude and orbit control, aerodynamic drag compensation, power generation,
428

and thermal management become the primary engineering challenges, for which no ight
429

heritage exists.
430

This survey is organised to trace the technology from its best-proven applications (TRL 9
431

materials, TRL 9 BEAM habitat, TRL 89 rigid solar arrays) through to the most speculative
432

future capabilities (TRL 23 pressure-stabilised membrane AOCS, TRL 34 vacuum-gap
433

actuation), making explicit at each stage what is demonstrated, what is extrapolated, and
434

what requires new research.
435

Why Soft? Why Inatable? Why Now?
436

Three converging developments make this survey timely.
437

Material advances. Vectran and Kevlar have matured to TRL 9 in space environments.
438

Perovskite/CIGS tandem solar cells, demonstrated at 2,100 W/kg with 85% proton radia-
439

tion retention after equivalent 50-year LEO exposure Lang et al. [2020], promise to integrate
440

power generation into inatable membrane layers at specic powers unachievable with con-
441

ventional rigid panels. Cryogenic metallic cable-based soft robots (Foster-Hall et al. 2025)
442

maintain full range of motion at −196 ◦C, solving the elastomer embrittlement problem for
443

deep-space applications Foster-Hall et al. [2025].
444

Mission context. The commercial station era (Orbital Reef, Axiom, LIFE, Starlab) cre-
445

ates the rst sustained market demand for habitable volume beyond ISS. ESA's ClearSpace-1
446

mission, targeting PROBA-1 for retrieval in the late 2020s, establishes ADR as an opera-
447

tional rather than experimental activity. The convergence of launch cost reduction (SpaceX
448

Falcon 9, Starship) with mission demand means that the technology development cost of
449

inatable systems is now justiable against a credible mission pull.
450

Paradigm shift.
As outlined in Section 1, the space environment is increasingly un-
451

derstood as a resource for soft robotic systems rather than an obstacle. Vacuum-gap ac-
452

tuation Sirbu et al. [2025], jamming without pumps Fitzgerald et al. [2020], and passive
453

microgravity compliance Araromi et al. [2015] represent a qualitative shift in what the space
454

environment enables. This survey maps these opportunities systematically across the full
455

technology stack.
456

The following sections develop the application use cases (Sections 3 and 4) before re-
457

viewing the enabling technology state-of-the-art (Sections 512), and concluding with a
458

consolidated gap analysis and research roadmap (Section 13).
459

3
Use Cases: Active Debris Removal
460

The orbital debris environmentcharacterised in Section 2.1represents the most urgent
461

operational motivation for soft inatable robotic systems in space. With over 54,000 esti-
462

mated objects larger than 10 cm, 15,800 tonnes of total orbital mass, and a 65-fold increase
463

in Starlink collision avoidance manoeuvres since 2021 ESA Space Debris Oce [2025], the
464

operational urgency is undeniable.
465

The scientic foundation for active debris removal (ADR) was established by Kessler and
466

Cour-Palais Kessler and Cour-Palais [1978], who developed the rst mathematical model pre-
467

dicting cascading collisional fragmentation in low Earth orbit (LEO). Their analysis identied
468

three debris population regimesstable, critical, and cascadingand predicted the forma-
469

tion of a debris belt within a century. Subsequent Monte Carlo simulations by Liou and
470

Johnson Liou and Johnson [2006, 2008] using the NASA LEGEND model with 200-year pro-
471

jections across 50 runs demonstrated that the LEO debris population had already crossed
472

the instability threshold: the number of objects would continue to grow even with zero future
473

launches. Their work quantied the minimum intervention rate, establishing that at least
474

ve large objects per year must be removed from the 8001000 km altitude bands to stabilise
475

the environment Liou et al. [2010]. At approximately 550 km altitude, debris spatial density
476

now equals active satellite densityan unprecedented situation that fundamentally changes
477

the risk calculus for orbital operations ESA Space Debris Oce [2025].
478

This section examines the role of soft and inatable systems in addressing the debris
479

challenge. We rst review conventional rigid capture approaches and their inherent fragmen-
480

tation risk (Section 3.1), then survey eight distinct soft and compliant capture mechanisms
481

(Section 3.2), and nally discuss inatable debris shields as passive protection infrastructure
482

(Section 3.3).
483

3.1
Rigid Capture Approaches and Fragmentation Risk
484

Active debris removal using rigid robotic manipulators has been the dominant paradigm in
485

mission planning for the past two decades. Zhao et al. Zhao et al. [2024] provide the most
486

recent comprehensive review in Progress in Aerospace Sciences of rigid manipulators for on-
487

orbit servicing and ADR, covering ight-heritage systems such as the Canadarm and the
488

European Robotic Arm (ERA), cancelled missions including ESA's e.deorbit, and planned
489

missions such as ClearSpace-1. The review documents the extensive engineering heritage of
490

rigid robotic arms but also explicitly acknowledges the potential for fragmentation generation
491

during debris capture Zhao et al. [2024].
492

Ledkov and Aslanov Ledkov and Aslanov [2022] survey the full spectrum of ADR meth-
493

ods in Progress in Aerospace Sciences, including nets, harpoons, robotic arms, tentacles, ion
494

beam shepherding, laser ablation, electrostatic tractors, and electrodynamic tethers. Their
495

analysis notes that contactless methods such as ion beam shepherdingcapable of deorbit-
496

ing a 2-tonne debris object in 34 monthscarry zero mechanical impact risk, but require
497

extended proximity operations and signicant power budgets. Contact-based methods, while
498

operationally faster, necessarily introduce mechanical loads to the target.
499

The only in-orbit ADR technology demonstration to date is the RemoveDebris mission,
500

documented by Aglietti et al. Aglietti et al. [2020]. This mission successfully demonstrated
501

net capture of a CubeSat at 5 cm/s relative velocity and 7 m separation distance, as well
502

as harpoon ring at 20 m/s into a target panel at 1.5 m range. Two results are particu-
503

larly instructive. First, the net capture succeeded but was conducted against a cooperative
504

2U CubeSat (expanded to approximately 1 m pyramidal target), which is not representative
505

of real debris targets of 500 kg8 tonnes tumbling at 15 deg/s. Second, and more critically,
506

the harpoon test resulted in the snapping of the carbon bre boom from impact forces, de-
507

spite the harpoon itself being retained by its tether Aglietti et al. [2020]. This structural
508

failure during a controlled test illustrates the magnitude of impulse loads that contact-based
509

capture imposes.
510

3.1.1
The Fragmentation Paradox
511

The central paradox of rigid-body ADR is that the very act of removing debris may generate
512

new fragments, potentially worsening the environment it aims to protect. This concern is
513

supported by multiple lines of evidence:
514

 Wang et al. Zhang et al. [2022] state explicitly that rigid behaviour has the potential
515

to generate fragments during [the] capturing phase, hence increase [the] risk of further
516

space debris.
517

 Chen et al. Chen et al. [2024a] assess that single contact-based caging [is] excessively
518

risky for fast-tumbling targets with unknown massmomentum transfer could create
519

new debris.
520

 Dynamic simulations of the cancelled e.deorbit mission show peak torques of 195 Nm
521

at the manipulator joints when attempting to capture a target tumbling at only 5 deg/s
522

(the ENVISAT upper stage) Stol et al. [2017], reaching the operational limits of the
523

robotic joints.
524

 The Aerospace Corporation's IMPACT model establishes 10 J/g specic energy as the
525

threshold for catastrophic fragmentation of a satellite Aerospace Corporation [2020].
526

ClearSpace-1, the rst contracted commercial debris removal mission (ESA, ¿86M con-
527

tract), plans to use four rigid robotic arms to capture the Proba-1 satellite (95 kg, 0.6×0.6×
528

0.8 m) ClearSpace SA and European Space Agency [2020]. The mission's planning was itself
529

disrupted by the debris problem: the original target, the VESPA upper stage, was struck by
530

a tracked debris object during mission preparation, illustrating the cascading urgency of the
531

debris environment ClearSpace SA and European Space Agency [2020]. Launch is currently
532

planned for approximately 2029.
533

To place the fragmentation risk in perspective, we note that a rigid robotic arm exerting
534

195 Nm of torque on a 100 kg target at a 0.5 m lever arm produces a contact force of 390 N.
535

If this force acts over a contact area of 10 cm2 on a honeycomb panel with typical crush
536

strength of 13 MPa, the resulting stress of 0.39 MPa falls below the crush threshold of
537

the primary structure. However, the fragmentation risk is not primarily to the strongest
538

structural components, but to the most vulnerable: degraded solar panel hinge joints, aged
539

thermal blanket fasteners, corroded aluminium alloy brackets, and antenna feed structures
540

that have experienced decades of thermal cycling, UV degradation, and atomic oxygen ero-
541

sion. These appendage materials may have lost 3060% of their original strength through
542

environmental degradation, reducing eective crush thresholds well below nominal values.
543

For a tumbling 1000 kg upper stage at 5 deg/s, the angular momentum is approximately
544

50 N·m·s, and the impulsive loads during despin are proportionally larger. Applying the
545

catastrophic fragmentation threshold of 10 J/g from the IMPACT model Aerospace Corpo-
546

ration [2020], Johnson et al. [2001]: if a rigid grasp concentrates 50 J of despin energy into a
547

100 g solar panel hinge assembly, the resulting specic energy of 0.5 J/g remains below the
548

10 J/g threshold, but contact with a 10 g degraded thermal blanket fastener at equivalent
549

energy would yield 5 J/gapproaching the threshold. A compliant grasp distributing the
550

same energy over a larger contact area and longer time period reduces peak specic energy
551

by one to two orders of magnitude.
552

The fragmentation risk is therefore physically plausible and supported by qualitative as-
553

sessments, though not yet experimentally quantied. This survey adopts the precautionary
554

principle: compliant capture is preferred until quantitative data become available, on the
555

basis that the consequences of inadvertent fragmentation during ADRpotentially generat-
556

ing hundreds of new tracked objectsare severe enough to warrant risk-averse technology
557

selection even in the absence of denitive comparative data. A comprehensive, quantita-
558

tive comparison of fragmentation probability as a function of contact compliance remains
559


![Table 8](paper-23-v1_images/table_8.png)
*Table 8*

the single highest-priority open experimental question the community must address (see
560

Section 13).
561

Table 5 summarises the principal ADR technology classes, their technology readiness
562

levels (TRL), contact characteristics, and assessed fragmentation risk.
563

3.2
Soft and Compliant Capture Mechanisms
564

The fragmentation risk inherent in rigid capture has motivated the development of soft and
565

compliant alternatives that absorb, rather than transmit, kinetic energy during the capture
566

interaction. Eight distinct soft and compliant capture approaches have been documented in
567

the literature, all currently at TRL 25. We review each in turn, organised by their operating
568

principle: adhesion-based, bistable/passive, inatable-arm, and net-plus-inatable systems.
569

3.2.1
Gecko-Inspired Dry Adhesive Grippers
570

The most mature soft capture technology is the gecko-inspired dry adhesive gripper demon-
571

strated by Jiang et al. Jiang et al. [2017]. Published in Science Robotics, this system uses
572

shear-activated van der Waals adhesion pads with a load-sharing tendon-pulley mechanism
573

that scales adhesion from small patches to large contact areas. Critically, a nonlinear pas-
574


![Table 9](paper-23-v1_images/table_9.png)
*Table 9*

Table 5: Comparison of active debris removal technology classes.
Fragmentation risk is
assessed qualitatively based on published evidence; a quantitative comparison remains an
open research gap.

Method
TRL
Contact
Frag. Risk
Key Limitation

Rigid robotic
arm
56
Direct, rigid
High
Peak torques at
joint limits; brittle
appendage damage
Harpoon
6
Penetrative
Very high
Boom failure in
RemoveDebris;
target perforation
Thrown net
7
Enveloping
Moderate
Impulse at net
closure;
entanglement
dynamics
Ion beam
shepherd
4
Contactless
None
34 month timeline;
high power
Laser ablation
3
Contactless
None
Pointing accuracy;
space weapon
concerns
Gecko adhesive
45
Shear adhesion
Very low
Clean surfaces
assumed; no
tumbling test
Soft/inatable
arm
23
Compliant
Low
Precision; pneumatic
in vacuum
Bistable gripper
23
Passive snap
Low
Energy barrier
tuning; untested in
vacuum
Net + inatable
(INSIDeR)
∼4
Controlled net
Low
System integration
unproven in orbit

sive wrist provides high stiness during normal manipulation but becomes compliant under
575

overload, oering inherent protection against excessive contact forces.
576

The gecko gripper was validated in actual microgravity during NASA parabolic ight
577

campaigns, achieving capture success rates of 100% for spherical targets, 75% for cubic tar-
578

gets, and 81% for cylindrical targets, with objects up to approximately 400 kg and diameters
579

exceeding 1 m Jiang et al. [2017]. Failures were attributed to human operator misalignment
580

rather than adhesive performance. The system achieves essentially zero mechanical impact
581

forcea fundamental advantage for fragmentation avoidance. We note, following the taxon-
582

omy of Shintake et al. Shintake et al. [2018], that the gecko gripper is more precisely classied
583

as a compliant end-eector mechanism on a rigid platform rather than a fully soft robotic
584

system; nevertheless, its compliant capture interface directly addresses the fragmentation
585

concern. At TRL 45, it represents the highest-readiness soft capture technology, though
586

signicant gaps remain: all testing used cooperative (stationary) targets, and performance
587

under space vacuum, UV radiation, atomic oxygen exposure, and thermal cycling has not
588

been demonstrated.
589

3.2.2
Dielectric Elastomer Minimum Energy Structure (DEMES) Grippers
590

Araromi et al. Araromi et al. [2015] developed a DEMES-based deployable gripper explic-
591

itly for the CleanSpace One ADR mission. The device uses dielectric elastomer actuators
592

(DEAs) bonded to a exible frame, achieving rollable compact storage and deployment to
593

a multi-segment gripper with bending angles exceeding 60°. Each arm produces forces in
594

the mN range, sucient only for microgravity manipulation of small, lightweight targets.
595

The system demonstrated over 860,000 actuation cycles with individual arm mass below
596

0.65 g Araromi et al. [2015]. At TRL 34, the DEMES gripper is notable as the only soft
597

capture device explicitly designed for an actual ADR mission, although the CleanSpace One
598

mission architecture subsequently evolved without the gripper ying. Key limitations in-
599

clude the high operating voltage (∼kV) required for DEAs in vacuum (arcing risk) and the
600

absence of cryogenic or thermal cycling testing.
601

3.2.3
Bistable and Passive Capture Grippers
602

Two distinct bistable gripper concepts have been proposed for ADR. Liu et al. Liu et al.
603

[2022] developed a bistable snap-through gripper that captures targets using the kinetic
604

energy of the collision itself, requiring no external power for the grasping action. The gripper
605

deforms on contact, absorbs kinetic energy, triggers a bistable snap, and locks into the closed
606

conguration.
The energy barrier is adjustable through pre-deformation of the bistable
607

elements, allowing tuning for dierent target masses and approach velocities Liu et al. [2022].
608

This passive capture concept eliminates the need for precise actuation timinga signicant
609

advantage for tumbling, non-cooperative targets.
610

Zhang et al. Zhang et al. [2023b] propose a Venus ytrap-inspired bistable origami gripper
611

actuated by a shape memory alloy spring actuator (SMASA) that provides slow energy
612

storage followed by rapid release, with a DEA bristle-locking structure that prevents target
613

escape after capture. Capture is achieved within approximately 300 ms, and the device has
614

been demonstrated on complex geometries including asteroid models and spacecraft mock-
615

ups Zhang et al. [2023b].
Both bistable concepts remain at TRL 23, with no vacuum,
616

thermal, or microgravity testing.
617

3.2.4
Thermally Qualied Soft Grippers
618

Addressing the thermal environment is critical for any space capture mechanism.
Ruiz
619

Vincueria et al. Ruiz Vincueria et al. [2023] developed a multi-layered soft gripper combining
620

TPU, silicone, PTFE, and aerogel layers, tested across the full orbital thermal range from
621

−180°C to +220°C. A counter-intuitive but operationally signicant nding is that grasping
622

forces increase by 220% at cryogenic temperatures due to cold stiening of the elastomeric
623

layers, while decreasing by at most 50% at the hot extreme Ruiz Vincueria et al. [2023]. The
624

gripper uses MoS2 solid lubricant for vacuum compatibility and is available in dual and quad
625

arm congurations. This work provides the most quantitative thermal performance data
626

for any soft capture device and explicitly compares its approach against the ClearSpace-1
627

and Astroscale rigid arm architectures. However, all testing was conducted in laboratory
628

conditions without vacuum, radiation, or microgravity validation (TRL 2).
629

Foster-Hall et al. Foster-Hall et al. [2025] introduce a fundamentally dierent approach
630

to the cryogenic challenge: metallic cable-driven soft robotic structures tested at −196°C in
631

liquid nitrogen. Unlike elastomeric soft robots that embrittle at cryogenic temperatures, the
632

modular metallic cable structures exhibited only 5% stiness increase over 100 actuation cy-
633

cles, maintained full range of motion, and showed no microfractures under scanning electron
634

microscopyconsistent with cold-working behaviour in stainless steel rather than brittle
635

failure Foster-Hall et al. [2025]. Two-dimensional grasping was demonstrated at −196°C. At
636

TRL 23, this work opens a new design paradigm for soft space robotics beyond elastomers,
637

though three-dimensional manipulation and vacuum testing remain to be demonstrated.
638

3.2.5
Inatable Robotic Arms for Capture
639

Palmieri et al. Palmieri et al. [2023] developed the POPUP robot: a 7-DOF manipulator
640

with inatable links and rigid electric motor joints, incorporating visual servoing via dual
641

cameras and high-stiness bre reinforcement. The inatable links provide signicant mass
642

and volume reduction compared to equivalent rigid arms, and simulation demonstrates debris
643

capture feasibility despite the inherent compliance of the links Palmieri et al. [2023]. A 3-
644

DOF ground prototype has been statically characterised (TRL 3), but key challenges remain:
645

the compliance of inatable links reduces end-eector positioning precision, the pneumatic
646

ination system must operate in vacuum, and no thermal or radiation testing has been
647

performed.
648

3.2.6
INSIDeR: Net Capture with Inatable Deployment
649

The Innovative Net and Space Inatable structure for active Debris Removal (INSIDeR)
650

is a patented CNES/ESA-funded concept that combines the proven in-orbit heritage of
651

net capture (demonstrated by RemoveDebris) with inatable deployment structures CT
652

Ingénierie et al. [2017, 2021].
The system architecture comprises an inatable ring and
653

two inatable masts that deploy and guide a capture net, followed by a deorbit tether for
654

removal. The complete capture sequence proceeds through six phases: ination of the ring
655

and masts, net deployment, approach boost, mast detachment and deation, net capture,
656

and tether-assisted deorbit CT Ingénierie et al. [2017].
657

A key innovation is that the inatable masts provide controlled, slow net dynamics,
658

eliminating the large impulse peaks associated with conventional spring-ejected nets and
659

thereby reducing momentum transfer to the target CT Ingénierie et al. [2021]. The system
660

packages into a cube of approximately 50 cm per side, forming a plug-and-play ADR kit
661

adaptable to any target mass, morphology, or tumbling rate. Developed over 15 years by
662

CT Ingénierie and AirCaptif (Michelin group) with CNES and ESA co-funding, INSIDeR has
663

reached TRL ∼4 at the system level (individual subsystem technologies at TRL 5+), with
664


![Table 10](paper-23-v1_images/table_11.png)
*Table 10*

a ground demonstrator under construction as of 2021 CT Ingénierie et al. [2021]. ABAQUS
665

nite element simulations have conrmed net capture feasibility.
666

Table 6 provides a comprehensive comparison of all documented soft and compliant cap-
667

ture approaches.
668


![Table 11](paper-23-v1_images/table_10.png)
*Table 11*

Flight
qualified

Concept
Validation

102

Tendon-driven

Gecko adhesive

SMA (one-shot)

101

Bistable gripper
Inflatable arm

Vacuum-gap
electrostatic

Force output (N)

100

10
1

Category / Est. mass

10
2

Adhesive

Pneumatic

Electroactive

0.1 kg

DEMES/DEA

Mechanical

2 kg

Shape memory

5 kg

10
3

Passive

INSIDeR
(net capture,

TRL 4)

1
2
3
4
5
6
7
8
9
10
Technology Readiness Level (TRL)

Figure 2: Force output versus technology readiness level (TRL) for soft and compliant cap-
ture approaches. Marker size indicates system mass. The gecko adhesive gripper occupies
the highest-TRL, highest-force quadrant, representing the most ight-ready soft capture
technology.

The most signicant observation from this landscape is the absence of orbital ight
669

heritage for any soft capture system. The gecko adhesive gripper, at TRL 4 with microgravity
670

validation, and INSIDeR, at TRL 4 with system-level ground demonstration, represent the
671

nearest-term candidates for ight demonstration. We identify the combination of a gecko
672

adhesive gripper mounted on an inatable arm with bre Bragg grating structural health
673

monitoring (see Section 8.1) as the most ight-ready near-term soft ADR demonstratora
674

system that leverages the highest-TRL end-eector, the mass eciency of inatable links,
675

and embedded sensing for operational awareness.
676


![Table 12](paper-23-v1_images/table_12.png)
*Table 12*

Table 6: Technology readiness and performance comparison of soft and compliant capture
mechanisms for active debris removal. No soft capture system has own an orbital capture
mission to date.

Approach
Key
Reference
TRL
Force
Output
µg Test
Key
Limitation

4a
≤400 kg
objects
Yes
Clean surfaces;
no tumbling

Gecko
adhesive
Jiang
2017 Jiang
et al. [2017]

3b
mN range
No
Very low force;
HV arcing

DEMES/DEA Araromi
2015 Araromi
et al. [2015]

Inatable
arm
Palmieri
2023 Palmieri
et al. [2023]

3
Not
quantied
No
Low precision;
pneumatic in
vacuum
Flytrap
origami
Zhang
2023 Zhang
et al. [2023b]

23
Bistable
snap
No
SMA slow reset;
HV in vacuum

Bistable
gripper
Liu 2023 Liu
et al. [2022]
2
Passive
(KE input)
No
Energy barrier
tuning
Cryo
metallic
Foster-Hall
2025 Foster-
Hall et al.
[2025]

23
Not
quantied
No
2D only; no
vacuum

Thermal
multi-layer
Ruiz
2024 Ruiz Vin-
cueria et al.
[2023]

2
+220% at
cryo
No
Lab only; no
vacuum

INSIDeR
(net+in.)
ESA SDC
2017/21 CT
Ingénierie
et al. [2017,
2021]

4
N/A (net)
Sim. only
System
integration

aTRL 4 per NASA NPR 7123.1B: parabolic ight (∼20 s µg per parabola) constitutes component
validation in a simulated relevant environment rather than a fully relevant orbital environment (TRL 5).
bTRL 3: 860,000 cycles demonstrated in ambient conditions, but no space environment testing (vacuum,
thermal cycling, radiation) performed.

3.3
Inatable Debris Shields
677

Beyond active capture, inatable structures oer a complementary approach to the debris
678

problem through passive shielding. Conventional rigid Whipple shields Arnold et al. [2009],
679

which use spaced aluminium bumper plates to disrupt and disperse hypervelocity projectiles
680

before they reach the pressure wall, are eective but carry signicant mass and volume
681

penalties.
The substitution of rigid bumper plates with exible fabric layersusing the
682

same high-strength materials (Nextel ceramic fabric, Kevlar, and ultra-high molecular weight
683

polyethylene, UHMWPE) that form the basis of inatable habitat wallsenables deployable
684

shields with dramatically improved packaging eciency.
685

Destefanis et al. Destefanis et al. [2006] demonstrated that stued Whipple shields using
686

Nextel and Kevlar layers protect against projectiles twice the diameter of those stopped by
687

standard aluminium Whipple shields at equal areal density. This nding established the
688

performance advantage of fabric-based shielding architectures that underlies both habitat
689

micrometeoroid and orbital debris (MMOD) protection and standalone shield concepts.
690

Cha et al. Cha et al. [2024] present the Inatable Multi-Shock Shield (IMSS), which ap-
691

plies waterbomb tessellation origami to create a deployable multi-bumper debris shield that
692

expands approximately 80% beyond its initial radius while achieving 90% volume savings
693

compared to an equivalent rigid Whipple shield. The IMSS uses UHMWPE bre for ballistic
694

protection within a ve-bumper conguration, with 50 mm bumper spacing accommodated
695

in a 400 mm stowed stack Cha et al. [2024]. A critical design feature is that all material
696

in the deployed conguration contributes to debris protectionthere is no structural dead
697

weight. The origami fold geometry that enables compact packaging simultaneously creates
698

the inter-bumper spacing required for eective hypervelocity projectile disruption, embody-
699

ing a dual-functionality design principle applicable to large deployable structures generally
700

(see Section 4.3 for related deployment mechanics).
701

At TRL 23, the IMSS concept requires further development in hypervelocity impact
702

validation, large-scale (>10 m) deployment demonstration, and ination system design.
703

Nevertheless, the material commonality between inatable debris shields, inatable habi-
704

tat MMOD layers, and inatable robotic arm structural fabrics reinforces the survey's
705

central thesis: the same high-strength fabric technology baseVectran, Kevlar, Nextel,
706

UHMWPEenables debris capture, debris protection, and habitable volume creation.
707

For very large-scale applications, inatable debris shields of 100 m class have been pro-
708

posed as orbital infrastructure to protect high-value assets or clear debris corridors. Such
709

structures would require the attitude and orbit control technologies discussed in Section 11
710

and the robotic in-orbit assembly capabilities reviewed in Section 12, linking the passive
711

protection concept back to the active robotic systems that are the primary focus of this
712

survey.
713

4
Use Cases: Habitats and Exploration
714

Inatable space structures for human habitation represent the second major application
715

domain where soft and exible technologies oer transformative advantages over conventional
716

rigid systems. The fundamental value proposition is mass eciency: high-strength fabrics
717

such as Vectran and Kevlar possess specic tensile strengths of 2,330 and 2,080 kN·m/kg
718

respectively at the fabric level (or 2,500 kN·m/kg for Kevlar 49 lament)more than an
719

order of magnitude greater than titanium alloy Ti-6Al-4V at 220 kN·m/kg or aluminium
720

7075 at 204 kN·m/kg Valle et al. [2019a]. This advantage translates directly into the ability
721

to launch habitable volumes that would be physically impossible with metallic construction
722

within current launch vehicle fairing constraints. A fabric-walled habitat is not merely a
723

lighter alternative to a metallic module; it enables architectural possibilitiesvolumes of
724

3001,400 m3that have no rigid equivalent.
725

This section traces the heritage of inatable space habitation from its origins in 1960 to
726

the present day (Section 4.1), reviews current commercial programs (Section 4.2), surveys
727

future concepts for lunar, Martian, and planetary applications (Section 4.3), and addresses
728

the critical issue of radiation shielding with an honest assessment of the BEAM solar particle
729

event ndings (Section 4.4).
730

4.1
Heritage Timeline: Echo to BEAM
731


![Table 13](paper-23-v1_images/table_15.png)
*Table 13*

The heritage of inatable space structures spans over six decades, progressing through a
732

non-linear TRL trajectory marked by both remarkable successes and programmatic setbacks.
733

Table 7 summarises the key milestones.
734


![Table 14](paper-23-v1_images/table_13.png)
*Table 14*

LIFE in-space

Volga airlock

IRVE-II

test
(~2026, planned)

(1965)

(2009)

Mars Pathfinder

BEAM
(2016)

TRL 9

TRL 7

TRL 6

(1997)

Echo 1
(1960)

Genesis I

LOFTID

TRL 9

TRL 8

(2006)

(2022)

TRL 9

TRL 8

TRL 7


![Table 15](paper-23-v1_images/table_14.png)
*Table 15*

IAE / Spartan 207

IRVE-3

ClearSpace-1
(~2029, planned)

(1996)

(2012)

Echo 2
(1964)

Genesis II

Sierra LIFE

TRL 7

TRL 7

TRL 5

(2007)

(UBP tests)

TransHab

InflateSail

TRL 9

TRL 8

(2024)

NASA

ESA / International

(1999)

(2017)

TRL 5

TRL 6

TRL 7

Commercial

Planned

1960
1970
1980
1990
2000
2010
2020
2030
Year

Figure 3: Heritage timeline of inatable space structures from Echo 1 (1960) to LOFTID
(2022), illustrating the progression from passive communication reectors through human-
rated habitats to active aerodynamic decelerators. Colour coding indicates programme ori-
gin; marker size reects achieved TRL.


![Table 16](paper-23-v1_images/table_16.png)
*Table 16*

Table 7: Heritage timeline of inatable space structures, from passive communication reec-
tors to human-rated orbital habitats. TRL ratings reect achieved (not planned) readiness
at programme conclusion or present status.

Year
Programme
TRL
Key Achievement

1960
Echo 1 (NASA)
9
30.5 m (100 ft) Mylar sphere; 8+ years
on-orbit; global communications relay
1965
Volga airlock
(USSR)
9
First human-rated inatable;
Voskhod-2 EVA (Leonov); 40 airbags,
3 independent groups, 7 min ination
1996
IAE/Spartan 207
(NASA)
7
14 m antenna; 28 m Kevlar/Neoprene
booms; Shuttle deployment
demonstration
1997
Mars Pathnder
airbags
9
Vectran fabric; operational landing on
3 missions (Pathnder, Spirit,
Opportunity)
19972000
TransHab (NASA
JSC)
56
8.2 m × 11 m; 5-layer shell; tested to
4× operating pressure; cancelled by
Congress (HR 1654)
200607
Genesis I/II
(Bigelow)
78
Orbital validation; 2.5+ years
on-orbit; pressure retention conrmed
2009
IRVE-II (NASA
LaRC)
7
3 m inatable reentry vehicle
experiment; suborbital demonstration
2016+
BEAM
(Bigelow/NASA)
9
16 m3; 1,415 kg; 8+ years on ISS;
converted to cargo storage; operational
2022
LOFTID (NASA)
78
6 m inatable aerodecelerator; orbital
reentry at Mach 30

4.1.1
Early Inatables: Echo and Volga (19601965)
735

Project Echo, initiated by NASA in 1960, deployed Echo 1 as a 30.5 m diameter Mylar
736

balloon serving as a passive communications reector Litteken [2019]. The satellite operated
737

for over eight years and enabled global communications experiments and geodetic measure-
738

ments. Echo 2 (1964) advanced the concept with a rigidisable aluminium foil/Mylar laminate
739

structure. While neither was habitable, the Echo programme demonstrated that large, thin-
740

walled inatable structures could survive the LEO environment for extended periods.
741

The Volga airlock, deployed for the Voskhod-2 mission in 1965, represents the rst human-
742

rated inatable space structure Litteken [2019]. Designed for Alexei Leonov's historic rst
743

spacewalk, the Volga used 40 airbags arranged in three independent groups to inate a 2.4 m
744

long, 1.2 m diameter cylindrical airlock in seven minutes. The successful EVA validated
745

the fundamental concept that pressurised inatable structures could safely support human
746

operations in space, albeit for a single use.
747

4.1.2
TransHab: Proving the Five-Layer Architecture (19972000)
748

The Transit Habitat (TransHab) programme at NASA Johnson Space Center represented the
749

most ambitious inatable habitat development prior to BEAM. Under Principal Architect
750

Kriss Kennedy Kennedy [2002] and shell lead Gerard Valle, the team developed an 8.2 m
751

diameter, 11 m long module with a ve-layer shell architecture that has become the standard
752

for all subsequent inatable habitat designs Valle et al. [2019a]:
753

1. Inner liner: Nomex scu protection layer.
754

2. Bladder: Multiple redundant layers, oversized relative to the restraint layer and car-
755

rying zero structural load.
756

3. Restraint layer: Tight basket-weave Kevlar/Vectran biaxial membrane, designed to
757

a safety factor of 4.0× per NASA-STD-5001.
758

4. MMOD shield: Ceramic (Nextel) bumper, open-cell foam spacer, and Kevlar rear
759

wallvacuum-packed for launch, with foam self-expanding in orbit.
760

5. Multi-layer insulation (MLI): 19 layers of double-aluminised Mylar/Kapton, with
761

perforated inner layers for venting during depressurisation.
762

TransHab was tested to 4× ambient pressure (>54 psig) in a September 1998 hydrostatic
763

burst test, and full-scale vacuum deployment was demonstrated Kennedy [2002]. Hyperveloc-
764

ity impact testing conrmed that the MMOD shield outperformed the aluminium structure
765

of ISS modules. The programme also pioneered the water wall radiation shelter concept,
766

positioning crew quarters within a rigid central core surrounded by water-lled containers
767

for radiation protection Kennedy [2002].
768

Despite reaching TRL 56, TransHab was cancelled by Congressional action (HR 1654,
769

2000). The technology investment was preserved through patent licensing to Bigelow Aerospace,
770

which continued development commercially Kennedy [2002].
771

4.1.3
Genesis and BEAM: Orbital Validation (20062016+)
772

Bigelow Aerospace launched Genesis I (2006) and Genesis II (2007) as uncrewed orbital test
773

modules, demonstrating pressure retention (69.672.4 kPa for Genesis II) and thermal per-
774

formance (average 26°C, range 4.532°C for Genesis I) over 2.5+ years Litteken [2019]. These
775

missions validated the TransHab-derived shell architecture in the actual orbital environment
776

for the rst time.
777

The Bigelow Expandable Activity Module (BEAM), launched to the International Space
778

Station in April 2016, represents the culmination of this heritage. BEAM provides 16 m3 of
779

habitable volume at a mass of 1,415 kg (88 kg/m3), compared to 137 kg/m3 for the Columbus
780

module and 205 kg/m3 for the Tranquility node Valle et al. [2019a]. While BEAM's mass-
781

per-volume ratio is higher than TransHab's projected 39 kg/m3reecting BEAM's small
782

size and relatively heavy end-ttingsthe comparison to metallic modules demonstrates the
783

eciency advantage of fabric-walled construction Valle et al. [2019a].
784

BEAM's deployment provided a critical engineering lesson. Initial expansion attempts
785

failed, and the module required 25 short pressure bursts over approximately 7 hours to
786

achieve full deploymentin contrast to the planned rapid ination sequence NASA Johnson
787

Space Center [2017]. The root cause was attributed to softgoods layers adhering after years
788

of compression in the launch conguration. For future free-ying deep-space modules where
789

ISS crew intervention would not be available, this deployment failure mode must be resolved
790

through autonomous ination protocols.
791

After its planned two-year demonstration, BEAM's mission was extended to at least 2028.
792

The module has been converted to active cargo storage (approximately 130 cargo transfer
793

bags), demonstrating practical volumetric value beyond its test objectives NASA Johnson
794

Space Center [2017]. No pressure loss, structural degradation, or signicant MMOD impacts
795

have been recorded in over eight years of operation.
The Distributed Impact Detection
796

System (DIDS) has continuously monitored for debris impacts throughout the mission.
797

4.2
Current Commercial Programs: LIFE, Orbital Reef, and Be-
798

yond
799

4.2.1
Sierra Space LIFE
800

The Large Integrated Flexible Environment (LIFE) programme by Sierra Space represents
801

the most advanced current inatable habitat development. The programme has conducted
802

a systematic Ultimate Burst Pressure (UBP) test campaign at NASA Marshall Space Flight
803

Center, producing two landmark results Sierra Space Corporation [2024]:
804

 January 2024 (full-scale): A full-scale LIFE 285 expandable structure (approx-
805

imately 300 m3, over 6 m tall) burst at 77 psi (531 kPa), exceeding NASA's rec-
806

ommended threshold of 60.8 psi (4× the 15.2 psi maximum operating pressure per
807

NASA-STD-5001) by 27% Sierra Space Corporation [2024].
808

 OctoberNovember 2024 (1/3 scale): The LIFE 10 module burst at 255 psi
809

(1,758 kPa), achieving a factor of safety of 16× for LEO operations (at 15.2 psi) and
810

23× for lunar surface operations (at 10.8 psi) Sierra Space Corporation [2024].
811

The LIFE product line spans three variants: LIFE 10 (∼100 m3 equivalent, 1/3 scale,
812

for lunar surface applications), LIFE 285 (∼300 m3, full-scale, for ISS-attached or free-
813

ying stations), and LIFE 500 (6001,440 m3, exceeding the total pressurised volume of the
814

ISS) Sierra Space Corporation [2024]. The restraint layer uses Vectran straps manufactured
815

by ILC Dover, the same organisation responsible for TransHab, Mars Exploration Rover, and
816

BEAM softgoods. Sierra Space is partnered with Blue Origin for the Orbital Reef commercial
817

space station, which received a $130M NASA Commercial LEO Destinations (CLD) award
818

in December 2021. An in-space test is targeted for no earlier than 2026.
819

4.2.2
Historical Context: B330 and Commercial Ecosystem Fragility
820

The history of Bigelow Aerospace provides a cautionary counterpoint. The B330 (330 m3,
821

18,50023,000 kg, 2436 layers totalling approximately 0.46 m wall thickness Valle et al.
822

[2019a]) was the most advanced commercial inatable habitat design as of 2019, with a full-
823

scale ground prototype (XBASE) tested under NASA's NextSTEP programme. The B330's
824

restraint design used a distinctive hoop webbing approach (US Patent 7,100,874) diering
825

from NASA's basket-weave architecture Valle et al. [2019a].
826

Bigelow Aerospace ceased operations in March 2020 following COVID-19 layos, and
827

BEAM's ownership was transferred to NASA JSC in December 2021. The collapse of the
828

most mature commercial inatable habitat programme illustrates that high TRL does not
829

guarantee commercial viability. Future programmes cannot rely on government safety nets
830

to preserve technology investments, and the commercial ecosystem supporting inatable
831

habitat development remains fragile.
832

4.2.3
NextSTEP Competitive Landscape
833

NASA's NextSTEP-2 programme (20162019) selected six companiesBigelow, Boeing,
834

Lockheed Martin, Orbital ATK, Sierra Nevada Corporation, and NanoRacksto develop
835

habitat prototypes for evaluation NASA [2016].
Lockheed Martin's inatable prototype
836

achieved a burst pressure of 285 psi with hundreds of sensors and high-speed cameras mon-
837

itoring the failure Lockheed Martin [2022]. However, this programme subsequently pivoted:
838

the Starlab commercial station (originally Lockheed Martin/NanoRacks) adopted a rigid
839

architecture with Airbus as partner, abandoning the inatable approach. Of the six original
840

NextSTEP-2 companies, only Sierra Space (evolved from Sierra Nevada Corporation) has
841

continued to develop inatable habitats. This consolidation, combined with Bigelow's exit,
842

suggests that the inatable habitat technology faces unresolved commercialisation challenges
843

that complement the technical risks discussed elsewhere.
844

4.3
Future Concepts: Lunar Surface, Mars Transit, Planetary En-
845

try
846

4.3.1
Lunar Surface Habitats
847

Multiple concepts have been proposed for inatable habitats on the lunar surface, where
848

the reduced gravity (1/6 g) and absence of orbital debris shift the design requirements
849

from MMOD protection toward radiation shielding and dust management. The ESA-Hassell
850

collaboration has designed a scalable inatable pod system at the Shackleton Crater (lunar
851

south pole), partially constructed from lunar regolith via 3D printing and expandable to
852

house up to 144 people Hassell Studio and European Space Agency [2024]. The ESA-SOM
853

Moon Village concept proposes a semi-inatable shell that doubles its internal volume upon
854

deployment, supporting a four-person crew for up to 300 days Skidmore, Owings & Merrill
855

and European Space Agency [2019]. The ESA Pneumocell concept is specically designed
856

for burial under 45 m of regolith, using the lunar soil itself as radiation shielding European
857

Space Agency [2018]an elegant solution that leverages the inatable structure's compliance
858

to conform to the excavated cavity.
859

For lunar operations, the MMOD layer that constitutes approximately 68% of the shell
860

mass in LEO Valle et al. [2019a] can be substantially reduced or eliminated, oering signi-
861

cant mass savings. However, lunar dust intrusion and abrasion present a new challenge for
862

exible fabric surfaces that has not been addressed in any inatable habitat design to date.
863

4.3.2
Mars Transit and Surface Applications
864

TransHab was originally conceived as a Mars transit vehicle, and the deep-space habitat
865

architecture inherits directly from this heritage. Valle et al. Valle et al. [2019a] present a
866

launch-to-activation deployment owchart for a deep-space inatable habitat, identifying key
867

operational challenges: autonomous deployment without crew intervention, up to 4 kW of
868

heater power required post-ination to bring the bladder above minimum operating tem-
869

perature, and up to 24 hours before crew entry is permitted. For a three-year Mars transit
870

mission at solar minimum with three solar particle events (SPEs), radiation shielding re-
871

quirements range from 25 cm to 400 cm of water equivalent depending on the allowable bone
872

marrow dose Valle et al. [2019a]a signicant design driver discussed further in Section 4.4.
873

Mars surface applications extend to entry systems. The Low-Earth Orbit Flight Test of
874

an Inatable Decelerator (LOFTID, 2022) demonstrated a 6 m diameter inatable aerodecel-
875

erator at Mach 30 during orbital reentry NASA [2022], achieving TRL 78 and establishing
876

the viability of inatable heat shields for planetary entry. The Inatable Reentry Vehicle
877

Experiment (IRVE-II, 2009) had previously validated a 3 m prototype in suborbital ight Lit-
878

teken [2019]. For Mars, where the thin atmosphere limits the eectiveness of parachutes for
879

large payloads, inatable aerodecelerators oer the only viable path to landing human-scale
880

masses (>20 tonnes) on the surface. More exotic concepts include the HAVOC Venus air-
881

ship and the Titan Aerover blimp, both leveraging inatable structures for buoyancy-based
882

exploration Litteken [2019].
883

4.3.3
European Programmes
884

European contributions to inatable habitat development include the ASI-funded FLECS
885

(Flexible Commercial Structure), the ESA-funded IHAB (Inatable Habitation) and IMOD
886

(Inatable Module) programmes, and the 2002 ESA/ESTEC First European Workshop
887

on Inatable Space Structures (ESA-WPP-200) ESA/ESTEC [2002]. These programmes
888

have contributed materials characterisation, hypervelocity impact testing of exible MMOD
889

shields (notably Destefanis et al. Destefanis et al. [2006]), and architectural concepts. How-
890

ever, it must be noted that no European inatable has own in a habitation role. After
891

more than two decades of investment, all European inatable habitat programmes remain at
892

TRL 24. The Volga airlock (1965) remains the only European-adjacent (Soviet-era) ight
893

precedent for a human-rated inatable in space.
894

4.4
Radiation Shielding: The BEAM SPE Findings and Design Im-
895

plications
896

Radiation shielding represents the single most serious unresolved technical challenge for
897

inatable habitats in deep space. The BEAM module has provided the only in-ight radiation
898

data for an inatable habitat, and the ndings demand honest assessment.
899

During the September 2017 solar particle event (SPE), radiation dosimeters inside BEAM
900

recorded approximately 22.5 mGy, compared to approximately 0.25 mGy measured in typ-
901

ical ISS metallic habitable modules during the same eventa ratio of 810× higher dose
902

inside the inatable module NASA Johnson Space Center [2017]. For galactic cosmic ra-
903

diation (GCR), which is continuous rather than episodic, BEAM dose rates were similar
904

to other ISS modules at baseline, indicating that the fabric shell provides adequate GCR
905

shielding in LEO where the Earth's magnetic eld supplies primary protection.
906

The SPE nding has signicant implications:
907

 Fabric walls alone are insucient for SPE protection. The multi-layer shell
908

(60+ individual layers, 3050 cm total thickness) provides substantially less shielding
909

than the aluminium structure of ISS modules during particle events.
910

 The mitigation is designed-in, not absent. Both TransHab and the LIFE archi-
911

tecture incorporate a rigid central core functioning as a storm shelter during SPEs.
912

Crew quarters are positioned within this core, surrounded by water wall containers
913

(a concept originating with Kennedy's TransHab design Kennedy [2002]) that provide
914

eective hydrogen-rich shielding. The inatable volume provides habitable space for
915

non-storm operations, while the rigid core provides radiation protection.
916

 Material selection matters. Polyethylene provides 27.8% mass savings compared to
917

aluminium for equivalent radiation shielding eectiveness, and three-layer composite
918

shields (combining high-Z, medium-Z, and low-Z materials) achieve up to 70% total
919

ionising dose improvement for electrons and 50% for protons Norbury et al. [2025].
920

For deep-space missions beyond Earth's magnetosphere, the GCR environment is more
921

severe and continuous. Valle et al. Valle et al. [2019a] model that a three-year deep-space
922

mission at solar minimum with three SPEs requires between 25 cm and 400 cm of water-
923

equivalent shielding depending on the allowable bone marrow dosetranslating to substan-
924

tial mass within the rigid core. Active magnetic shielding and pharmaceutical countermea-
925

sures remain at low TRL and are not viable near-term solutions.
926

The honest framing is that inatable habitats are not radiation protection structures,
927

and were never designed to be. They are mass-ecient volume structures with integrated
928

MMOD protection. Radiation protection is the responsibility of the rigid core and water wall
929

architecture. The BEAM SPE data conrms this design philosophy rather than undermining
930

it, but the data must be presented without minimisation to maintain credibility with the
931

radiation protection community. The absence of post-2017 follow-up publications detailing
932

BEAM's continued radiation environment data over its now eight-year mission represents a
933

gap in the available evidence base that future studies should address.
934

5
State of the Art: Materials and Structures
935

The material systems underpinning inatable space structures occupy a unique design space:
936

they must combine the tensile strength of structural metals, the exibility to package into
937

compact launch volumes, and the environmental durability to survive atomic oxygen, ultra-
938

violet radiation, and micrometeoroid impacts for mission lifetimes spanning years to decades.
939

This section reviews the four dominant fabric families, the canonical multi-layer shell archi-
940

tecture derived from TransHab, established rigidisation technologies, and the environmental
941

degradation mechanisms that govern long-term performance.
942

5.1
Space-Rated Fabrics: Vectran, Kevlar, Zylon, Nextel
943

Four high-performance fabric families dominate inatable space structure design, each oc-
944

cupying a distinct functional niche determined by the intersection of mechanical properties,
945

environmental tolerance, and ight heritage.
946

Vectran HT (liquid crystal polymer, Kuraray Co.) has emerged as the preferred ma-
947

terial for restraint layers in inatable habitats. With a tensile strength of approximately
948

3.0 GPa at a density of 1.40 g/cm3, Vectran achieves a specic strength of 2,330 kN·m/kg
949

an order of magnitude above Ti-6Al-4V (220 kN·m/kg) and Al 7075 (204 kN·m/kg) Valle
950

et al. [2019b]. Vectran's principal advantage over the earlier-generation Kevlar is its superior
951

creep resistance: under sustained load at the NASA-mandated factor of safety of 4.0 (corre-
952

sponding to 25% of ultimate tensile strength), Vectran fabric exhibits no failure over extended
953

test periods of months Weadon [2013]. This characteristic is critical because creep is the life-
954

limiting mechanism for restraint layers in pressure-stabilised structures. However, Weadon's
955

systematic characterisation revealed that time-to-failure is exponentially sensitive to load
956

level, and manufacturing variability in ultimate tensile strength (±10% for 12K webbing,
957

±6% for 6K webbing) introduces signicant uncertainty in lifetime predictionat 7585%
958

UTS, time-to-failure ranges from 4 minutes to 5.5 months for identical test congurations
959

Weadon [2013].
This nding underscores the importance of quality control in inatable
960

habitat fabrication. Two important qualications must be noted. First, Weadon's creep
961

characterisation was conducted at room temperature; no published Vectran creep dataset
962

exists for space-representative thermal cycling conditions (approximately −100◦C to +120◦C
963

for LEO), and the eective creep rate under such cycling may dier signicantly from room-
964

temperature datathis represents a critical materials gap for habitat lifetime prediction.
965

Second, the no failure over extended test periods result at 25% UTS, while encouraging, is
966

based on a limited number of specimens at the design operating point; given the wide man-
967

ufacturing variability, condence intervals on lifetime prediction remain large, and the creep
968

behaviour exhibits bimodal characteristics where some specimens show substantially earlier
969

failure than others at identical load levels. Vectran's ight heritage includes Mars Pathnder
970

airbags (1997), BEAM restraint layers (2016present), and the Sierra Space LIFE program
971

Litteken [2019].
972

Kevlar 49 (poly-paraphenylene terephthalamide, DuPont) was the original restraint
973

layer material for TransHab, with a tensile strength of approximately 3.0 GPa at the fabric
974

level and 3.6 GPa at the individual lament level, at a density of 1.44 g/cm3 Valle et al.
975

[2019b], DuPont [2019]. The corresponding specic strength is 2,080 kN·m/kg (fabric) or
976

2,500 kN·m/kg (lament); throughout this survey, fabric-level properties are reported un-
977

less otherwise noted, as these are the engineering-relevant values for woven restraint layers.
978

While Kevlar's fabric-level specic strength is comparable to Vectran's, its higher creep rate
979

under sustained biaxial loading led to its replacement by Vectran in subsequent habitat de-
980

signs Kennedy [2002]. Kevlar retains an important role as a rear-wall material in multi-layer
981

micrometeoroid and orbital debris (MMOD) shields, where its combination of high energy
982

absorption and relatively low cost makes it the material of choice for fragment capture layers
983

Destefanis et al. [2003]. Space environment characterisation by Destefanis et al. conrmed
984

that Kevlar suers UV-induced discoloration and embrittlement but shows acceptable perfor-
985

mance when shielded from direct solar exposure within the MMOD sub-assembly Destefanis
986

et al. [2009].
987

Zylon (poly-p-phenylene-2,6-benzobisoxazole, PBO; Toyobo Co.) oers the highest ten-
988

sile strength of any commercially available high-performance bre at 5.8 GPa, yielding a
989

specic strength of 3,840 kN·m/kg Toyobo Co., Ltd. [2005]. However, Zylon exhibits catas-
990

trophic UV degradation: strength loss of approximately 35% within 6 months of unshielded
991

exposure, rendering it unsuitable for any application without comprehensive UV protec-
992

tion Toyobo Co., Ltd. [2005], Said et al. [2006]. Despite this limitation, Zylon has found
993

niche space applications where UV shielding is inherently provided: SpaceX Crew Dragon
994

parachute risers and NASA high-altitude balloon tendons Litteken [2019]. For inatable
995

structures, Zylon could serve in interior tensile elements (e.g., oor suspension webbings
996

within pressurised habitats) where the multi-layer shell provides UV shielding, but its UV
997

sensitivity eectively precludes use in any externally exposed role.
998

Nextel 440 (3M alumina-boria-silica ceramic fabric) occupies a unique position as the
999

only ceramic bre used in inatable space structures. With a density of 3.05 g/cm3 and con-
1000

tinuous use temperature of 1370◦C, Nextel is employed exclusively as the outer bumper layer
1001

in MMOD shielding Christiansen and Davis [2019], Destefanis et al. [2003]. Upon hyperveloc-
1002

ity impact, Nextel fragments incoming particles into smaller, more widely dispersed debris,
1003

reducing the energy density impinging on subsequent shield layers. The stued Whipple
1004

conguration (Nextel bumper + open-cell foam + Kevlar rear wall) protects against projec-
1005

tiles approximately twice the diameter of those defeated by a standard aluminium Whipple
1006

shield at equal areal density Destefanis et al. [2003]. Nextel is inherently immune to UV and
1007

atomic oxygen degradation due to its ceramic composition, but its high density limits its use
1008

to the thin bumper layer.
1009

Two additional materials complete the palette for inatable structures.
Beta cloth
1010

(PTFE-coated breglass) serves as the outermost atomic oxygen protection cover layer, with
1011

LDEF ight data demonstrating excellent durability over 68 months of LEO exposure Linton
1012

et al. [1993], Banks et al. [2004]. Kapton H (polyimide, DuPont) is the workhorse lm for
1013

multi-layer insulation, operating from −269◦C to +400◦C, though it is susceptible to atomic
1014

oxygen erosion at a rate of 3.0×10−24 cm3/atom Banks et al. [2004], Finckenor and Dooling
1015

[1999].
1016


![Table 17](paper-23-v1_images/table_17.png)
*Table 17*

Table 8 presents a comprehensive comparison of these material systems across eight
1017


![Table 18](paper-23-v1_images/table_18.png)
*Table 18*

performance parameters relevant to inatable space structures.
1018

Table 8: Comparison of space-rated materials for inatable structures.

Material
Type
σUTS
ρ
Tmax
UV
AO
Primary
TRL
(GPa)
(g/cm3)
(◦C)
Sens.
Resist.
Role

Vectran HT
LCP bre
3.0
1.40
330
Mod.
Low
Restraint
9
Kevlar 49
Aramid
3.0
1.44
427
High
Low
MMOD rear
9
Zylon AS
PBO bre
5.8
1.54
650
V. High
Low
Interior only
7
Nextel 440
Ceramic

3.05
1370
None
N/A
MMOD bumper
9
Kapton H
Polyimide
0.23
1.42
400
Low
Low
MLI layers
9
Beta cloth
PTFE/glass
0.34

650
Low
High
AO cover
9


![Table 19](paper-23-v1_images/table_19.png)
*Table 19*

Table 9: Specic strength comparison: high-performance fabrics versus structural metals
(data from Valle et al. Valle et al. [2019b]).

Material
σUTS (GPa)
ρ (g/cm3)
Specic Strength (kN·m/kg)
Ratio to Ti-6Al-4V

Zylon AS
5.8
1.54
3,840
17.5×
Kevlar 49 (fabric)
3.0
1.44
2,080
9.5×
Vectran HT
3.0
1.40
2,330
10.6×
Ti-6Al-4V
0.95
4.43
220
1.0×
Al 7075-T6
0.57
2.81
204
0.9×

5.2
Multi-Layer Shell Architecture
1019

The TransHab program (19972000) established the canonical ve-layer shell architecture
1020

that remains the reference design for all subsequent inatable habitats Kennedy [2002, 2016].
1021

From innermost to outermost, the layers are:
1022

1. Liner: Nomex fabric backed by Kevlar felt provides the crew-contact interior surface,
1023

oering acoustic attenuation and a substrate for equipment mounting.
1024

2. Bladder: Three redundant layers of polymeric gas barrier (Combitherm or urethane-
1025

coated Nylon), each sandwiched between Kevlar felt separators. The bladder is deliber-
1026

ately oversized relative to the restraint layer so that it carries no structural loadthe
1027

positive pressure dierential is transmitted entirely to the restraint layer Kennedy
1028

[2016]. The triple redundancy ensures continued pressure containment after a single-
1029

layer puncture.
1030

3. Restraint layer: The primary load-carrying element, comprising Kevlar (TransHab)
1031

or Vectran (BEAM and subsequent designs) in a biaxial basket-weave conguration.
1032

TransHab's restraint layer was designed to sustain 12,500 lb/in hoop loading and
1033

High-performance

High-perf. fabric
Structural metal
Ceramic
Polymer film

fabrics

Zylon AS

Kevlar 49

Specific strength (kN·m/kg)

Vectran HT

103

Ceramics

Nextel 440

Ti-6Al-4V

Al 7075-T6
Kapton H

Structural metals

102

200
400
600
800
1000
1200
1400
1600
Maximum service temperature (°C)

Figure 4: Materials Ashby chart comparing specic strength versus maximum service tem-
perature for space-rated fabrics and structural metals.
High-performance fabrics (Vec-
tran, Kevlar, Zylon) occupy a design space inaccessible to metals, combining an order-
of-magnitude advantage in specic strength with adequate thermal performance for LEO
applications.

6,000 lb/in axial loading at a factor of safety of 4.0 per NASA-STD-5001 Kennedy
1034

[2016]. The restraint layer attaches to rigid bulkheads via clevis ttings that transfer
1035

membrane loads to the metallic core structure. Ground testing demonstrated sustained
1036

pressure at 4× operating pressure (60 psid) without failure, and burst at 196 psid in
1037

sub-scale articles Kennedy [2002].
1038

4. MMOD shield: A stued Whipple conguration comprising Nextel 440 ceramic fab-
1039

ric bumper layers, open-cell polyurethane foam spacers, and Kevlar rear walls Deste-
1040

fanis et al. [2003]. The MMOD assembly is vacuum-packed during launch to maintain
1041

the folded conguration and expands passively on orbit when exposed to vacuum.
1042

TransHab's MMOD design was tested against projectiles up to 1.7 cm diameter at
1043

hypervelocity, meeting the no-penetration probability requirement of PNP ≥0.9820
1044

Kennedy [2002]. Damage tolerance testing by Trevino et al. demonstrated that a 2 in
1045

× 3.5 in hole in the restraint layer at 25% of burst pressure resulted in load redistri-
1046

bution without catastrophic failurean inherent advantage of woven textile structures
1047

over metallic shells Edgecombe et al. [2009].
1048

5. Thermal protection system (TPS): Multi-layer insulation comprising nylon-reinforced
1049

double-aluminized Mylar and double-aluminized Kapton layers, with inner layers perfo-
1050

rated for gas venting during deployment Finckenor and Dooling [1999]. The outermost
1051

element is an atomic oxygen cover of Beta glass fabric for LEO operations Kennedy
1052

[2016]. Eective emittance values for properly installed MLI range from 0.015 to 0.05,
1053

though practical performance with seams, penetrations, and attachment hardware typ-
1054

ically falls at the upper end of this range Finckenor and Dooling [1999], Gilmore [2002].
1055

The total shell assembly comprises 60+ individual layers deployed to a thickness of 30
1056

50 cm Valle et al. [2019b]. For TransHab, the overall packaged dimensions were 10.5 m length
1057

with a deployed width of 8.3 m, yielding an internal habitable volume of approximately
1058

161 m3 and a total packaged shell volume of 329 m3 Kennedy [2016]. BEAM, the ight-
1059

demonstrated derivative, achieves a habitable volume of 16 m3 in a 1,415 kg module Valle
1060


![Table 20](paper-23-v1_images/table_20.png)
*Table 20*

et al. [2019b].
1061

Table 10: Layer-by-layer specication of the TransHab/BEAM shell architecture. The her-
itage convention identies ve functional sub-assemblies; the AO cover (Beta cloth) is the
outermost element of the TPS sub-assembly but is listed separately here for clarity, yielding
six table rows for ve sub-assemblies.

Sub-assy
Layer
Material(s)
Function
Key Specica

1
Liner
Nomex + Kevlar felt
Crew contact, acoustic
Non-structural
2
Bladder (×3)
Combitherm / Urethane-Nylon
Gas barrier
3× redundant,
3
Restraint
Vectran basket-weave
Primary structure
FOS = 4.0, 12
4
MMOD
Nextel + foam + Kevlar
Debris protection
PNP ≥0.9820

5
TPS/MLI
Aluminized Mylar/Kapton
Thermal control
εe = 0.0150.0
AO cover
Beta glass fabric
AO protection (outer TPS)
LDEF-validate

TransHab / BEAM Shell Architecture

(five functional sub-assemblies, not to scale)

Exterior (space environment)

AO protection cover

(Beta cloth)
Atomic oxygen barrier

4. Thermal Protection

MLI blankets
(Kapton + Dacron)
Thermal insulation

Sub-assembly

MMOD: Nextel + Kevlar

(Stuffed Whipple)
Projectile fragmentation

3. MMOD Shield

Sub-assembly

MMOD: Open-cell foam

(Solimide)
Energy absorption

Restraint layer
(Vectran webbing)
Primary structural element

2. Restraint
Sub-assembly

Bladder
(Combitherm film)
Gas barrier

Liner
(Nomex/Kevlar felt)
Crew-contact surface

1. Liner

Interior (pressurised)

Figure 5: TransHab/BEAM multi-layer shell architecture, showing the ve functional sub-
assemblies from the crew-contact liner (innermost) to the atomic oxygen protection cover
(outermost). The restraint layer (Vectran basket-weave) carries all pressure loads; the blad-
der, MMOD shield, and thermal protection system are non-structural. Total deployed thick-
ness: 3050 cm; total number of individual layers: 60+.

5.3
Rigidization Technologies
1062

While habitats remain pressure-stabilised throughout their operational life (at a factor
1063

of safety of 4.0), many inatable componentsparticularly booms, masts, and structural
1064

supportsrequire rigidisation after deployment to eliminate dependence on continued gas
1065

containment. Cadogan and Scarborough established the canonical classication of rigidisa-
1066

tion technologies into three families Cadogan and Scarborough [2001]:
1067

Mechanical (strain hardening): Aluminum-polymer laminates (e.g., 14.5 µm Al /
1068

16 µm Mylar / 14.5 µm Al) undergo plastic deformation during ination, work-hardening
1069

the aluminium layers and locking the deployed shape Schenk et al. [2014]. This approach
1070

has the longest ight heritage, from Echo 2 (1964) through InateSail (2017), where a 1 m
1071

strain-rigidized mast achieved deployment in approximately 2 seconds via CO2 pressurization
1072

Viquerat et al. [2019], Underwood et al. [2017].
Lenticular boom cross-sections achieve
1073

packaging ratios of approximately 10:1, while circular cross-sections achieve approximately
1074

5:1 under z-fold Schenk et al. [2014]. Current TRL: 89.
1075

Physical (sub-Tg and shape memory): Resin-impregnated composites heated above
1076

their glass transition temperature (Tg) become pliable for packaging; upon deployment and
1077

cooling below Tg in the space thermal environment, the resin solidies and rigidizes the struc-
1078

ture Cadogan and Scarborough [2001], Defoort et al. [2005]. This approach is reversible in
1079

principle, enabling re-stowage. Shape memory polymers extend this concept with engineered
1080

Tg transitions. Current TRL: 45.
1081

Chemical (UV-curable): Cationic epoxy resins cure upon exposure to solar UV radi-
1082

ation, achieving the highest post-rigidisation stiness of the three approaches Allred et al.
1083

[2002]. The Rigidization on Command (ROC) technology demonstrated by Adherent Tech-
1084

nologies achieves mechanical properties equivalent to thermally cured composites using sun-
1085

light alone Adherent Technologies Inc. [2001]. However, UV curing requires unobstructed
1086

solar access and is sensitive to shadowing by other spacecraft elements. Current TRL: 45.
1087

An emerging fourth approach uses shape memory alloy (SMA) elements integrated into
1088

inatable toroidal structures. Patel et al. developed an analytical framework for SMA-based
1089

rigidisation where NiTi alloy wires, embedded in the inatable wall and heated above their
1090

austenite nish temperature, contract and lock the deployed geometry Rastogi et al. [2024].
1091

This approach remains at the analytical stage (TRL 23) but oers the potential for active
1092


![Table 21](paper-23-v1_images/table_21.png)
*Table 21*

shape control during rigidisation.
1093

Table 11: Rigidization technology comparison for inatable space structures.

Method
Mechanism
TRL
Heritage
Best Application

Strain hardening
Al-polymer plastic deformation
89
Echo 2, InateSail
Thin booms, sails
Sub-Tg resin
Glass transition solidication
45
Ground demos
Structural booms
UV curing
Solar-initiated polymerization
45
Ground demos
Max. stiness booms
SMA rigidisation
Thermoelastic contraction
23
Analytical only
Toroidal structures

A critical distinction: large inatable habitats (BEAM, TransHab, LIFE) do not em-
1094

ploy rigidisation. They remain pressure-stabilised structures throughout their operational
1095

life, relying on the continuous pressure dierential across the multi-layer shell to maintain
1096

structural integrity at a factor of safety of 4.0 Valle et al. [2019b]. Rigidization is primar-
1097

ily relevant for booms, masts, and structural supports where prolonged gas containment is
1098

impractical or where a loss-of-pressure failure mode is unacceptable.
1099

5.4
Environmental Degradation: AO, UV, Radiation, Creep
1100

Four environmental mechanisms govern the long-term performance of inatable structures
1101

in the space environment, each aecting dierent layers of the shell assembly.
1102

Atomic oxygen (AO) is the dominant surface degradation threat in LEO. At ISS
1103

altitude (∼400 km), AO ux is approximately 1015 atoms/cm2/s, and Kapton H exhibits
1104

an erosion yield (Ey) of 3.0 × 10−24 cm3/atomthe practical erosion rate (thickness loss
1105

per unit time) is Ey × Φ, where Φ is the AO ux, which varies with altitude, solar activity,
1106

and ram direction; at ISS altitude this corresponds to approximately 1 µm/year Banks
1107

et al. [2004]. Unprotected Mylar, Kevlar, and Vectran all exhibit comparable erosion rates.
1108

SiO2 coatings reduce Kapton erosion by 23 orders of magnitude, and novel AO-resistant
1109

polymers (TOR, COR) developed at NASA Glenn demonstrate near-zero erosion Banks
1110

et al. [2004].
In practice, inatable habitats are protected by the outermost Beta cloth
1111

layer, which is inherently AO-resistant due to its PTFE coating. In-situ measurements from
1112

JAXA's SLATS satellite (160560 km altitude range) have recently provided direct on-orbit
1113

validation of erosion models Verker et al. [2023].
1114

UV degradation primarily aects Kevlar (discoloration and embrittlement) and Zylon
1115

(catastrophic strength loss of ∼35% in 6 months) Destefanis et al. [2009], Toyobo Co., Ltd.
1116

[2005]. Vectran shows moderate UV sensitivity. The multi-layer shell architecture naturally
1117

provides UV shielding for interior layers, but any externally exposed fabric elements require
1118

dedicated UV protection.
1119

Radiation eects on high-performance fabrics are comparatively modest for LEO mis-
1120

sions. The primary radiation concern for inatable habitats is crew dose rather than material
1121

degradationBEAM measurements during a September 2017 solar particle event recorded
1122

22.5 mGy inside BEAM versus approximately 0.25 mGy in adjacent ISS metallic modules,
1123

an 810× ratio attributable to the lower areal density of the fabric shell NASA Johnson Space
1124

Center [2017]. Polyethylene supplemental shielding oers 27.8% mass savings over equivalent
1125

aluminium shielding, and multi-layer congurations achieve up to 70% total ionizing dose
1126

improvement for electrons and 50% for protons Norbury et al. [2025].
1127

Creep is the life-limiting mechanism for Vectran and Kevlar restraint layers under sus-
1128

tained biaxial pressure loading. Weadon's characterisation demonstrated three-stage vis-
1129

coelastic creep with exponential sensitivity to the ratio of applied load to ultimate tensile
1130

strength Weadon [2013]. At the design operating point of 25% UTS (FOS = 4.0), specimens
1131

showed no failure over test periods of months. However, the wide manufacturing variability
1132

in UTS (±10%) dominates lifetime uncertaintynot the average material properties them-
1133

selves. Combined synergistic eects (AO + UV + thermal cycling + sustained load) remain
1134

inadequately characterised, representing a research gap that limits condence in multi-decade
1135

lifetime predictions for deep-space habitats Zhai et al. [2023].
1136

6
State of the Art: Deployment Mechanics
1137

The deployment of inatable structures in the space environment presents a unique engineer-
1138

ing challenge: a large, compliant membrane must transition from a compactly folded launch
1139

conguration to a precise deployed geometry under vacuum conditions where gas dynamics,
1140

thermal gradients, and material memory eects all inuence the nal state. This section
1141

reviews fold pattern selection, ination control strategies, and lessons from ight heritage.
1142

6.1
Fold Patterns and Packaging Eciency
1143

The choice of fold pattern determines deployment reliability, packaging eciency, and the
1144

number of actuators required for controlled deployment. Three primary pattern families are
1145

employed, each optimised for a dierent structural geometry.
1146

Miura-ori Miura [1985] is the foundational pattern for at membrane deployment. The
1147

tessellation of parallelogram facets creates a one-degree-of-freedom rigid-foldable mecha-
1148

nism: the entire membrane deploys via a single actuator force without requiring elastic
1149

deformation of the panels. This property is critical for fragile thin lms (metallized My-
1150

lar, ceramic-coated Kapton) that cannot sustain repeated fold stress. The negative Pois-
1151

son's ratio characteristiccontraction in one direction when extended in the perpendicular
1152

directionassists controlled deployment by preventing bunching Miura [1985]. Compaction
1153

is theoretically unlimited: an N × M panel array compacts to a stack of 2 panels thick,
1154

achieving compaction ratios of N/2 in each direction. Miura-ori is optimal for solar sails,
1155

antenna reectors, and drag sails where at-membrane deployment is required.
1156

For cylindrical structures (booms, masts), Schenk and Guest adapted the Miura-
1157

ori pattern to cylindrical geometry, enabling origami-based compaction of inatable booms
1158

with geometrically determined deployment kinematics Schenk et al. [2013, 2014]. The z-fold
1159

variant oers the simplest implementation and highest packaging ratio but lower deployment
1160

reliability, as individual folds must sequentially release without jamming. Wrapping (coiling)
1161

provides more controlled deployment at lower packaging ratios. The lenticular boom cross-
1162

section achieves ∼10:1 packaging ratios versus ∼5:1 for circular cross-sections Schenk et al.
1163

[2014].
1164

For habitats, a 7-gore S-fold approach is employed: the bladder and restraint layers
1165

are folded in an S-pattern around the rigid central core, with individual MMOD and MLI
1166

gore panels attached separately Kennedy [2016], Valle et al. [2019b]. The habitat packaging
1167

ratio is substantially lower than for membranes or booms because the rigid core occupies a
1168

signicant fraction of the stowed volume. TransHab achieved a stowed-to-deployed volume
1169

ratio of approximately 2.1:1 (habitable volume), while BEAM achieves approximately 4.4:1
1170

(16 m3 deployed / ∼3.6 m3 stowed) Valle et al. [2019b].
1171

6.2
Ination Sequencing and Control
1172

Ination rate control is critical for successful deployment: ination that is too rapid generates
1173

shock waves in the gas column that can damage thin lms and cause asymmetric expansion,
1174

while ination that is too slow allows thermal gradients to develop that aect the nal
1175

geometry Jenkins [2001]. Minimum tension requirements must be maintained throughout
1176


![Table 22](paper-23-v1_images/table_22.png)
*Table 22*

BEAM
(habitat)

400:1

TransHab

340:1

(habitat)


![Table 23](paper-23-v1_images/table_23.png)
*Table 23*

LOFTID
(aeroshell)

Inflatable

180:1

InflateSail
(sail boom)

150:1

PowerSphere

120:1

(power)

ROSA
(solar array)

20:1

TRAC boom

50:1

(structural)


![Table 24](paper-23-v1_images/table_24.png)
*Table 24*

Rigid
deployable

Mesh reflector

30:1

(antenna)

Coilable mast

15:1

(structural)

Hinged panel

8:1

(solar array)

0
100
200
300
400
Deployed-to-stowed volume ratio

Figure 6: Deployed-to-stowed volume ratio comparison between inatable and rigid deploy-
able structures. Inatable systems achieve packaging ratios an order of magnitude higher
than rigid deployable alternatives, with BEAM demonstrating a 400:1 ratio. Data compiled
from mission documentation and manufacturer specications.


![Table 25](paper-23-v1_images/table_25.png)
*Table 25*

Table 12: Packaging eciency by structure type for inatable space systems.

Structure Type
Fold Pattern
Packaging Ratio
Heritage Example

Flat membrane (sail)
Miura-ori / z-fold
∼500:1 (membrane)
InateSail (10 m2)
Boom (lenticular)
Origami / coil
∼10:1
InateSail (1 m boom)
Boom (circular)
z-fold
∼5:1
Various CubeSat booms
Habitat (with rigid core)
7-gore S-fold
25:1
BEAM (∼4.4:1), TransHab
Origami shield
Waterbomb tessellation
∼5:1 (80% expansion)
IMSS concept Cha et al. [20

ination to prevent wrinkling, which can create permanent creases in metallized lms and
1177

compromise thermal or RF performance.
1178

The BEAM deployment sequence provides the most instructive ight data on ination
1179

control challenges. Initial deployment in May 2016 failed to expand BEAM beyond a small
1180

fraction of its intended volume. Over the following 7 hours, mission controllers executed
1181

25 sequential pressure bursts, each providing a small increment of expansion, before BEAM
1182

reached its full deployed geometry NASA Johnson Space Center [2017]. This arduous recov-
1183

ery illustrates a fundamental tension: the folded softgoods develop stronger memory eects
1184

during extended stowed periods than ground testing predicted, requiring more expansion
1185

energy than designed. For autonomous missions (lunar surface habitats, Mars transit mod-
1186

ules), such manual intervention is not viable, and deployment reliability must be established
1187

at substantially higher condence levels Valle et al. [2019b].
1188

Several ination methodologies have been demonstrated or proposed. Stored gas (typi-
1189

cally CO2 or N2) provides the most controllable ination but requires tanks, regulators, and
1190

plumbing that add mass and failure modes. InateSail used a cold-gas CO2 system for boom
1191

deployment Viquerat et al. [2019]. Sublimation-based ination eliminates gas handling
1192

hardware: benzoic acid or naphthalene powder generates sucient vapour pressure at ambi-
1193

ent space temperatures to inate simple structures, though residual air in the packed struc-
1194

ture can cause premature partial ination Horn [2017]. The PowerSphere concept employed
1195

passive vapour-pressure ination from sublimation powder for a multifunctional sphere Cado-
1196

gan et al. [2006]. Active pressure control using real-time pressure-volume feedback with
1197

variable ination rates has been studied analytically by Wei et al., who demonstrated that
1198

instantaneous optimal control of ination rate can minimise deployment loads and improve
1199

nal shape accuracy Li et al. [2022a].
1200

6.3
Flight Heritage: InateSail, LOFTID, BEAM Deployment Lessons
1201

Three ight demonstrations provide the primary deployment heritage for inatable struc-
1202

tures, each operating at a dierent scale and in a dierent deployment regime.
1203

InateSail (2017) demonstrated the most compact packaging and fastest deployment: a
1204

1 m aluminium-Mylar laminate boom (14.5 µm Al / 16 µm Mylar / 14.5 µm Al) and 10 m2
1205

aluminized Mylar drag sail packaged into a 0.5U volume (∼50 mm cube), deploying and
1206

strain-rigidizing in approximately 2 seconds via CO2 pressurization Viquerat et al. [2019].
1207

The deployed membrane-to-stowed volume ratio of approximately 500:1 represents the high-
1208

est documented packaging eciency for a complete deployable system. InateSail de-orbited
1209

from 505 km in 72 days, compared to an estimated 4+ years without the sail, validating the
1210

drag deorbit concept at TRL 89 Viquerat et al. [2019].
1211

IRVE-3 (Inatable Reentry Vehicle Experiment, 2012) demonstrated a 3 m diameter in-
1212

atable aeroshell surviving Mach 10 reentry with peak heating of 14.4 W/cm2 Hughes et al.
1213

[2005]. Its successor, LOFTID (Low-Earth Orbit Flight Test of an Inatable Decelerator,
1214

2022), scaled this concept to 6 m diameter and survived Mach 30 reentry, achieving TRL 89
1215

for inatable aerodynamic decelerators. These demonstrations establish the thermal protec-
1216

tion performance of exible fabric systems under extreme heating conditions, conrming that
1217

multi-layer woven ceramic and polymer fabrics can provide thermal protection comparable
1218

to rigid ablative shields at a fraction of the mass.
1219

BEAM (2016present) provides the denitive deployment lesson for large pressurised
1220

habitats. Beyond the 25-burst recovery described above, BEAM demonstrated that pack-
1221

aged softgoods develop adhesion between layers during extended stowage that signicantly in-
1222

creases deployment energy requirements NASA Johnson Space Center [2017]. Post-deployment,
1223

thermal performance was more benign than predicted because folded softgoods create addi-
1224

tional insulation beyond the designed MLI performance. BEAM has now operated on ISS for
1225

over 8 years, providing the most extensive in-service data for any inatable habitat. These
1226

deployment lessons directly inform the design of future autonomous systems: residual fold
1227

adhesion must be characterised and accounted for, deployment energy budgets must include
1228

substantial margin, and passive deployment mechanisms (sublimation, spring) may be more
1229

reliable than active pressurization for autonomous operations.
1230

6.4
Comparison with Rigid Deployable Alternatives
1231

The survey's thesisthat inatables oer advantages over rigid systemsrequires adequate
1232

characterisation of the rigid deployable baseline. Three competing technology classes merit
1233

explicit comparison.
1234

Composite booms (e.g., CFRP bi-stable tape springs, Triangular Rollable and Col-
1235

lapsible (TRAC) booms) achieve packaging ratios exceeding 50:1 and are ight-proven at
1236

TRL 9 Murphey et al. [2015], Banik and Murphey [2010].
The TRAC boom, used on
1237

LightSail-2 and the Aeroboom Innovative Mechanism (AIM), provides high deployed stiness
1238

with no ination requirement. Sickinger and Herbeck Sickinger and Herbeck [2004] charac-
1239

terised CFRP boom deployment for solar sails, demonstrating that non-inatable composite
1240

booms are the dominant competing technology for CubeSat-class deployables.
1241

Mesh reector antennas (e.g., Harris/L3Harris AstroMesh, 1222 m deployed diam-
1242

eter, TRL 9) achieve large deployed apertures through cable-net tensioned mesh without
1243

requiring ination Santiago-Prowald and Rodrigues [2018]. These are the primary competi-
1244

tor to inatable antenna concepts and represent the state of the art for deployable high-gain
1245

antennas.
1246

Mechanically hinged trusses (e.g., NASA Langley's Compact Telescoping Array,
1247


![Table 26](paper-23-v1_images/table_26.png)
*Table 26*

CIRAS) provide high stiness and precise geometry through articulated rigid elements, at
1248

the cost of higher mass and complexity compared to inatable deployment.
1249


![Table 27](paper-23-v1_images/table_27.png)
*Table 27*

Table 13 presents a comparative assessment.
1250

Table 13: Comparison of inatable and rigid deployable technologies.

Technology
Pkg Ratio
Deployed Sti.
Mass/m
TRL
Key Limitation

TRAC composite boom
50100:1
High
Low
9
Length <10 m
AstroMesh reector
1020:1
High
Medium
9
Complex cable-net
Mech. hinged truss
310:1
Very high
High
9
Mass, complexity
Inatable boom (Al-lam.)
510:1
Med. (post-rigid.)
Very low
89
Rigidisation req'd
Inatable membrane
100500:1
Low (press.-stab.)
Very low
79
Pressure maint.

The inatable approach oers its greatest advantage at the largest scales (>10 m), where
1251

composite boom stiness-to-length scaling becomes unfavourable and mesh reector cable-
1252

net complexity grows prohibitively. For CubeSat-class deployables (<3 m), TRAC booms
1253

are the dominant technology; for medium-scale antennas (522 m), mesh reectors compete
1254

strongly. Inatables become uniquely enabling above approximately 30 m, where no rigid
1255

deployable alternative exists at acceptable mass.
1256

7
State of the Art: Actuation for Soft Space Systems
1257

The space environment imposes four principal constraints on actuator selection for soft in-
1258

atable systems: (1) ultrahigh vacuum eliminates ambient pressure support for unsealed
1259

pneumatic systems; (2) extreme temperature cycling (−150◦C to +120◦C in LEO) chal-
1260

lenges elastomers, smart materials, and ionic actuators; (3) high-energy particle and UV
1261

radiation degrades polymers, electrodes, and electrolytes; and (4) the absence of conven-
1262

tional lubricants eliminates standard gearing options. Against this backdrop, research has
1263

converged on several non-pneumatic actuation principles. This section reviews six technol-
1264

ogy families, organised from highest space-mission specicity to most novel, and presents a
1265

comparative assessment for inatable system integration.
1266

7.1
Dielectric Elastomer Actuators and DEMES
1267

Dielectric Elastomer Actuators (DEAs) convert high-voltage electrical input into mechanical
1268

deformation of a thin elastomer membrane sandwiched between compliant electrodes. Di-
1269

electric Elastomer Minimum Energy Structures (DEMES) extend this principle by bonding a
1270

pre-stretched DEA membrane to a exible frame, creating a self-deploying bending actuator
1271

that rolls compactly for stowage Araromi et al. [2014, 2015].
1272

The most mission-specic DEA application is the DEMES gripper developed by Araromi et al.
1273

for ESA's CleanSpace One microsatellite, targeting the 820 g SwissCube CubeSat for active
1274

debris removal Araromi et al. [2014]. The four-arm gripper achieves the following specica-
1275

tions: mass less than 0.65 g per arm, tip angle change of approximately 60◦, gripping force
1276

of 0.8 mN at 5 mm deection (up to 2.2 mN in optimised frame variants), and over 860,000
1277

actuation cycles at 1 Hz and 2000 V without degradation. The actuator stores rolled to a
1278

14 mm diameter cylinder and deploys by burning a retaining Nylon wire. A mechanically
1279

elegant property emerges from the force-displacement characteristic: grip force increases as
1280

the target drifts away from the actuator tip, creating a passive negative feedback loop that
1281

enhances capture stability without active control Araromi et al. [2014].
1282

Li et al. subsequently extended the 2D DEMES concept to a three-dimensional congura-
1283

tion specically designed for on-orbit servicing, enabling triaxial manipulation of irregularly
1284

shaped targets Liang et al. [2023]. The 3D conguration achieves higher load capacity and
1285

more favorable specic force output than planar DEMES.
1286

The critical limitation of DEA/DEMES for space applications is force output: the sub-
1287

millinewton to millinewton range, while sucient for microgravity contact-only operations
1288

on CubeSat-class targets, is inadequate for structural loads or capture of debris exceeding
1289

a few kilograms. DEA membranes (PDMS, acrylic) are also vulnerable to outgassing in
1290

vacuum and UV degradation, though neither has been systematically quantied under space
1291

conditionsa notable gap.
1292

7.2
Vacuum-Gap Electrostatic Actuators: Vacuum as Enabler
1293

A paradigm-shifting development emerged in 2025 with Sîrbu et al.'s introduction of vacuum-
1294

gap electrostatic multilayer actuators Sirbu et al. [2025]. These devices use thin-lm polymer
1295

multilayer structures enclosing vacuum gaps that zip closed upon electrical activationa
1296

mechanism that fundamentally benets from, rather than suers from, the space vacuum.
1297

In terrestrial operation, the vacuum gaps must be maintained against atmospheric pressure;
1298

in space, the ambient ultrahigh vacuum (∼10−7 Pa in LEO) is the default state.
1299

The performance specications represent a qualitative advance over existing soft actuator
1300

technologies: actuators weighing 0.7 g deliver forces exceeding 4 N, operate at bandwidths
1301

above 100 Hz, and achieve specic power of 1.4 kW/kg Sirbu et al. [2025]. For comparison,
1302

DEMES achieves 0.82.2 mN force at comparable massvacuum-gap actuators thus exceed
1303

DEA performance by three orders of magnitude in force at the same mass scale. The gearless,
1304

lubricant-free construction eliminates two major space reliability concerns.
1305

(a) Terrestrial operation

(b) Space operation

Atmospheric pressure opposes vacuum gap

Ambient vacuum provides functional gap

Electrode (+)

Electrode (+)

Fe
V

Fe
V

Vacuum gap

Vacuum gap

(pumped)

(ambient)

Flexible
membrane

Flexible
membrane

Electrode (-)

Electrode (-)

Ambient vacuum = functional gap

Atmospheric pressure

must be overcome

No pump required

1 atm

~0 Pa (vacuum)

Sirbu et al. 2025: 0.7 g, >4 N force, >100 Hz bandwidth, specific power 614 W/kg

Figure 7: Vacuum-gap electrostatic actuator operating principle (after Sirbu et al. 2025 Sirbu
et al. [2025]). (a) In terrestrial operation, vacuum gaps between electrodes must be main-
tained against atmospheric pressure, requiring a vacuum pump. (b) In space, the ambient
vacuum provides the functional dielectric gap directly, eliminating the pump and enabling
higher bandwidth (>100 Hz) at extremely low mass (0.7 g, >4 N, specic power 614 W/kg).

The thin-lm polymer construction of vacuum-gap actuators is structurally analogous
1306

to the multilayer membrane systems already used in inatable habitat construction. The
1307

possibility of laminating vacuum-gap actuator layers to the inner liner of an inatable robotic
1308

arm, combined with bre optic shape sensors woven into the restraint webbing, suggests
1309

a pathway toward fully sensorized, actively controlled inatable manipulatorsa system
1310

architecture not yet demonstrated in the literature. The primary unresolved qualication
1311

gaps are thermal cycling (−150◦C to +120◦C), radiation tolerance, and scale-up beyond the
1312

current laboratory-scale prototypes.
1313

7.3
Ionic Electroactive Polymers: Space Tolerance Assessment
1314

Ionic electroactive polymer (IEAP) actuators operate through ion migration within a polymer
1315

membrane, producing bending deformation at low voltages (15 V). Punning et al. conducted
1316

the only systematic, large-sample space environment tolerance study for this actuator class,
1317

testing 320 samples across 7 IEAP material types under six space-relevant conditions: X-ray
1318

irradiation (167.4 Gy), gamma irradiation (2036 Gy from 60Co), UV exposure (180 hours,
1319

xenon lamp), vacuum (<1 mbar, 2 weeks), and cryogenic storage at 77 K (liquid N2, 2 weeks)
1320

and 4.22 K (liquid He) Punning et al. [2014].
1321

The results establish three design rules for space IEAP deployment:
1322

1. Use ionic liquid electrolytes: IEAP types employing ionic liquid (IL) electrolytes
1323

(EMIBF4, EMITF, EMITFSI) showed no notable degradation under vacuum or cryo-
1324

genic conditions. Aqueous IPMC actuators (Type A) dry out in vacuum, requiring
1325

encapsulation for space use.
1326

2. Provide UV shielding for external applications: UV irradiation destroys PE-
1327

DOT and PEO-based IEAP materials via photo-oxidation. This is the primary space
1328

environment threat. Materials using carbonaceous or conducting polymer electrodes
1329

with ionic liquid electrolytes (Types B, C, G) survived UV testing with no notable
1330

eect.
1331

3. Plan for cryogenic dormancy: All tested IEAP types survived cryogenic stor-
1332

age (77 K for 2 weeks, 4.22 K for 15 minutes) and recovered full functionality upon
1333

warmingthe materials cannot operate while frozen but survive and revive Punning
1334

et al. [2014].
1335

A counter-intuitive nding is that X-ray radiation initially increases IEAP performance
1336

through radiation-induced doping of conducting polymer electrodes, an eect that normalizes
1337

within a few actuation cycles Punning et al. [2014]. The force output of current IEAPs
1338

remains in the low-millinewton range, limiting applications to sensing-adjacent tasks and
1339

micro-manipulation.
1340

7.4
Tendon-Driven Continuum Manipulators
1341

Tendon-driven continuum manipulators represent the highest-force soft actuation approach
1342

compatible with space constraints. NASA's Tendril robot (Mehling et al., 2006) established
1343

the heritage origin: a 1:1000 aspect-ratio inspection robot designed for conned-space inspec-
1344

tion inside the Space Shuttle external tank Mehling et al. [2006]. The Tendril architecture
1345

multiple antagonistic tendons routed along a compliant backboneprovides both the force
1346

density and bandwidth necessary for structural manipulation tasks.
1347

Ouyang et al. proposed a hybrid rigid-continuum dual-arm space robot combining a rigid
1348

primary arm for strength and reach with a continuum secondary arm for dexterity and com-
1349

pliance Ouyang et al. [2021]. The Generalized Jacobian Matrix analysis demonstrated coor-
1350

dinated motion planning for free-oating operations, establishing the mathematical frame-
1351

work for hybrid architectures where inatable continuum arms complement rigid primary
1352

manipulators.
1353

For space-compatible tendon routing, MoS2 solid lubricant enables vacuum-compatible
1354

sliding contacts, addressing the lubrication challenge that would otherwise limit tendon-
1355

driven systems to short operational lifetimes Ruíz et al. [2023].
The primary limitation
1356

of tendon-driven approaches is that routing tendons over long lengths (>1 m) introduces
1357

increasing friction and hysteresis, requiring careful mechanical design.
1358

7.5
Shape Memory Alloys for Deployment
1359

Shape memory alloys (SMAs), principally NiTi (Nitinol), have the highest ight TRL (8
1360

9) among actuator technologies applicable to soft inatable systems, though primarily for
1361

one-shot deployment rather than cyclic actuation. Nitinol achieves up to 10% recoverable
1362

strain and cycle life up to 600,000 activation cycles under controlled conditions Costanza and
1363

Tata [2020]. Space heritage includes Mars Pathnder deployment hinges, numerous CubeSat
1364

solar array release mechanisms, and ESA satellite solar array root hinges Costanza and Tata
1365

[2020], Blanc et al. [2013].
1366

For inatable structures specically, the critical limitation of SMA is its slow cooling
1367

rate in the vacuum thermal environment. Without convective cooling, SMA actuators rely
1368

on radiative heat transfer alone, limiting cyclic actuation frequency to well below 1 Hz for
1369

typical element sizes. This eectively restricts SMA to single-deployment or low-frequency
1370

repositioning applications in space.
1371

An emerging application combines SMA with inatable structures: Patel et al. devel-
1372

oped an analytical framework for SMA-based rigidisation of inatable toroidal structures,
1373

where NiTi wires embedded in the inatable wall contract upon heating to lock the deployed
1374

geometry Rastogi et al. [2024]. This represents a potential fourth rigidisation approach be-
1375

yond the three families established by Cadogan and Scarborough Cadogan and Scarborough
1376

[2001], though it remains at the analytical stage (TRL 23).
1377

7.6
Jamming in Vacuum: A Novel Opportunity
1378

Variable stiness by granular or layer jamming presents a counter-intuitive advantage in
1379

the space environment that has not been previously identied in the literature. In terres-
1380

trial soft robotics, jamming requires a dedicated vacuum pump to evacuate the jammed
1381

medium's enclosure, with external atmospheric pressure (∼101 kPa) providing the conning
1382

force Fitzgerald et al. [2020]. Zhang et al. noted that jamming structures are more likely
1383

to be used in soft space robots because of scalability, easy fabrication, and low cost Zhang
1384

et al. [2023c], but did not explore the vacuum-specic advantage.
1385

In the space environment, this constraint inverts: the ambient vacuum of LEO (∼10−7 Pa)
1386

serves as the external conning medium, while an inatable structure's internal pressuriza-
1387

tion (∼100 kPa) provides the pressure dierential across the membrane wall. A sealed jam-
1388

ming structure integrated into or attached to a pressurised inatable therefore achieves sti-
1389

ness modulation without any vacuum pumpa simplication unavailable on Earth. Layer
1390

jamming, which achieves stiness ratios exceeding 25:1 in terrestrial systems Fitzgerald et al.
1391

[2020], could be particularly well-suited for variable-stiness robotic elements embedded in
1392

inatable arms.
1393

(a) Terrestrial jamming

(b) Space jamming

Requires vacuum pump

Ambient vacuum = confining pressure

Patm = 101.3 kPa

Pvacuum 
 0 Pa

(external confining pressure)

(ambient space vacuum)

Granular

Granular

medium
(particles)

medium
(particles)

No pump

Vacuum

Inflatable structure

needed

pump

(internal pressure)

1 atm

~0 Pa

Evacuates interior

Stiffness transition: compliant (unjammed) to rigid (jammed) via pressure differential

Figure 8: Jamming-in-vacuum principle for variable stiness in space. (a) Terrestrial cong-
uration: a vacuum pump evacuates the sealed granular membrane, and atmospheric pressure
(∼101 kPa) provides the external conning force that locks the particles. (b) Space congu-
ration: the ambient space vacuum provides external conning pressure directly; the internal
pressurisation of the host inatable structure provides the pressure dierential. The vacuum
pump is eliminated, and the stiness transition from compliant to rigid is achieved passively.

The primary engineering challenges are: (1) selecting space-compatible granular media
1394

that do not outgas (candidates include hollow glass microspheres and sintered ceramic gran-
1395

ules); (2) maintaining gas-tight seals over mission duration against micrometeoroid puncture;
1396

and (3) characterising friction behaviour of jammed interfaces in vacuum, where the absence
1397

of adsorbed water layers may alter surface friction coecients. This insight represents a logi-
1398

cal deduction from known physics and inatable structure operating principles, and requires
1399

experimental validationa 5-year research priority identied in Section 13.3.
1400

7.7
Sealed Pneumatic Actuation in Space
1401

The opening constraint of this sectionthat ultrahigh vacuum eliminates ambient pressure
1402

support for unsealed pneumatic systemsdoes not preclude sealed pneumatic actuators that
1403

carry their own gas supply. BEAM itself is the supreme example of a sealed pneumatic
1404

structure in space. Ataka et al. Ataka et al. [2020] demonstrated a closed-loop pneumatic
1405

eversion arm with observer-based control that is directly relevant to inatable continuum
1406

manipulators for space inspection tasks. Eversion robots, which navigate their environment
1407

through growth by turning inside-out Hawkes et al. [2017], are particularly promising for
1408

space applications because the growth mechanism inherently manages the gas supply within
1409

the extending structure.
1410

Sealed pneumatic actuation with onboard gas storage is viable for missions where the total
1411

number of actuation cycles is bounded (limiting gas consumption) or where the inatable
1412

structure's own pressurisation system can serve as the gas source. The mass penalty of gas
1413

storageapproximately 0.52 kg per litre at 200 bar, depending on tank technologymakes
1414

this approach less competitive for sustained cyclic actuation than electrical alternatives, but
1415

appropriate for deployment and one-shot or low-cycle capture operations.
1416

7.8
Electroadhesion and Magnetic Actuation: Emerging Approaches
1417

Two additional actuation families, while not yet proposed for space inatable systems, merit
1418

assessment for completeness.
1419

Electroadhesion (electrostatic adhesion to a target surface) diers from the vacuum-gap
1420

actuators of Section 7.2 in operating principle: Coulombic attraction to an external target
1421

surface rather than internal gap zipping. Guo et al. Guo et al. [2020] demonstrated elec-
1422

troadhesion pads integrated with soft robotic grippers for manipulation of non-cooperative
1423

surfaces, achieving adhesion pressures of 15 kPa on conductive substrates. For debris cap-
1424

ture on metallic spacecraft surfaces, electroadhesion oers a contactless-force alternative to
1425

mechanical grasping. The primary space qualication gaps are dielectric breakdown in par-
1426

tial vacuum (outgassing-induced), surface contamination from space debris, and radiation
1427

degradation of the dielectric layer.
1428

Magnetic soft actuators with programmed 3D magnetisation proles Kim et al. [2018]
1429

represent a fundamentally dierent approach that avoids the vacuum and temperature limi-
1430

tations of pneumatics and elastomers. While not yet proposed for space, magnetic actuation
1431

in the eld-free environment of orbit would require onboard eld sources (permanent magnets
1432

or electromagnets), adding mass but eliminating the outgassing and embrittlement concerns
1433

of polymer-based actuators. This approach remains at TRL 2 for space applications.
1434


![Table 28](paper-23-v1_images/table_28.png)
*Table 28*

Table 14 presents a comparative assessment of the nine actuation technologies assessed
1435


![Table 29](paper-23-v1_images/table_29.png)
*Table 29*

for inatable space systems.
1436

Table 14: Actuator technology comparison for soft inatable space systems.

Technology
Force
Speed
Mass
TRL
Critical Space Gap
(Space)

DEA/DEMES
0.82.2 mN
∼1 Hz
<0.65 g
34
UV, outgas., low force
Vacuum-gap electrost.
>4 N
>100 Hz
0.7 g
34
Radiation, thermal
IL-IEAP (types B,C)
Very low
Medium
Excellent
34
UV (shield), frozen op.
Tendon-driven
High
High
Good
56
Long-path friction
SMA (one-shot)
Medium
Slow
Low
89
Slow cooling, fatigue
Jamming (layer)
Stiness only
Medium
Good
23
Unvalidated in vacuum
Sealed pneumatic
High
Medium
Mod. (gas)
45
Gas supply mass
Electroadhesion
15 kPa
Fast
Low
23
Surface contam., diel. brkdn
Magnetic (programmed)
Medium
Fast
Mod. (magnet)
12
Requires onboard eld

8
State of the Art: Sensing and Structural Health Mon-
1437

itoring
1438

Structural health monitoring (SHM) for inatable space structures must address three simul-
1439

taneous requirements: detection of micrometeoroid and orbital debris (MMOD) impacts that
1440

may compromise pressure integrity, continuous monitoring of creep deformation in restraint
1441

layers under sustained pressure loading, and shape sensing for actively controlled inatable
1442

manipulators. Fibre Bragg Grating (FBG) sensors have emerged as the leading technology
1443

platform for all three functions, with a coherent maturation pathway from rigid spacecraft
1444

heritage through soft actuator integration to inatable habitat application.
1445

8.1
Fibre Bragg Grating Sensors: From Proba-2 to Inatable Web-
1446

bing
1447

The FBG sensing principlewavelength-selective reection from a periodic refractive index
1448

modulation inscribed in a bre coreenables wavelength-division multiplexing (WDM) and
1449

time-division multiplexing (TDM) of large sensor arrays on a single bre strand. A single
1450

bre can carry 100+ independent FBG sensors, each at a distinct Bragg wavelength, pro-
1451

viding distributed strain and temperature measurement with no electrical connections at
1452

the measurement points Mckenzie et al. [2021]. Temperature sensitivity is approximately
1453

10 pm/◦C in the 15001600 nm wavelength range.
1454

ESA's 20+ year investment in bre optic sensing for spacecraft culminated in the Fiber
1455

Sensor Demonstrator (FSD) aboard Proba-2, launched in November 2009the rst bre
1456

optic sensor network demonstrated in the space environment Mckenzie et al. [2021]. The
1457

FSD incorporated 12 temperature sensors, a high-temperature thruster sensor, and a xenon
1458

tank pressure sensor, establishing TRL 78 for FBG technology on rigid spacecraft platforms.
1459


![Table 30](paper-23-v1_images/table_30.png)
*Table 30*

DEA / DEMES

Vacuum-gap
electrostatic

Tendon-driven

SMA (deployment)

Sealed pneumatic

Jamming
(vacuum-enabled)

Electroadhesion

Space TRL
Force output
Bandwidth
Vacuum
compatibility
Mass
efficiency

IEAP / IPMC

TRL 6
(flight
qualified)

0
2
4
6
8
10
Rating (0 = lowest, 10 = highest)

Figure 9: Comparative assessment of actuation technologies for soft inatable space sys-
tems across ve performance dimensions: space TRL, force output, bandwidth, vacuum
compatibility, and mass eciency. Ratings on a 010 scale reect the combined evidence
from literature reviewed in Sections 7.17.8. Vacuum-gap electrostatic actuators Sirbu et al.
[2025] and jamming Fitzgerald et al. [2020] score highest on vacuum compatibility, reecting
the vacuum as enabler paradigm shift.

Radiation tolerance of appropriately selected bre types (nitrogen-doped, uorine-doped) has
1460

been conrmed through ground testing, with Type II and Type III FBGs showing the highest
1461

radiation hardness Girard et al. [2022], Baba et al. [2025].
1462

The critical transition from rigid spacecraft to inatable structures was demonstrated by
1463

Bally Ribbon Mills (BRM) and Luna Innovations under a NASA SBIR program Bally Ribbon
1464

Mills and Luna Innovations [2020]. High-Denition Fibre Optic Sensing (HD-FOS) elements
1465

were woven directly into Vectran structural restraint webbing during the manufacturing
1466

processnot bonded after fabrication. Testing on 0.61 m and 2.74 m (1/3-scale) inatable
1467

habitat test articles at NASA Johnson Space Center demonstrated detection of:
1468

 Creep deformation under sustained pressure loading
1469

 Internal pressure changes during ination and operational cycling
1470

 Micrometeoroid impact events (conrmed via hypervelocity impact testing on inated
1471

articles)
1472

The partnership included NASA, Sierra Nevada Corporation, ILC Dover, BRM, and Luna
1473

Innovations, targeting applications for the Lunar Gateway and Mars transit habitats Bally
1474

Ribbon Mills and Luna Innovations [2020]. However, these results have been reported only
1475

in technical briefs and SBIR documentation, not in peer-reviewed publicationsa gap that
1476

limits independent assessment of sensitivity metrics, minimum detectable impact size, and
1477

long-term reliability.
1478

The TRL assessment for FBG sensing across application domains is:
1479

 FBG on rigid spacecraft: TRL 78 (Proba-2 FSD ight heritage, 2009)
1480

 FBG in Vectran restraint webbing: TRL 45 (NASA JSC ground testing, 0.61 m and
1481

2.74 m articles)
1482

 FBG in operational inatable habitat (ight): TRL 23 (not yet demonstrated)
1483


![Table 31](paper-23-v1_images/table_31.png)
*Table 31*

Table 15: Sensing technology comparison for inatable structural health monitoring.

Technology
Accuracy
Channels
Space
Demo
TRL
/Fiber
Heritage
Scale

FBG (rigid s/c)
±10 µε / ±1◦C
100+
Proba-2 (2009)
Satellite
78
FBG (Vectran webbing)
Creep/MMOD det.
Multiple
JSC ground
2.74 m
45
Multicore FOSS
0.64 mm tip
Multicore
Lab only
Actuator
34
DFOS (Rayleigh)
±1 µε
Continuous
Lab only
m-scale
23
Capacitive (stretch.)
±5% strain
Per-sensor
Lab only
cm-scale
23
Resistive (fabric)
±2% strain
Per-sensor
Lab only
cm-scale
23
Piezoelectric (PVDF)
Impact detection
Array
Lab only
Panel
23

8.2
Multicore Fibre Optic Shape Sensing
1484

For soft actuator shape sensing, Galloway et al. demonstrated the rst integration of a
1485

monolithic multicore Fibre Optic Shape Sensor (FOSS) into a bre-reinforced soft pneumatic
1486

actuator Galloway et al. [2019]. The multicore bre contains multiple sensing cores within a
1487

single cladding, enabling three-dimensional shape reconstruction from dierential curvature
1488

measurements without requiring multiple separate bre installations. Key results include a
1489

mean tip position error of 0.64 mm, successful reconstruction of six distinct planar shape
1490

proles, and simultaneous detection of collision events, environmental shape changes, and
1491

material stiness variations within a single sensing modality.
1492

The eld has advanced signicantly since Galloway's initial demonstration. Paloschi et al. Paloschi
1493

et al. [2020] developed improved 3D shape reconstruction algorithms for multicore optical
1494

bres, comparing transformation matrix approaches with Frenet-Serret equations for real-
1495

time applications and demonstrating that transformation matrix methods achieve superior
1496

accuracy for large-curvature deformations characteristic of soft actuators. Sefati et al. Sefati
1497

et al. [2021] demonstrated real-time 3D shape reconstruction using multicore FBG sensors
1498

on continuum robot manipulators, achieving sub-millimetre accuracy relevant to the tendon-
1499

driven continuum arms discussed in Section 7.4. These advances collectively bring multicore
1500

FOSS from a proof-of-concept to a viable shape sensing modality for soft space manipulators,
1501

though the interrogator hardware miniaturisation and radiation tolerance gaps remain.
1502

The multicore FOSS approach oers two advantages over distributed single-core FBG
1503

arrays for soft structure applications. First, the monolithic construction eliminates the need
1504

for multiple bre routing paths through complex soft geometries. Second, the dierential
1505

curvature measurement provides inherent common-mode rejection of temperature-induced
1506

wavelength shifts, improving strain measurement accuracy in the thermally variable space en-
1507

vironment. The primary barriers to space qualication are the mass and power requirements
1508

of the multicore FOSS interrogator (readout) hardware, which has not yet been miniaturized
1509

for spacecraft integration, and the radiation tolerance of the multicore bre itself, which has
1510

not been characterised.
1511

For broader context, Rajan et al. Rajan et al. [2021] provide a comprehensive review of
1512

FBG sensors for structural health monitoring across aerospace applications, conrming that
1513

FBG-based SHM is the most mature bre optic sensing technology for spacecraft structures
1514

and identifying the key challenges for transitioning from rigid to exible substrates.
1515

8.3
Capacitive, Resistive, and Alternative Soft Sensors
1516

While FBG sensors dominate the space-qualied sensing landscape, alternative soft sensing
1517

technologies merit assessment for completeness. Zhang et al. Zhang et al. [2023a] devote
1518

signicant attention to stretchable capacitive sensors, resistive fabric sensors, and liquid
1519

metal strain sensors for soft space robots. The advantages of these technologies include: no
1520

requirement for specialised interrogator hardware (unlike FBG, which requires wavelength-
1521

swept laser sources), simpler integration into soft structures via printing or embedding, and
1522

lower per-sensor cost. However, for space applications, three signicant limitations arise:
1523

 Electromagnetic interference (EMI) sensitivity: Capacitive and resistive sensors
1524

operate in the electrical domain and are vulnerable to the charged particle environ-
1525

ment of LEO, solar radio bursts, and EMI from onboard electronics. FBG sensors,
1526

operating in the optical domain, are inherently immune to EMIa decisive advantage
1527

for spacecraft.
1528

 Radiation vulnerability: Liquid metal sensors (e.g., eutectic gallium-indium, EGaIn)
1529

and conductive polymer sensors have not been characterised for radiation tolerance.
1530

Ionising radiation can alter the resistivity of conductive polymers and the wetting prop-
1531

erties of liquid metals, degrading sensor calibration over mission-duration timescales.
1532

 Multiplexing limitations: A single optical bre can carry 100+ independent FBG
1533

sensors via wavelength-division multiplexing; achieving comparable channel density
1534

with electrical sensors requires extensive wiring harnesses that add mass and failure
1535

modes to exible structures.
1536

For inatable habitat applications, capacitive pressure sensors could complement FBG
1537

strain sensors by providing direct pressure measurement at locations inaccessible to bre
1538

routing. For soft robotic manipulators, resistive bend sensors oer simplicity advantages for
1539

prototype development, though FBG remains the preferred technology for ight systems.
1540

8.4
Distributed Fibre Optic Sensing: Rayleigh and Brillouin Scat-
1541

tering
1542

Distributed bre optic sensing (DFOS) by Rayleigh or Brillouin scattering provides contin-
1543

uous strain and temperature proles along the entire bre length, rather than at discrete
1544

FBG grating locations. Rayleigh-based DFOS (e.g., Luna Inc. ODiSI platform) achieves
1545

spatial resolution of approximately 0.65 mm with strain resolution better than ±1 µε, while
1546

Brillouin-based systems provide sensing over distances up to 100 km at reduced spatial res-
1547

olution (typically 0.51 m). For inatable habitats with large membrane areas requiring
1548

continuous monitoring (rather than point-by-point FBG interrogation), DFOS oers the
1549

potential for comprehensive strain mapping of the entire restraint layer from a single bre
1550

installation.
1551

The principal barriers to space deployment of DFOS are: (i) interrogator size, mass,
1552

and power (current laboratory DFOS systems exceed 10 kg and 50 W, compared to <2 kg
1553

and <10 W for space-grade FBG interrogators); (ii) sensitivity to bre bending loss, which
1554

is exacerbated by the tight bend radii in folded inatable structures during stowage; and
1555

(iii) the absence of any space-environment characterisation data. DFOS is assessed at TRL 2
1556

3 for space inatable applications, but its unique capability for continuous spatial coverage
1557

makes it a high-priority development target for large-scale habitat SHM systems.
1558

8.5
Distributed Impact Detection
1559

The Distributed Impact Detection System (DIDS) installed on BEAM represents the highest-
1560

TRL implementation of impact sensing for inatable habitats. DIDS uses distributed sensors
1561

to detect and locate MMOD impacts on the inatable shell, providing real-time structural
1562

integrity monitoring.
1563

Beyond the BEAM DIDS, two emerging approaches extend impact detection capabilities.
1564

The BRM/Luna FBG-in-Vectran-webbing system described in Section 8.1 detected hyperve-
1565

locity impacts during ground testing, with the woven integration providing inherent coverage
1566

of the restraint layer structural grid Bally Ribbon Mills and Luna Innovations [2020]. Sepa-
1567

rately, Liao et al. demonstrated on-demand fabrication of PVDF-trFE piezoelectric sensors
1568

via in-space manufacturing techniques, enabling the production of impact detection arrays
1569

directly on deployed inatable structures White et al. [2024]. This approach could address
1570

the challenge of instrumenting structures that are too large or complex to pre-instrument
1571

before launch.
1572

Wei et al. proposed a complementary SHM approach based on low-frequency vibration
1573

response characterisation of pressurised inatable structures, where changes in modal fre-
1574

quencies indicate structural degradation Li et al. [2022b]. This global monitoring approach
1575

could complement the local sensing provided by FBG arrays, together forming a hierarchical
1576

SHM architecture: global vibration monitoring for overall structural health assessment, and
1577

local FBG sensing for precise damage location and magnitude quantication.
1578

The pathway from current demonstrated capabilities to a ight-qualied inatable SHM
1579

system requires: (1) formal peer-reviewed publication of the BRM/Luna FBG-in-webbing
1580

results with full sensitivity characterisation; (2) space qualication of FOSS interrogator
1581

hardware (mass, power, radiation tolerance); (3) development of data fusion algorithms
1582

combining local FBG and global vibration sensing; and (4) a ight demonstration, potentially
1583

as an ISS external payload experiment, to bridge the TRL 45 to TRL 78 gap.
1584

9
State of the Art: Power Systems for Large Inatables
1585

The integration of electrical power generation with inatable space structures is a critical
1586

enabling challenge for large deployable platforms. Unlike rigid spacecraft, where solar ar-
1587

rays are mechanically decoupled from the primary structure, inatable systems present the
1588

possibilityand the engineering challengeof co-locating photovoltaic generation on the de-
1589

ployable membrane itself. This section reviews the exible solar array landscape, the singular
1590

attempt at inatable-power integration (PowerSphere), and energy storage considerations for
1591

mission architectures ranging from 100 m-class debris shields to inatable habitats.
1592

9.1
Flexible Solar Array Landscape: ROSA to Perovskite
1593

The current state of the art in exible solar arrays for space is dened by the Roll-Out Solar
1594

Array (ROSA), which achieved TRL 9 via ISS ight demonstration in June 2017 as part
1595

of the STP-H5 experiment Spence et al. [2018]. The demonstration unit (5.40 m × 1.67 m)
1596

deployed successfully using stored strain energy in carbon-bre-reinforced polymer (CFRP)
1597

slit-tube booms, requiring no motors. The subsequent production variant, iROSA, scaled
1598

to 6 m × 13.7 m wings generating over 28 kW per wing at beginning of life with XTJ Prime
1599

triple-junction cells at 30.7% eciency.
Six iROSA wings installed on the ISS between
1600

2021 and 2023 added over 120 kW of generation capacity. At system level (blanket plus
1601

booms, excluding spacecraft attachment hardware), ROSA achieves a specic power exceed-
1602

ing 100 W/kgapproximately 3.7× the legacy ISS silicon rigid-panel arrays at ∼27 W/kg
1603

Spence et al. [2018], Yan et al. [2025]. Critically, however, ROSA's exible photovoltaic
1604

blanket is deployed on rigid composite booms; the deployment mechanism is structurally
1605

distinct from inatable substrate concepts.
1606

Beyond ROSA, three deployment architectures compete for next-generation high-power
1607

arrays Yan et al. [2025]: (i) Z-fold accordion panels on a central mast, representing the
1608

ISS legacy approach at TRL 9; (ii) fan-fold blankets on deployable masts, exemplied by
1609

China's CST arrays on the Wentian laboratory module (2022), achieving approximately 4×
1610

the specic power of rigid baselines; and (iii) roll-out arrays (ROSA/iROSA class). Mega-
1611

ROSA and SOLAROSA concepts target 200500 W/kg for systems exceeding 100 kW, though
1612

these remain at TRL 45 Yan et al. [2025]. For very large arrays approaching the kilometre
1613

scale (Space Solar Power Station concepts), wireless power transmission between modules
1614

has been identied as a necessity Yan et al. [2025].
1615

A paradigm shift in exible photovoltaic technology is emerging from perovskite-based
1616

tandem cells. Lang et al. Lang et al. [2020] provided the critical nding that perovskite/CIGS
1617

(copper indium gallium selenide) tandem cells are radiation-hard, while perovskite/silicon
1618

heterojunction (SHJ) tandems are emphatically not. Under 68 MeV proton irradiation at a
1619

uence of 2 × 1012 p+/cm2 (equivalent to over 50 years at ISS altitude), perovskite/CIGS
1620

tandems retained approximately 85% of initial power conversion eciency, whereas per-
1621

ovskite/SHJ devices degraded catastrophically to ∼1% retention due to proton-induced deep
1622

trap states in the silicon bottom cell Lang et al. [2020]. The perovskite top cell itself was
1623

essentially unaected, with quasi-Fermi level splitting changing by only 0.004 eV. With ac-
1624

tive layers only 4.38 µm thick (2.8 mg/cm2), perovskite/CIGS achieves a specic power of
1625

7,400 W/kg at the active-layer level, or 2,100 W/kg when including a 25 µm exible poly-
1626

imide substrate Lang et al. [2020]. More recently, Kim et al. Kim et al. [2024] demonstrated
1627

23.6% ecient exible perovskite/CIGS tandems surviving 100,000 bending cycles with a
1628

specic power of approximately 6,150 W/kg at the cell level.
1629

These gures represent a 1060× improvement over ROSA's system-level specic power,
1630


![Table 32](paper-23-v1_images/table_32.png)
*Table 32*

though the comparison requires careful attention to system boundaries: cell-only gures
1631

exclude interconnects, encapsulant, wiring harness, and structural substrate, which collec-
1632

tively reduce specic power by a factor of 36× at the system level. Table 16 summarises
1633


![Table 33](paper-23-v1_images/table_33.png)
*Table 33*

the specic power versus TRL landscape across exible photovoltaic technologies.
1634

Table 16: Specic power versus TRL for exible photovoltaic technologies for space applica-
tions. Cell-only and system-level gures are distinguished where data are available.

Technology
Specic Power (W/kg)
Eciency (%)
TRL
Ref.

Legacy ISS SAW (rigid)
∼27 (system)
14
9
Spence et al. [20
ATK UltraFlex
∼150 (system)
2830
9

ROSA/iROSA
>100 (system)
30.7
9
Spence et al. [20
Mega-ROSA (target)
>200400
30.7
45
Yan et al. [202
Perovskite/CIGS (25 µm sub.)
2,100 (cell+sub.)
19.2
34
Lang et al. [202
Perovskite/CIGS (Kim 2024)
∼6,150 (cell)
23.6
34
Kim et al. [202
PowerSphere (a-Si, measured)
∼7 (system)
10
45
Cadogan et al. [2
PowerSphere (proj. w/ III-V)
∼85 (projected)
2730

Cadogan et al. [2


![Table 34](paper-23-v1_images/table_34.png)
*Table 34*

Category / Cell efficiency

1400

Target region:
high specific power

= 12%

Heritage

Thin film

= 25%

+ flight-qualified

Perovskite
(single jxn)

Emerging

= 33%

1200

Inflatable

Specific power (W/kg)

1000

800

Perovskite/Si

tandem

600

CIGS thin-film

400

a-Si thin-film

200

PowerSphere

III-V MJ (rigid)
ROSA/iROSA

(integr.)

(III-V flex)

0

1
2
3
4
5
6
7
8
9
10
Technology Readiness Level (TRL)

Figure 10: Specic power versus technology readiness level for exible photovoltaic technolo-
gies relevant to inatable space structures. Marker size indicates cell eciency. Perovskite-
based technologies Lang et al. [2020], Kim et al. [2024] oer 1060× improvements over
heritage ROSA systems Spence et al. [2018] at the cell level, but remain at TRL 34. The
green shaded region indicates the target design space for next-generation inatable-power
integration: high specic power (>400 W/kg) at ight-qualied TRL (>6).

9.2
The Inatable-Power Integration Gap: PowerSphere and Be-
1635

yond
1636

The most direct attempt to integrate thin-lm photovoltaics with an inatable deployable
1637

structure was NASA's PowerSphere programme (20042009), led by ILC Dover (structure),
1638

NASA Glenn Research Center (cells), and Sandia National Laboratories (interconnects)
1639

Cadogan et al. [2003], Scheiman et al. [2005]. The PowerSphere Engineering Development
1640

Unit was a 0.6 m diameter UV-rigidisable inatable geodetic sphere clad with thin-lm amor-
1641

phous silicon (a-Si) solar cells on a polyimide substrate. The complete system comprised a
1642

1 kg PowerSphere subsystem mounted on a 3 kg bus, with 15 cells per hemisphere (9 hexag-
1643

onal, 6 pentagonal) connected via copper wrap-around ex-circuit interconnects that could
1644

survive folding during stowage without cracking Cadogan et al. [2003], Scheiman et al. [2005].
1645

The UV-rigidisation mechanism is particularly signicant for the survey's themes. Thirty
1646

hinges per sphere used S-glass bre reinforced with ATI-P600-2 UV-curing epoxy (glass tran-
1647

sition temperature Tg = 211 ◦C), encapsulated in UV-transparent 1-mil Mylar lm. Upon
1648

exposure to solar UV radiation (λ < 385 nm) for 1045 minutes post-deployment, the resin
1649

polymerised, converting the inatable into a self-supporting rigid structure and eliminating
1650

the requirement for long-term ination gas retention Cadogan et al. [2003]. Ination was
1651

achieved passively through vapour pressure from sublimation powder at approximately 1 psi
1652

(∼6.9 kPa).
1653

Thermal cycling tests (−120 ◦C to +80 ◦C, 1000 cycles per NASA specication) demon-
1654

strated cell and interconnect survival with less than 2% power degradation, although one
1655

of four interconnect coupons failed, prompting the addition of a titanium binder layer as a
1656

design modication. Cell interconnect technology was partially validated on the MISSE-5
1657

experiment aboard the ISS Cadogan et al. [2003]. At 10% a-Si cell eciency, the 0.6 m
1658

prototype generated approximately 29 W at design point, yielding a system specic power
1659

of ∼7.25 W/kg. With projected III-V triple-junction cells at 2730% eciency, the concept
1660

was estimated to reach ∼85 W/kg.
1661

The PowerSphere programme reached TRL 45 but never ew. Planned missionsthe
1662

PowerSphere Flight Experiment and PSIREX (Pico Satellite Inatable Reector Experiment)
1663

were not implemented, and the programme appears inactive since the nal publication by
1664

Jenkins in 2009 on thermal cycling results Jenkins et al. [2009]. No successor programme inte-
1665

grating thin-lm photovoltaics with inatable structure deployment has been identied. This
1666

represents a critical gap: ROSA (TRL 9) demonstrates that exible photovoltaic blankets
1667

function reliably in space, and PowerSphere (TRL 45) demonstrated that cells can survive
1668

fold/deploy on an inatable substrate, but nobody is currently pursuing inatable-integrated
1669

photovoltaics. A revival of the PowerSphere concept using modern perovskite/CIGS cells
1670

which oer 200300× higher specic power than the original a-Si cells and validated radiation
1671

hardness Lang et al. [2020]represents a logical and compelling research direction.
1672

9.3
Energy Storage: Li-ion, RFC, and Mission-Dependent Selection
1673

Energy storage for large inatable structures follows established space heritage, with tech-
1674

nology selection driven primarily by eclipse duration and mission architecture. The current
1675

standard is lithium-ion, with state-of-the-art cell-level specic energy of 200300 Wh/kg
1676

and system-level (including battery management, thermal control, and structure) of 100
1677

160 Wh/kg Sharma and Santasalo-Aarnio [2025]. The ISS lithium-ion upgrade programme
1678

(20172021), replacing nickel-hydrogen (Ni-H2) with 24 lithium-ion Orbital Replacement
1679

Units (ORUs) at 4 kWh each, provides direct heritage for large-structure lithium-ion energy
1680

storage.
1681

For a 100 m-class inatable debris shield in LEO (90-minute orbit, 36-minute eclipse),
1682

the power demand is driven by supporting subsystems rather than the passive membrane
1683

itself. Station-keeping via electric propulsion dominates at 150 kW depending on orbit and
1684

attitude strategy (Section 11.3); attitude control, telemetry, and sensors add 17 kW. A total
1685

system power demand in the range of 250 kW is appropriate, requiring 440 kWh of eclipse
1686

energy storagetranslating to 25250 kg of lithium-ion battery mass at 160 Wh/kg system
1687

level. This is a non-trivial but manageable fraction of the estimated 5,000 kg total system
1688

mass.
1689

For missions requiring extended eclipse storagenotably lunar surface operations (354-
1690

hour lunar night) or deep-space transitregenerative fuel cells (RFCs) oer 4001,000 Wh/kg
1691

at system level but remain at TRL 56 for space applications Sharma and Santasalo-Aarnio
1692

[2025]. Supercapacitors (515 Wh/kg) are poorly suited for eclipse energy storage but may
1693

serve pulsed-load applications such as electric propulsion ignition or deployment actuators.
1694


![Table 35](paper-23-v1_images/table_35.png)
*Table 35*


![Table 36](paper-23-v1_images/table_36.png)
*Table 36*

Table 17 summarises the energy storage technology comparison.
1695

Table 17: Energy storage technologies for large inatable space structures.

Technology
Sp. Energy (Wh/kg)
Cycle Life
TRL
Best Use Case

Li-ion (cell)
200300
>30,000
9
LEO eclipse storage
Li-ion (system)
100160
>30,000
9
LEO eclipse storage
Ni-H2 (legacy)
3060
>40,000
9
Heritage only
RFC (H2/O2)
4001,000

56
Lunar night, deep space
Supercapacitor
515
>500,000
7
Pulsed loads
RTG
N/A

9
No-sun environments

10
State of the Art: Thermal Management
1696

Thermal management for inatable space structures presents unique challenges that stem
1697

from the fundamental nature of the structural material: thin fabric membranes oer minimal
1698

thermal mass, poor through-thickness conductivity, and large surface area-to-volume ratios.
1699

These characteristics amplify the orbital thermal cycling environment and demand thermal
1700

control approaches that are compatible with the fold/deploy lifecycle, vacuum exposure, and
1701

the mechanical exibility of the host structure. This section reviews established approaches
1702

(multi-layer insulation, loop heat pipes), the JWST sunshield as a large-area deployable
1703

thermal barrier precedent, and emerging technologies (variable emissivity coatings, phase
1704

change materials) that oer particular promise for inatable applications.
1705

10.1
Multi-Layer Insulation for Inatable Shells
1706

Multi-layer insulation (MLI) is the primary passive thermal control technology for spacecraft
1707

and achieves eective emittance εe = 0.0050.05 for 1040 layer blankets Gilmore [2002],
1708

Finckenor and Dooling [1999]. For conventional rigid spacecraft, MLI is draped over external
1709

surfaces with controlled layer separation maintained by low-conductance spacers (typically
1710

Dacron netting). For inatable structures, MLI integration is more complex: the insulation
1711

must survive folding, accommodate deployment kinematics, and maintain layer separation
1712

without rigid structural support.
1713

The TransHab/BEAM heritage shell architecture represents the current standard for
1714

inatable habitat thermal design Kennedy [2002], Valle et al. [2019a]. In this architecture,
1715

MLI forms the outermost thermal protection sub-assembly of a ve-layer softgoods stack,
1716

ordered (outer to inner) as: (1) BETA cloth exterior for atomic oxygen protection; (2) nylon-
1717

reinforced double-aluminised Mylar/Kapton MLI layers with perforated inner surfaces for
1718

venting during deployment; (3) Nextel/Kevlar stued-Whipple MMOD shield; (4) Vectran
1719

restraint layer carrying hoop and axial pressure loads; and (5) multi-redundant gas-tight
1720

bladder. The MLI sub-assembly in TransHab comprised over 20 individual reector layers
1721

with eective emittance on the order of 0.0150.05 Finckenor and Dooling [1999].
1722

BEAM's on-orbit thermal performance has been characterised as more benign than
1723

predicted NASA Johnson Space Center [2017], an observation attributed to the additional
1724

insulation provided by folded softgoods layers that act as low-conductance barriers even
1725

when not specically designed as MLI. This nding has positive implications for inatable
1726

structure design: the inherent multi-layer nature of the fabric wall stack provides a degree
1727

of passive thermal buering beyond that of the dedicated MLI layers alone.
1728

10.2
The JWST Sunshield as Deployable Thermal Barrier Prece-
1729

dent
1730

The James Webb Space Telescope (JWST) sunshield is the largest deployed thermal barrier
1731

ever own and provides the benchmark for what large-area passive thermal isolation can
1732

achieve Arenberg et al. [2016]. At 21.2 m × 14.2 m (approximately 300 m2), the kite-shaped
1733

sunshield comprises ve layers of Kapton E polyimide membrane: Layer 1 (sun-facing) at
1734

50 µm thickness, Layers 25 at 25 µm. All layers are coated with 100 nm aluminium on both
1735

sides for reectivity; Layers 1 and 2 additionally carry 50 nm doped silicon on the sun-facing
1736

surface for enhanced emissivity and electrostatic discharge grounding.
1737

The thermal performance is extraordinary: the sun-facing side of Layer 1 reaches approx-
1738

imately +110 ◦C while the telescope-facing side of Layer 5 operates at −233 ◦Ca gradient
1739

of 343 ◦C across ve layers. Incoming solar power of approximately 200250 kW is attenu-
1740

ated to ∼23 mW transmitted to the cold side, an attenuation ratio of approximately 106:1
1741

Arenberg et al. [2016]. This performance is achieved through the V-groove geometry: angled
1742

layers radiate inter-layer thermal energy sideways to deep space through the vacuum gaps
1743

between membranes.
1744

However, the JWST sunshield is not an inatable structure. Layer separation is main-
1745


![Table 37](paper-23-v1_images/table_37.png)
*Table 37*

tained by six rigid spreader bars, with centre gaps of ∼2550 mm expanding to ∼250 mm
1746

at the edges. The deployment system required 139 of JWST's 178 release mechanisms, 400
1747

pulleys, 90 cables (∼0.5 km total), 8 motors, and 70 hinges Arenberg et al. [2016]. Table 18
1748


![Table 38](paper-23-v1_images/table_38.png)
*Table 38*

compares the JWST sunshield and TransHab shell architectures.
1749

Table 18: JWST sunshield versus TransHab inatable shell comparison.

Feature
JWST Sunshield
TransHab Shell

Primary function
Thermal isolation
Structural + MMOD + thermal
Layer count
5 membranes
5 sub-assemblies (60+ layers)
Layer material
Kapton E (all 5)
Vectran, Kevlar, Nextel, Mylar
Structural role
None (spreader bars)
Vectran restraint carries pressure
Energy attenuation
106:1
∼150 ◦C gradient
Deployment
139 mechanisms, 8 motors
Ination pressure
Deployed area
300 m2
220 m2 (cylinder)

It should be noted that JWST operates at the Sun-Earth L2 point, not in LEOthe
1750

thermal environment is fundamentally dierent (no orbital cycling, no atmospheric drag, no
1751

atomic oxygen), and this limits the direct applicability of JWST thermal performance num-
1752

bers to LEO inatable structures. Nevertheless, for inatable debris shields or large-area
1753

thermal barriers, the JWST heritage demonstrates that multi-layer Kapton stacks achieve
1754

extreme thermal gradients at 20+ metre scales. Adapting this concept to a fully inat-
1755

able deployment mechanismreplacing rigid spreader bars with ination-maintained layer
1756

separationremains an open engineering challenge. A hybrid approach combining inatable
1757

outer layers with rigid-bar-maintained inner separation represents a plausible intermediate
1758

architecture.
1759

10.3
Variable Emissivity Coatings and Smart Radiators
1760

Variable emissivity materials (VEMs) oer electronic louver functionality for dynamic ther-
1761

mal regulation without mechanical moving partsa capability uniquely suited to large inat-
1762

able surfaces where conventional mechanical louvers are impractical due to mass, complexity,
1763

and incompatibility with membrane substrates. Two technology families have received sus-
1764

tained development: passive thermochromic coatings and active electrochromic devices.
1765

Among passive thermochromic approaches, vanadium dioxide (VO2) based coatings are
1766

technically most advanced. Kim et al. Kim et al. [2019] performed the rst direct calorimetric
1767

measurement of a VO2-based switchable radiator in a simulated space environment (vacuum
1768

10−7 Torr, cold block at 108 K). Their multilayer structureSi substrate / VO2 (40100 nm)
1769

/ BaF2 dielectric spacer (1,500 nm) / Au back reector (200 nm)operates as a Fabry-Pérot
1770

resonant absorber.
In the low-temperature insulating state (T < 340 K), hemispherical
1771

emissivity is εL = 0.16; above the phase transition (T > 340 K, metallic VO2), εH = 0.51,
1772

yielding ∆ε = 0.35. The practical consequence is a net radiated power dierence of 480 W/m2
1773

between 300 K and 373 Ka factor of 7× in radiative cooling capacity Kim et al. [2019].
1774

The silicon substrate provides an incidental benet: protection of the VO2 lm from atomic
1775

oxygen erosion, addressing a known degradation mechanism. An earlier design by Hendaoui
1776

et al. Hendaoui et al. [2013] achieved a higher normal emissivity swing of ∆ε = 0.49 but
1777

without the atomic oxygen protection.
1778

The sole ight-demonstrated variable emissivity technology is the EclipseVEDTM elec-
1779

trochromic coating (Ashwin-Ushas Corporation), own on the MidSTAR-1 satellite in 2007,
1780

achieving TRL 78. EclipseVED operates by applying a low voltage (13 V) to an elec-
1781

trochromic polymer lm, switching emissivity across the range ε ≈0.190.90 in the 812 µm
1782

thermal infrared band. It requires no mechanical actuators, making it compatible with large-
1783

area application including inatable surfaces. The principal limitation is the requirement for
1784


![Table 39](paper-23-v1_images/table_39.png)
*Table 39*

a thin-lm conductor and electrical interconnects across the deployed areaa tractable but
1785

non-trivial integration challenge for inatable structures.
1786

Table 19 compares variable emissivity technologies.
1787

For the survey's inatable structures context, VEMs oer a path to autonomous ther-
1788

mal self-regulation: at high temperature (sunlit, electronics active), emissivity increases to
1789

reject heat; at low temperature (eclipse), emissivity decreases to conserve heat. This self-
1790

regulating behaviour eliminates active heaters in many scenarios, reducing power demand
1791

on power-constrained large inatable platforms. The principal barrier to inatable applica-
1792

tion is substrate compatibility: VO2 coatings currently require rigid silicon substrates, while
1793

EclipseVED has been demonstrated only on rigid aluminium panels. Developing these tech-
1794

nologies on exible polymer substrates (Kapton, polyimide) is a near-term research priority.
1795


![Table 40](paper-23-v1_images/table_40.png)
*Table 40*

Table 19: Variable emissivity coating technologies for spacecraft thermal control.

Technology
∆ε
Tswitch
Power
TRL
Flight Heritage

VO2 (Kim 2019)
0.35 (hemi.)
67 ◦C
Zero
34
None
VO2 (Hendaoui 2013)
0.49 (normal)
67 ◦C
Zero
3
None
EclipseVED (electrochromic)
∼0.71
Voltage ctrl
13 V
78
MidSTAR-1 (2007)
MEMS louvers
∼0.8 (e.)
Bimetal
Zero
78
Multiple

10.4
Loop Heat Pipes for Deployed Structures
1796

Loop heat pipes (LHPs) are the preferred heat transport technology for active thermal
1797

systems in space, oering passive capillary-driven two-phase uid transport with zero pump
1798

power, distances up to several tens of metres, and heat loads up to 5+ kW per evaporator
1799

Maydanik [2005]. The capillary driving force is generated by a sintered porous wick conned
1800

to a compact evaporator body; vapour and liquid travel in separate smooth-wall transport
1801

lines. A compensation chamber at the evaporator provides thermal buering and enables
1802

active setpoint control to ±0.5 ◦C via low-power heaters (15 W). Working uids for space
1803

include ammonia (−40 to +70 ◦C, the standard), propylene (−60 to +50 ◦C, when ammonia
1804

freeze risk exists), and ethane (−100 to +30 ◦C) for cryogenic applications.
1805

LHP spaceight heritage extends over 35 years, beginning with the Granat astrophysics
1806

satellite in 1989 and encompassing over 30 systems own by 2005 across Russian, American,
1807

and European programmes Maydanik [2005]. The Hughes HS-702 communications satel-
1808

lite (1999) demonstrated the rst LHP-coupled deployable radiatorthe directly relevant
1809

precedent for inatable structures, as the LHP exible transport lines accommodated the
1810

mechanical hinge between the deployed radiator panel and the spacecraft bus. NASA's EOS
1811

Terra and Aqua missions, ICESat/GLAS, and Swift/BAT all employed LHP thermal control.
1812

For inatable habitats, LHPs are the natural technology for transporting waste heat
1813

from interior systems (avionics, crew metabolic load) to external deployable radiators. The
1814

exible transport lines can be routed through deployment hinges and accommodate the
1815

geometric changes between stowed and deployed congurations. Current single-evaporator
1816

LHP systems transport 50700 W in typical spacecraft congurations, with multi-loop archi-
1817

tectures providing aggregate capacities exceeding 10 kW for large platforms. The principal
1818

engineering challenge for inatable integration is the condenser interface: bonding or me-
1819

chanically attaching the condenser panel to the exible membrane requires a solution to the
1820

rigid-to-exible interface problem discussed in Section 12.3.
1821

10.5
Phase Change Materials in Fabric Layers: The TRL 23 Gap
1822

Phase change materials (PCMs) oer passive thermal buering by absorbing and releasing
1823

latent heat during orbital day/night transitions. For LEO inatable habitats experiencing
1824

90-minute thermal cycles, the most promising PCM candidates are n-eicosane (melting point
1825

36.4 ◦C, latent heat 247253 J/g) and n-octadecane (28.2 ◦C, 244 J/g) Diaconu et al. [2023].
1826

PCM-based thermal control for rigid electronics enclosures has extensive spaceight her-
1827

itage spanning from Apollo Lunar Roving Vehicle battery management (1971) through Mars
1828

rovers (Spirit, Opportunity, Curiosity, Perseverance; TRL 9) and ISS experiments (TRL 56)
1829

Diaconu et al. [2023].
1830

However, integration of PCMs into exible fabric layers for inatable structuresthe
1831

conguration needed to provide distributed thermal buering across large membrane areas
1832

remains at TRL 23. Five specic technical barriers have been identied:
1833

1. Microgravity containment: Liquid-phase PCM migrates freely in zero-g. Microen-
1834

capsulation (1100 µm capsules) addresses this at small scale, but capsule integrity
1835

during the fold/deploy lifecycle has not been tested for space-grade materials.
1836

2. Fold/deploy cycling: PCM-loaded fabric must survive hundreds to thousands of
1837

fold/deploy cycles without capsule rupturea requirement with no demonstrated so-
1838

lution in the space-qualied materials literature.
1839

3. Outgassing: PCM solvents and vapour can contaminate optical surfaces (solar cells,
1840

sensors). Space-qualied encapsulation that meets ASTM E595 outgassing require-
1841

ments has not been characterised for PCM-textile systems.
1842

4. Thermal conductivity: Raw paran PCMs have thermal conductivity k ≈0.2 W/(m·K)
1843

approximately 1,000× lower than aluminiumresulting in slow thermal response. Car-
1844

bon nanotube or graphene additives can improve conductivity to 15 W/(m·K) but at
1845

the cost of reduced fabric exibility and increased mass.
1846

5. Atomic oxygen interaction: PCM capsule shells (typically PMMA or gelatin) may
1847

erode under atomic oxygen ux in LEO, releasing PCM material and contaminating
1848

the local environment.
1849

Despite these barriers, the potential benet is substantial. A 1 kg/m2 layer of microencap-
1850

sulated n-eicosane would provide ∼250 J/g × 1,000 g/m2 = 250 kJ/m2 of thermal storage
1851

sucient to buer the rst ∼10 minutes of eclipse entry for a membrane with low thermal
1852

mass, signicantly reducing peak-to-trough temperature excursions. The technology needs
1853

a structured development programme analogous to what IRVE provided for exible thermal
1854

protection systems.
1855

11
State of the Art: Attitude and Orbit Control
1856

Attitude and orbit control for large inatable space structures is dominated by a single over-
1857

arching challenge: control-structure interaction (CSI). When structural exibility approaches
1858

or overlaps the attitude control bandwidth, conventional rigid-body AOCS designs become
1859

inadequate or unstable. For 100 m-class inatable structures, where the lowest natural fre-
1860

quencies may fall well below 0.1 Hz, CSI is not merely a complicationit is the central design
1861

driver. This section reviews the CSI challenge, the theoretical framework of gyroelastic body
1862

dynamics, the drag budget for large LEO structures, and the critical gap in AOCS theory
1863

for pressure-stabilised membranes.
1864

11.1
Control-Structure Interaction for Flexible Spacecraft
1865

CSI has been studied since the 1970s in the context of large space systems including the
1866

Solar Power Satellite concept, the Space Station, and large deployable antennas. For me-
1867

chanically sti structuresrigid trusses, mesh antennas, deployable solar arraysthe lowest
1868

structural modes typically fall in the 0.11 Hz range for 1030 m scale structures, and struc-
1869

tural damping ratios ζ ≈0.0010.005 are small but predictable Nicassio et al. [2022]. The
1870

standard approach is modal truncation and notch ltering: identify the structural modes, ex-
1871

clude them from the control bandwidth, and ensure adequate frequency separation between
1872

rigid-body and exible modes.
1873

For inatable (pressure-stabilised) structures, the CSI problem is qualitatively dierent
1874

in four respects. First, structural stiness is primarily provided by membrane tension arising
1875

from internal pressure (σhoop = pR/t for a cylindrical geometry) rather than material bending
1876

stiness, and this stiness changes if pressure is lost due to microleaks or thermal cycling.
1877

Second, the lowest natural frequencies scale inversely with structure size and can be ≪
1878

0.1 Hz for 100 m-class structures, potentially falling within the AOCS bandwidth. Third,
1879

membranes cannot carry compressive stressthey wrinkle, creating local zones of nonlinear
1880

stiness that invalidate linear modal analysis. Fourth, actuator forces transmitted through a
1881

exible membrane diuse spatially rather than transmitting cleanly through a rigid structural
1882

path, degrading actuator-to-mode coupling. No paper in the published literature explicitly
1883

addresses AOCS for pressure-stabilised inatable structures at the 100 m scale.
1884

Angeletti et al. Nicassio et al. [2022] developed a minimum complexity hybrid ODE-
1885

PDE model for large exible spacecraft that provides a useful methodological template: the
1886

rigid bus is treated as an ODE system (6 DOF) coupled to the exible structure as a PDE
1887

system (beam/plate). Even a 2-mode truncation captured over 80% of the relevant dynamics
1888

for control design. However, this framework assumes conventional bending stiness and is
1889

not directly applicable to pressure-stabilised membranes.
1890

11.2
Gyroelastic Body Theory and Distributed Momentum Man-
1891

agement
1892

The theoretical foundation for distributed attitude actuators on exible structures was estab-
1893

lished by D'Eleuterio and Hughes in a series of foundational papers D'Eleuterio and Hughes
1894

[1984, 1986, 1987]. The 1984 paper introduced the concept of gyricitythe distribution of
1895

angular momentum per unit volume embedded within an elastic continuum. The governing
1896

equations couple elastic deformation to rigid-body rotation through the gyricity distribu-
1897

tion g(x), showing that distributed angular momentum fundamentally modies elastic wave
1898

propagation and natural frequencies. The key theoretical nding is that gyroelastic systems
1899

have complex eigenvalues (gyroelastic frequency splitting), providing passive damping-like
1900

behaviour without explicit energy dissipationanalogous to Zeeman splitting in quantum
1901

mechanics D'Eleuterio and Hughes [1984].
The 1986 companion paper D'Eleuterio and
1902

Hughes [1986] derived the modal parameters (complex mode shapes, orthogonality condi-
1903

tions, participation factors) needed for practical numerical analysis, while the 1987 paper
1904

D'Eleuterio and Hughes [1987] extended the framework to complete spacecraft systems,
1905

treating a vehicle with distributed angular momentum storage as a unied gyroelastic body.
1906

Damaren and D'Eleuterio Damaren and D'Eleuterio [1989] solved the optimal gyricity
1907

distribution problem using calculus of variations: the spatial distribution g∗(x) that min-
1908

imises a quadratic performance index concentrates angular momentum where modal kinetic
1909

energy is highestat the antinodes of the dominant vibration modes. This is directly analo-
1910

gous to collocating sensors at modal antinodes and provides the theoretical basis for actuator
1911

placement optimisation on large exible structures.
1912

The most recent quantitative validation of distributed momentum management was pro-
1913

vided by Cachim et al. Cachim et al. [2024], who compared centralized (6 large reaction
1914

wheels on the bus) versus distributed (33 small reaction wheels throughout the structure)
1915

attitude control for a ∼30 m hexagonal plate-like structure (4,200 kg, Jxx = 2.2×105 kg·m2).
1916

Using LQG control with 25 retained modes below 80 Hz, the distributed conguration
1917

achieved 3.3× faster settling (30 s versus 100 s), 7× less structural deformation (0.33 µm
1918

versus 2.3 µm) during a 0.5◦slew, and improved ne pointing (RMS error 0.038 versus
1919

0.068 arcsec), at the cost of approximately 2× more total torque Cachim et al. [2024]. The
1920

structure was modelled as a Kirchho plate (bending-only, shear neglected), which is valid
1921

for thin plates with thickness-to-span ratio >1:30 but is not applicable to pressure-stabilised
1922

membranes.
1923

11.3
Drag Budget for 100 m-Class LEO Structures
1924

A 100 m-class inatable structure in LEO faces a severe drag penalty due to its extreme area-
1925

to-mass ratio. At 500 km altitude, atmospheric density varies from ρ ≈5×10−13 kg/m3 (solar
1926

minimum) to ρ ≈3×10−12 kg/m3 (solar maximum)a factor of 6× variation driven by solar
1927


![Table 41](paper-23-v1_images/table_41.png)
*Table 41*

EUV heating of the upper atmosphere Jiang et al. [2022], Andreussi et al. [2022]. For a 100 m
1928

diameter circular membrane presented broadside to the velocity vector (Ae ≈7,850 m2), the
1929


![Table 42](paper-23-v1_images/table_42.png)
*Table 42*

drag force FD = 1

2ρv2CDA yields the estimates in Table 20.
1930

Table 20: Drag force estimates for a 100 m inatable structure at 500 km altitude. CD ≈2.4
3.2 for at membrane in free molecular ow with atomic oxygen accommodation.

Scenario
ρ (kg/m3)
Ae (m2)
CD
FD (N)

Solar min, edge-on
5 × 10−13
100
2.4
0.007
Solar min, broadside
5 × 10−13
5,000
2.4
0.35
Solar min, broadside (max)
5 × 10−13
7,850
3.2
0.72
Solar max, broadside
3 × 10−12
5,000
3.2
14
Solar max, broadside (max)
3 × 10−12
7,850
3.2
21

The drag coecient range of CD = 2.43.2 for a at membrane in free molecular ow is
1931

based on the standard models of Sentman Sentman [1961] and Moe and Moe Moe and Moe
1932

[2005], where the upper bound corresponds to complete diuse reection with full thermal
1933

accommodation on atomic oxygen surfaces.
1934

The area-to-mass ratio is the fundamental problem: if the 100 m structure totals 5,000 kg,
1935

A/m ≈1.6 m2/kg (broadside), compared to ∼0.02 m2/kg for the ISSapproximately 80×
1936

higher. Using the ballistic coecient β = m/(CDA), the orbital decay time at 500 km during
1937

100 m diameter structure, CD = 2.2


![Table 43](paper-23-v1_images/table_43.png)
*Table 43*

Broadside, solar max

103

Broadside, solar min

500 km ref.

Edge-on, solar min

SRP reference (0.035 N)

102

101

3.64 N

100

Drag force (N)

0.28 N

10
1

0.02 N

10
2

10
3

Conventional

10
4

spacecraft
drag range

10
5

10
6

300
400
500
600
700
800
Altitude (km)

Figure 11: Drag force versus altitude for a 100 m diameter inatable structure in LEO,
showing solar minimum and solar maximum atmospheric conditions. The shaded region
illustrates the factor-of-six variation in atmospheric density driven by the solar cycle, which
dominates the orbit maintenance propellant budget.

solar maximum can be estimated at approximately 36 months for broadside orientation,
1938

conrming that orbital lifetime without propulsion would be months, not years.
1939

Second-Order Eects
1940

Three additional forces merit consideration for a complete 100 m-class force budget:
1941

Solar radiation pressure (SRP): For a 100 m diameter membrane at 500 km, the SRP
1942

force is FSRP = (P⊙/c) · A · (1 + r) ≈(4.56 × 10−6 N/m2) × 7,850 m2 × 1.5 ≈0.054 N, where
1943

P⊙= 1,361 W/m2 is the solar ux, c is the speed of light, and r ≈0.5 is the reectivity.
1944

This SRP force is comparable to the drag at solar minimum edge-on (0.007 N) and non-
1945

negligibleat solar minimum with edge-on orientation, SRP may actually dominate over
1946

atmospheric drag.
1947

Attitude-dependent cross-section: The table presents edge-on (100 m2) and broad-
1948

side (7,850 m2) as discrete cases, but a real membrane oscillates between attitudes unless
1949

actively controlled. The time-averaged eective area depends on AOCS capabilitycoupling
1950

the drag analysis to the AOCS gap (C4). Passive spin stabilisation about the minimum-
1951

inertia axis would yield a time-averaged Aeff intermediate between edge-on and broadside,
1952

approximately 0.5×Abroadside ≈3,900 m2, roughly halving the broadside drag but still orders
1953

of magnitude above edge-on.
1954

Propellant mass rate derivation: The xenon propellant consumption for Hall thruster
1955

drag compensation can be derived as ˙m = FD/(g0Isp), where g0 = 9.81 m/s2 and Isp = 3,000 s
1956

for a representative Hall thruster. For the solar-minimum broadside case (FD = 0.35 N):
1957

˙m = 0.35/(9.81 × 3,000) = 1.19 × 10−5 kg/s = 1.03 kg/day = 376 kg/year. For the solar-
1958

maximum broadside case (FD = 21 N): ˙m = 21/(9.81 × 3,000) = 7.14 × 10−4 kg/s =
1959

61.7 kg/dayclearly unsustainable without in-orbit refuelling. The corresponding thrust
1960

power is P = FDve/(2η), where ve = g0Isp = 29,430 m/s and η = 0.6 (thruster eciency):
1961

yielding 8.6 kW for the solar-minimum broadside case and 515 kW for the solar-maximum
1962

broadside case. The 150 kW range stated in Section 13.2 corresponds to solar-minimum
1963

conditions with partial edge-on attitude control.
1964

Air-Breathing Electric Propulsion (ABEP), which collects atmospheric gas for use as
1965

propellant, has been proposed for drag compensation in Very Low Earth Orbit (VLEO,
1966

150450 km) Andreussi et al. [2022]. However, at 500 km the atmospheric density is approx-
1967

imately 100× lower than at the 250350 km altitudes where ABEP is designed to operate,
1968

reducing achievable thrust to 0.0010.1 mNorders of magnitude insucient for the 0.35
1969

21 N drag forces computed above. Conventional electric propulsion (Hall eect or gridded
1970

ion thrusters) with onboard xenon propellant is the only viable station-keeping option. This
1971

propulsion requirement fundamentally constrains mission architecture and represents a sig-
1972

nicant fraction of the overall mass budget.
1973

11.4
The Missing Theory:
AOCS for Pressure-Stabilised Mem-
1974

branes
1975

The gyroelastic body framework of D'Eleuterio and Hughes assumes elastic continua with
1976

Cauchy stress tensor constitutive relationsvalid for beams, plates, and shells with inher-
1977

ent bending stiness. Extending this framework to pressure-stabilised inatable membranes
1978

requires four theoretical modications that represent a signicant gap in the published lit-
1979

erature:
1980

1. Pressure-stiness coupling: For an inatable structure, the eective stiness Ke =
1981

Kmembrane + Kpressure, where the pressure contribution depends on ination state and
1982

couples to deformation through the ideal gas law. When pressure changes due to mi-
1983

croleaks or thermal cycling, natural frequencies shift and gyroelastic modes recongure
1984

a time-varying system for which xed-gain controllers may become unstable.
1985

2. Wrinkling constraint: Membranes cannot carry compressive stress; they wrinkle,
1986

creating zones where σn = max(0, Tmembrane · εn). This state-dependent nonlinearity
1987

causes mode shapes to change with the deformation state, invalidating the linear modal
1988

analysis assumption that underpins both the D'Eleuterio framework and the Cachim
1989

optimisation.
1990

3. Orthotropic fabric constitutive model: Space fabrics (Vectran, Kevlar) are woven
1991

materials with highly anisotropic stinesswarp versus weft direction stiness can
1992

dier by 25×. The isotropic elastic continuum in the D'Eleuterio formulation requires
1993

replacement with an orthotropic constitutive model.
1994

4. Gas-structure interaction coupling: For large inatable volumes, internal gas has
1995

its own dynamics (acoustic modes, pressure wave propagation). This is analogous to
1996

liquid sloshing in fuel tanksa well-studied problembut the gas-structure coupling
1997

for inatable membranes has received no published treatment.
1998

Each of these extensions builds upon established prior work, and the timeline can be
1999

estimated with some granularity:
2000

 Pressure-stiness coupling (estimated 34 years): The coupling of ination pressure
2001

to membrane stiness is well-understood for simple geometries through the gossamer
2002

structure dynamics literature Jenkins [2001]. The novel challenge is coupling this to the
2003

gyroelastic formulation, requiring a pressure-dependent constitutive model within the
2004

D'Eleuterio framework. This is the most tractable extension and could be addressed
2005

within a focused doctoral programme.
2006

 Wrinkling constraint (estimated 34 years): Tension-eld theory Stein and Hedgepeth
2007

[1961] provides a well-established framework for membranes that cannot sustain com-
2008

pression. Roddeman et al. Roddeman et al. [1987] developed the modern computa-
2009

tional treatment. Integrating wrinkling-induced state-dependent stiness into gyroe-
2010

lastic eigenvalue analysis is non-trivial but has analogues in rotor dynamics (cracked
2011

shaft models with breathing cracks).
2012

 Orthotropic fabric constitutive model (estimated 12 years): Replacing isotropic
2013

with orthotropic constitutive relations requires substituting the appropriate fourth-
2014

order stiness tensor into the D'Eleuterio equations. The D'Eleuterio formulation uses
2015

the general Cauchy stress tensor, making the extension algebraically systematic. This
2016

is the most tractable extension and could constitute the early phase of a doctoral
2017

programme or a Master's thesis.
2018

 Gas-structure interaction coupling (estimated 45 years): This is the most novel
2019

and uncertain extension.
The fuel-sloshing analogy Abramson [1966] is useful but
2020

incompletegas is compressible while classical sloshing models assume incompressibil-
2021

ity. Coupled gas-membrane problems have been studied in the aeroelasticity literature
2022

(utter of inated membrane wings Leclercq and de Langre [2018]), providing a starting
2023

point, but the three-dimensional coupling for large inatable volumes in the gyroelastic
2024

context has no precedent. This is the genuine multi-year research challenge.
2025

The total estimated timeline is 1215 years if pursued sequentially by individual doctoral
2026

candidates, or 57 years if pursued in parallel by a coordinated research group with 23
2027

concurrent doctoral projects. The sequential estimate of 1015 years stated in Section 13
2028

is therefore conservative but reasonable. This is among the most signicant fundamental
2029

research gaps identied in this survey.
2030

12
State of the Art: Robotic In-Orbit Assembly
2031

The vision of large inatable space structures100 m-class debris shields, large-aperture
2032

antenna reectors, or orbital habitats exceeding ISS volumewill likely require in-orbit
2033

assembly of subsystems that exceed the launch vehicle fairing envelope or are too complex
2034

for single-deployment architectures.
This section reviews the state of in-space servicing,
2035

assembly, and manufacturing (ISAM) robotics, the E-Walker concept for walking robots on
2036

large structures, and the critical gap in rigid-to-exible interface technology that currently
2037

prevents assembly on inatable substrates.
2038

12.1
Assembly Robot Heritage and Current Programmes
2039

In-orbit robotic assembly heritage begins with the ISS, whose construction (19982011) relied
2040

on the Canadarm2 Space Station Remote Manipulator System (SSRMS): a 17.6 m, 7-DOF
2041

arm operating from xed Power Data Grapple Fixtures (PDGFs) on the truss structure.
2042

Canadarm2 demonstrated that large-scale orbital assembly is achievable with telerobotic
2043

systems, but at the cost of extensive EVA support and ground-in-the-loop operations.
2044

The ISAM landscape has expanded substantially since ISS assembly. NASA's 2025 State
2045

of Play report catalogues 524 capability entries across 145 developers in 21 countries, with
2046

over $2 billion in government investment NASA [2025].
Current programmes span mul-
2047

tiple technology readiness levels: GITAI's S2 experiment demonstrated autonomous ISS
2048

solar array assembly (2021); Project GHOST validated tool manipulation in orbit (2024);
2049

DARPA's NOM4D programme targets LEO truss assembly demonstration by Caltech in
2050

2026; and NASA Langley's CIRAS/TALISMAN/SAMURAI/NINJAR ground demonstra-
2051

tions have validated multi-robot truss assembly at 15 m scale Li et al. [2022c], Doggett et al.
2052

[2018]. The European PULSAR project targets autonomous assembly of a 12 m telescope
2053

mirror Rognant et al. [2019]. Northrop Grumman's MEV-1 (2020) and MEV-2 (2021) rep-
2054

resent the rst commercial ISAM operations, though these are servicing (docking with client
2055

spacecraft) rather than structural assembly.
2056

A critical observation for the present survey is that all 524 entries in the NASA ISAM
2057

catalogue address assembly of rigid structurestrusses, beams, modular satellites, and mir-
2058

ror segments NASA [2025]. Not a single entry addresses assembly on or of inatable/exible
2059

substrates. This is not a mere omission; it reects a fundamental gap in the technology base:
2060

the rigid-to-exible interface problem remains unsolved (Section 12.3).
2061

12.2
Walking Robots for Large Structure Assembly: E-Walker
2062

The End-over-End Walking Robot (E-Walker) represents the current state of the art in
2063

walking manipulators designed for ISAM missions Nair et al. [2022, 2024]. Inheriting the
2064

Canadarm2 design philosophy of end-over-end locomotion via grapple xtures, the E-Walker
2065

is a 7-DOF dexterous manipulator at full scale of approximately 475 kg with 350 kg pay-
2066

load capacitysucient to handle one primary mirror segment for a 25 m Large Aperture
2067

Space Telescope (LAST). Maximum joint torque reaches ∼70 Nm at Joint 2, and nite ele-
2068

ment analysis conrms maximum link deection of only 0.04 mm under full payload, with a
2069

buckling safety factor exceeding 129 Nair et al. [2022].
2070

A scaled prototype (1.3 m, 12 kg, 2 kg payload at 1:6 scale) has been demonstrated in
2071

ground testing. Nair et al. Nair et al. [2024] evaluated 11 concepts of operations for 25 m
2072

telescope assembly, concluding that a dual E-Walker conguration is optimal. The 8 m E-
2073

Walker requires 4.5 m less workspace than an equivalent xed-base arm, making walking
2074

locomotion particularly advantageous for assembly tasks distributed over large structures.
2075

However, all E-Walker analysis assumes a rigid assembly substrate. The grapple x-
2076

tures are ISS-standard PDGFs requiring rigid interfaces with ±10 mm capture tolerance and
2077

multi-kN load capacity. When the E-Walker applies 70 Nm joint torques during assembly
2078

operations, Newton's third law transmits equal and opposite reactions into the mounting
2079

substrate. On the ISS rigid truss, these are absorbed globally; on an inatable membrane,
2080

they would cause local deformation, potential wrinkling, and excitation of global vibration
2081

modes. The 475 kg robot's every movement in microgravity creates reaction forces that, on
2082

a exible membrane, propagate as structural disturbances.
2083

12.3
The Rigid-to-Flexible Interface Gap
2084

All existing docking and assembly interfaces assume rigid-to-rigid connections. Chen et al.
2085

Liu et al. [2024] designed an androgynous docking port with ±23.5 mm translation tolerance
2086

for on-orbit assemblya practical engineering specication for robotically-assisted mating
2087

of rigid modules. ISS Power Data Grapple Fixtures, common berthing mechanisms, and all
2088

ISAM interface concepts in the literature share this rigid-to-rigid assumption.
2089

No published work specically addresses distributed rigid-module attachment to inat-
2090

able membranes in the space environment. However, several bodies of adjacent work provide
2091

relevant design heritage that should be acknowledged:
2092

 Tensegrity structures: Tensegrity platforms Skelton and de Oliveira [2009] inher-
2093

ently address the rigid-to-exible interface through bar-cable connections.
NASA
2094

Ames' Super Ball Bot Sabelhaus et al. [2015] demonstrates rigid node attachment to
2095

tensioned cables in a recongurable structure; the load-spreading problem at hardpoint-
2096

membrane interfaces is structurally analogous to the bar-cable joint in tensegrity.
2097

 Deployable antenna feed support: Large deployable mesh antennas (Harris/L3
2098

AstroMesh, Northrop Grumman CRAF) attach a rigid feed assembly to a tensioned
2099

cable-net/mesh reector surface Santiago-Prowald and Rodrigues [2018].
The feed
2100

support struts connect rigid hardware to a exible, tension-stabilised structurea
2101

direct analogue to the rigid-module-on-inatable-membrane problem.
2102

 Solar sail boom-membrane attachment: Solar sail designs (e.g., IKAROS, NEA
2103

Scout) attach rigid booms to thin-lm membranes via reinforced corner ttings. The
2104

stress concentration and load distribution at these attachment points have been anal-
2105

ysed in the solar sail literature Fernandez et al. [2014].
2106

The gap remains genuine: none of these analogues addresses the full combination of vac-
2107

uum, thermal cycling, atomic oxygen, micrometeoroid exposure, and zero-gravity dynamics
2108


![Table 44](paper-23-v1_images/table_44.png)
*Table 44*

on an inatable pressure-stabilised substrate. The adjacent work provides starting points
2109

for analysis but not validated solutions.
2110

Table 21 summarises the technology readiness of assembly interface approaches.
2111

The closest ight analog is the BEAM-ISS interface: a rigid berthing ring connects the
2112

inatable module to the ISS Node 3 (Tranquility) common berthing mechanism. This is a
2113

single rigid-to-inatable joint at the berthing interface, not a distributed attachment system
2114

across the membrane surface. No demonstrated technology exists for attaching multiple rigid
2115


![Table 45](paper-23-v1_images/table_45.png)
*Table 45*

Table 21: Assembly interface technology readiness for space structures.

Interface Type
TRL
Heritage
Notes

Rigid-to-rigid (PDGF)
9
ISS
Operational since 2001
Rigid-to-rigid (androgynous)
34
Ground demo
Chen et al. 2024
Rigid-to-exible (hardpoint)
23
BEAM ring
Conceptual only
Rigid-to-exible (distributed)
12
None
No published work

subsystems (reaction wheels, solar array drives, communications antennas) to an inatable
2116

membrane at distributed locations. This is a novel nding of this survey and represents a
2117

critical research gap.
2118

12.4
Assembly-Enabled Inatable Platforms: Design Requirements
2119

Based on the analysis in Sections 12.212.3, a set of design requirements for assembly-enabled
2120

inatable platforms can be identied:
2121

1. Embedded rigid attachment rings: Metallic rings (0.51 m diameter) must be sewn
2122

into the inatable fabric at pre-determined assembly points during manufacturing, with
2123

integrated load-spreading plates to distribute reaction forces over sucient membrane
2124

area. The stress concentration factor at such embedded hardpoints (25× local stress
2125

amplication) must be accounted for in the membrane structural design.
2126

2. Compliance layer: A 35 mm silicone or elastomeric foam layer between each rigid
2127

attachment ring and the membrane accommodates local deformation and provides
2128

vibration isolation, preventing point-load damage to the fabric.
2129

3. Pre-integration requirement: Retrotting hardpoints onto an already-deployed
2130

inatable is impractical. All assembly interfaces must be designed in and manufactured
2131

as part of the inatable structure before launch. This implies that the assembly concept
2132

of operations must be fully dened before the inatable is manufactureda signicant
2133

systems engineering constraint.
2134

4. Active vibration isolation: Small dampers or isolation mounts between each E-
2135

Walker grapple point and the membrane surface attenuate reaction forces from assem-
2136

bly operations, reducing excitation of global membrane vibration modes.
2137

5. Pressure-aware operations: Assembly operations that change the mass distribu-
2138

tion (adding subsystems) alter both the inertia tensor and the natural frequencies of
2139

the inatable structure.
AOCS must accommodate these time-varying dynamics
2140

connecting to the gap identied in Section 11.4.
2141

The E-Walker on an inatable platform is conditionally feasible with pre-integrated hard-
2142

points, compliance layers, and active vibration isolation. However, none of these solutions has
2143

been demonstrated even at component level for space applications. A ground demonstration
2144

programmeanalogous to NASA Langley's CIRAS/TALISMAN truss assembly demonstra-
2145

tions but on an inatable test articlewould represent a signicant advance toward closing
2146

this gap.
2147

13
Challenges, Open Questions, and Research Roadmap
2148

The preceding eight technology surveys (Sections 512) have documented a paradox that
2149

denes the current state of soft inatable robotic systems for space: individual enabling tech-
2150

nologies have reached moderate-to-high readiness levelsVectran restraint layers at TRL 9
2151

(Section 5), shape memory alloy deployment actuators at TRL 89 (Section 7.5), bre Bragg
2152

grating sensors on rigid spacecraft at TRL 78 (Section 8.1)yet no integrated soft inat-
2153

able robotic system has been demonstrated in space. This section consolidates the research
2154

gaps identied throughout the survey, assesses their severity and interdependence, proposes
2155

a structured research roadmap spanning 5-year and 15-year horizons, and identies the most
2156

viable path to a near-term ight demonstration.
2157

13.1
Critical Research Gaps
2158

A systematic analysis of the technology areas reviewed in Sections 512 reveals 5 critical
2159

gaps, 9 moderate gaps, and 10 minor gaps. Here we consolidate the 5 critical gaps, each of
2160

which represents a showstopper for at least one major application domain.
2161

C1: Absence of Quantitative Soft-versus-Rigid Fragmentation Comparison.
The
2162

central motivation for soft capture in active debris removal (Section 3.2) rests on the propo-
2163

sition that compliant mechanisms reduce fragmentation risk relative to rigid robotic arms.
2164

Qualitative evidence supports this argument: Wang et al. Wang et al. [2023] identied the
2165

potential to generate fragments during the capturing phase for rigid systems; Chen et
2166

al. Chen et al. [2024a] concluded that single contact-based caging is excessively risky for
2167

fast-tumbling targets; and the RemoveDebris harpoon test demonstrated structural fail-
2168

ure of a carbon bre boom at 20 m s−1 impact Aglietti et al. [2020]. The e.deorbit mission
2169

study computed peak joint torques of 195 N m for capture of an 8-tonne ENVISAT tumbling
2170

at 5 ◦s−1 Flores-Abad et al. [2014]. However, no published study provides a quantitative
2171

fragmentation probability as a function of contact compliance. The catastrophic fragmenta-
2172

tion threshold (10 J g−1 specic energy from the IMPACT model Johnson et al. [2001]) has
2173

never been applied to a soft-versus-rigid capture force comparison. The fragmentation risk
2174

is physically plausible and supported by qualitative assessmentsparticularly for degraded
2175

appendages (solar panels, thermal blankets, antennas) that may have lost 3060% of their
2176

original strength through decades of space environment exposurebut remains experimen-
2177

tally unquantied. This survey adopts the precautionary principle that compliant capture
2178

is preferred until quantitative data become available, on the basis that the consequences of
2179

inadvertent fragmentation are severe enough to warrant risk-averse technology selection. We
2180

propose this as the single highest-priority experimental investigation the community should
2181

undertake, requiring hypervelocity and low-velocity impact testing with debris surrogates at
2182

varying contact compliance levels.
2183

C2: No Soft Robotic Capture System Has Flown in Space.
Despite eight distinct
2184

soft or compliant capture approaches documented in Section 3.2gecko adhesive (TRL 4
2185

5), DEMES grippers (TRL 34), bistable soft grippers (TRL 23), cryogenic metallic cable
2186

robots (TRL 3), inatable origami arms (TRL 3), ytrap origami (TRL 23), thermally
2187

qualied multi-layer grippers (TRL 2), and the INSIDeR system concept (TRL ∼4)none
2188

has own. The gecko adhesive gripper of Jiang et al. Jiang et al. [2017] achieved microgravity
2189

validation with 100% capture success rate on spherical targets and capacity exceeding 400 kg,
2190

making it the most mature candidate. However, this gripper operates on a rigid robotic arm
2191

platform and is more accurately classied as a compliant end-eector on a conventional
2192

manipulator (Section 3.1). The gap between ground/parabolic-ight demonstration and or-
2193

bital ight requires addressing space environment qualication (vacuum outgassing, thermal
2194

cycling, radiation exposure over mission-duration timescales) for which limited data exist.
2195

C3:
Rigid-to-Flexible Assembly Interface Lacks Specic Published Research.
2196

Section 12.3 identied that no published work specically addresses distributed rigid-module
2197

attachment to inatable membranes in the space environment, though adjacent work in
2198

tensegrity structures Skelton and de Oliveira [2009], Sabelhaus et al. [2015], deployable an-
2199

tenna feed supports Santiago-Prowald and Rodrigues [2018], and solar sail boom-membrane
2200

attachments Fernandez et al. [2014] provides relevant design heritage. All heritage docking
2201

interfacesISS PDGF, Common Berthing Mechanism, ClearSpace-1 capture arms, and the
2202

androgynous interfaces reviewed by Chen et al. Chen et al. [2024b]assume rigid-to-rigid
2203

mating. At the 100-metre scale required for large inatable debris shields (Section 11.3) or
2204

solar power platforms, the inatable structure becomes a platform onto which functional
2205

modules must be assembled in orbit Nair et al. [2024], Li et al. [2019]. The reaction force
2206

problemhow to apply assembly torques to a membrane that deforms under the applied
2207

loadhas no published solution specic to the space inatable context. Embedded metallic
2208

hardpoint rings represent a plausible design concept informed by the tensegrity and antenna
2209

feed analogues, but require detailed nite element analysis of stress concentration at the
2210

rigid-exible interface, none of which has been published.
2211

C4: No Published AOCS Theory for Pressure-Stabilized Inatable Structures.
2212

The control-structure interaction literature reviewed in Section 11.1 addresses rigid trusses,
2213

mesh antennas, and mechanically stiened deployable arraysstructures with inherent sti-
2214

ness independent of pressurization. Pressure-stabilized inatable structures exhibit funda-
2215

mentally dierent dynamics: stiness is a function of ination pressure (a time-varying pa-
2216

rameter), membranes wrinkle under compression introducing piecewise-linear stiness non-
2217

linearity, fabric is anisotropic, and internal gas couples to structural modes D'Eleuterio
2218

and Hughes [1984], Jenkins [2001]. The D'EleuterioHughes gyroelastic body framework
2219

D'Eleuterio and Hughes [1984, 1986, 1987] provides the most promising theoretical founda-
2220

tion, but requires four extensions: (i) pressure-dependent constitutive model for membrane
2221

elements, (ii) wrinkling constraints reecting piecewise-linear stiness transitions, (iii) or-
2222

thotropic fabric constitutive laws, and (iv) gas-structure coupling for internal atmosphere
2223

dynamics. Each extension constitutes a substantial theoretical undertaking; collectively they
2224

dene a research programme of 1015 years.
2225

C5: Inatable-Power Integration Gap.
The PowerSphere programme (Section 9.2)
2226

demonstrated thin-lm photovoltaic integration with an inatable substrate using amor-
2227

phous silicon cells, achieving 7.25 W kg−1 at 10% cell eciency Cadogan et al. [2003]. The
2228

programme has been inactive since approximately 2009, and no successor has been identied.
2229

Meanwhile, perovskite/CIGS tandem cells have achieved 2100 W kg−1 with 25 µm substrates
2230

and greater than 85% power retention after more than 50 years of LEO-equivalent proton irra-
2231

diation Lang et al. [2020]. The technology exists to revive inatable-integrated photovoltaics
2232

at 20300× the specic power of the original PowerSphere, yet no programme is pursuing
2233

this integration. The gap is institutional rather than technical: exible PV researchers and
2234

inatable structure researchers operate in separate communities with no overlap programme.
2235


![Table 46](paper-23-v1_images/table_46.png)
*Table 46*

Vectran restraint
Kevlar MMOD
Nextel bumper
Zylon (interior)
SMP rigidisation

Materials &

Structures

BEAM inflation
InflateSail
LOFTID
Origami packaging
Active controlled

Deployment

Mechanics

SMA hinges
Tendon-driven
DEA/DEMES
Vacuum-gap electrostatic
Jamming (vacuum)

Actuation

FBG (heritage)
FBG in webbing
Multicore FOSS
Distributed impact
Capacitive soft

Sensing &

SHM

ROSA/iROSA
CIGS thin-film

Power
Systems

Li-ion batteries
Perovskite PV
PowerSphere-type

MLI (heritage)
JWST sunshield
VO2 coatings
LHP deployed
PCM in fabric

Thermal
Management

CMG (heritage)
Distributed CMG

AOCS

EP drag comp.
CSI for inflatables

Gyroelastic body

Canadarm2
Walking robots

In-Orbit
Assembly

Docking i/f
Rigid-flex i/f

Autonomous assembly
Concept
(TRL 1-3)

Flight proven

Validated
(TRL 4-6)

(TRL 7-9)

1
2
3
4
5
6
7
8
9
Technology Readiness Level (TRL)

Figure 12: Technology readiness landscape across the eight enabling technology areas re-
viewed in Sections 512. Each marker represents a specic sub-technology; colour indicates
TRL band (red: concept TRL 13; orange: validated TRL 46; green: ight-proven TRL 7
9).
While heritage components (Vectran, FBG, ROSA, MLI, Canadarm2) have reached
TRL 79, the integrative technologies required for soft inatable robotic systemsvacuum-
gap actuators, jamming in vacuum, rigid-to-exible interfaces, distributed momentum man-
agement, and PCM in fabricremain at TRL 23.

13.2
Integration Challenges at System Level
2236

Beyond individual technology gaps, the fundamental barrier to ight-ready soft inatable
2237

robotic systems is system integration. The preceding sections documented integration decits
2238

across multiple interfaces:
2239

 ActuationStructure: Vacuum-gap electrostatic actuators (Section 7.2) achieve >4 N
2240

force at 0.7 g mass Sirbu et al. [2025] using thin-lm polymer multilayer construction
2241

that is structurally analogous to inatable membrane wall architecturesyet no study
2242

has attempted to laminate actuator layers into an inatable arm liner. Similarly, the
2243

jamming-in-vacuum concept (Section 7.6) has a sound physical basis Fitzgerald et al.
2244

[2020] but zero experimental validation in relevant conditions.
2245

 SensingStructure: FBG sensors woven into Vectran webbing have been demon-
2246

strated at NASA JSC on 0.61 m and 2.74 m test articles (TRL 45) Bally Ribbon
2247

Mills and Luna Innovations [2020], while multicore bre optic shape sensing achieves
2248

0.64 mm position accuracy in soft actuators Galloway et al. [2019]. The same FBG
2249

technology could provide both structural health monitoring for inatable walls and
2250

proprioceptive sensing for inatable robotic armsa unied sensing architecture that
2251

has not been proposed or demonstrated.
2252

 PowerThermalStructure: A large inatable membrane with thin-lm PV on the
2253

sun-facing surface, MLI on the space-facing surface, and variable-emissivity coatings
2254

for thermal regulation represents a multi-functional surface that would merge the power
2255

and thermal subsystems into a single membrane layer. The PowerSphere concept ap-
2256

proached this integration using 2004-era materials Cadogan et al. [2003]; 2025-era per-
2257

ovskite/CIGS cells on Kapton or Mylar substrates would share the same polymer base
2258

as inatable MLI layers Lang et al. [2020], making the integration pathway plausible.
2259

 AOCSDeployment: BEAM's deployment anomaly (25 ination bursts over 7 hours;
2260

Section 6.3) illustrates that deployment is a dynamic event with angular momen-
2261

tum consequences. For a free-ying 100-metre inatable, each ination pulse imparts
2262

momentum to the structure, and as the structure changes shape during deployment
2263

its modal frequencies shiftpotentially crossing into the AOCS controller bandwidth
2264

D'Eleuterio and Hughes [1984]. No published work addresses the coupled deployment
2265

AOCS problem for inatables.
2266

 DragPowerThermal Cascade: At 500 km altitude, a 100-metre broadside inat-
2267

able experiences drag forces of 0.3521 N depending on solar activity (Section 11.3). To
2268

illustrate the cascade quantitatively, consider a worked example for the solar-minimum
2269

broadside case (FD = 0.35 N) and the solar-maximum broadside case (FD = 21 N):
2270

Step 1  Thrust:
Hall thruster at Isp = 3,000 s, exhaust velocity ve = g0Isp =
2271

29,430 m/s.
2272

Step 2  Power: Pthrust = FDve/(2η) where η = 0.6. Solar-min broadside: P =
2273

0.35 × 29,430/1.2 = 8.6 kW. Solar-max broadside: P = 21 × 29,430/1.2 = 515 kW.
2274

Step 3  Solar array: At 300 W m−2 (BOL, triple-junction) and 100 W kg−1 system-
2275

level specic power: solar-min requires 29 m2 / 86 kg; solar-max requires 1,717 m2 /
2276

5,150 kgexceeding the entire platform mass budget.
2277

Step 4  Waste heat: At 40% combined losses (thruster + PPU): solar-min generates
2278

3.4 kW waste; solar-max generates 206 kW waste.
2279

Step 5  Radiator: At 200 W m−2 radiator capacity: solar-min requires 17 m2; solar-
2280

max requires 1,030 m2.
2281

This cascade demonstrates that the solar-maximum broadside scenario is infeasible
2282

without active attitude control to reduce Aeff, conrming that drag budget and AOCS
2283

capability are inextricably coupled. Edge-on operation at solar minimum (0.007 N
2284

drag, ∼0.17 kW power, <1 m2 array) is feasible; all other scenarios require either ac-
2285

tive attitude management, altitude selection, or both. The 150 kW range previously
2286

stated applies to solar-minimum conditions with partial attitude control.
No pub-
2287

lished analysis traces this full cascade end-to-end for inatable platforms, and a com-
2288

plete parametric study spanning altitude, solar cycle, attitude strategy, and propulsion
2289

technology is identied as a future research need.
2290

A unifying observation emerges: the integration barriers are not gaps within individual
2291

technology disciplines but gaps between disciplines. The soft robotics community, the inat-
2292

able structures community, the space power community, and the GNC community each have
2293

mature capabilities; the intersections remain unexplored. This fragmentation of the research
2294

landscape is itself a structural challenge that programmatic measures (cross-disciplinary
2295

funding calls, joint ground demonstrators) must address.
2296

13.3
Proposed Research Roadmap: 5-Year and 15-Year Horizons
2297

Based on the gap analysis above and the technology readiness levels documented in Sec-
2298

tions 512, we propose a two-horizon research roadmap. The 5-year horizon (20262031)
2299

targets ground validation and component-level ight demonstration; the 15-year horizon
2300

(20262041) targets system-level ight demonstration and initial operational capability.
2301

5-Year Horizon (20262031).
Five priority activities are identied, each addressing one
2302

or more critical or moderate gaps:
2303

1. Jamming-in-vacuum experimental validation (addresses M1). Ground experi-
2304

ment: vacuum chamber with sealed granular/layer jamming specimen connected to a
2305

pressurized chamber simulating an inatable interior. Measure stiness ratio versus
2306

pressure dierential and compare to terrestrial baselines. Space-compatible granular
2307

media candidates include hollow glass microspheres and metallic powder. This ex-
2308

periment is well-dened, moderate-cost, and publishable regardless of outcome.
If
2309

successful, it validates variable-stiness robotic elements that are simpler in orbit than
2310

on Eartha paradigm inversion for soft space robotics.
2311

2. FBG-in-Vectran-webbing ight demonstration (addresses M6). Current ground
2312

demonstrations at NASA JSC Bally Ribbon Mills and Luna Innovations [2020] have
2313


![Table 47](paper-23-v1_images/table_47.png)
*Table 47*

Now

5-year milestones

2026 2028

Jamming-in-vacuum validation

2027 2030

FBG flight demonstration

2026 2029

Perovskite fold-deploy testing

2027 2030

Rigid-flexible interface prototype

2026 2031

Gyroelastic theory extension

15-year milestones

2029 2035

Soft gripper flight demo

2031 2037

10 m inflatable with PV

2033 2039

Assembly robot on inflatable

5-year milestone

15-year milestone

2035 2041

AOCS-qualified inflatable

Dependency

2026
2028
2030
2032
2034
2036
2038
2040
Year

Figure 13: Research roadmap for soft inatable robotic space systems spanning 5-year and
15-year horizons.
Near-term milestones focus on ground validation of critical unknowns
(jamming-in-vacuum, FBG ight, perovskite fold-deploy, rigid-exible interface); long-term
milestones target integrated ight demonstrations (soft gripper capture, 10 m inatable with
PV, assembly robot on inatable substrate, AOCS-qualied inatable).

reached TRL 45. The next step is a ight experiment on an ISS external payload
2314

platform (e.g., MISSE or Bartlett) exposing FBG-instrumented Vectran webbing to the
2315

LEO environment (atomic oxygen, UV, thermal cycling, MMOD) for 1224 months.
2316

Success would advance the technology to TRL 67 and establish the ight heritage
2317

base for inatable SHM.
2318

3. Perovskite/CIGS fold-deploy-power testing (addresses C5, M5). Deposit per-
2319

ovskite/CIGS tandem cells on 25 µm polymer substrates identical to those used for
2320

inatable MLI. Subject samples to 1000 fold/deploy mechanical cycles, 1000 thermal
2321

vacuum cycles (−100 ◦C to 120 ◦C), and atomic oxygen exposure at LEO-equivalent
2322

uences. Measure power output degradation after each environmental stress. This
2323

establishes whether the remarkable radiation hardness of perovskite/CIGS Lang et al.
2324

[2020] survives the additional mechanical and environmental stresses of inatable in-
2325

tegration.
2326

4. Rigid-to-exible interface ground prototype (addresses C3). Design, fabricate,
2327

and test embedded metallic load-spreader rings sewn into representative multi-layer
2328

inatable fabric during manufacture. Characterize load distribution, stress concentra-
2329

tion factors, and modal response under simulated assembly loading. Compare FEA
2330

predictions with experimental measurements. This ground programme would produce
2331

the rst published dataset on rigid-to-exible assembly interfaces for space inatables.
2332

5. Gyroelastic theory extension for pressure-stabilized membranes (addresses
2333

C4). Mathematical extension of the D'EleuterioHughes framework D'Eleuterio and
2334

Hughes [1984, 1986] incorporating pressure-dependent stiness and fabric orthotropy.
2335

Numerical validation against commercial FEM codes for representative inatable ge-
2336

ometries (cylinder, torus, sphere). Publication of the extended theory would establish
2337

the foundational AOCS framework that any 100-metre-class inatable mission will
2338

require.
2339

15-Year Horizon (20262041).
Four system-level demonstrations dene the long-term
2340

roadmap:
2341

1. Soft gripper ight for debris capture (addresses C1, C2). A CubeSat or small-
2342

satellite class mission demonstrating compliant capture of a cooperative (then non-
2343

cooperative) target in LEO. The gripper subsystem (gecko adhesive, DEMES, or suc-
2344

cessor technology) operates on an inatable arm with integrated FBG sensing. This
2345

mission provides the rst orbital data on soft capture dynamics and validates the frag-
2346

mentation risk reduction argument with ight telemetry.
2347

2. 10-metre inatable with integrated photovoltaics (addresses C5). A free-ying
2348

technology demonstrator deploying a 10-metre-class inatable membrane with lami-
2349

nated perovskite/CIGS cells, demonstrating fold/deploy survival and power generation
2350

in the orbital environment. This bridges the gap between ROSA-class rigid-boom ex-
2351

ible arrays (TRL 9) and the 100-metre inatable solar platforms envisioned for future
2352

missions.
2353

3. Assembly robot on inatable substrate (addresses C3). A ground or parabolic-
2354

ight demonstration of a walking or crawling robot (E-Walker class Nair et al. [2024])
2355

operating on an inatable test article, attaching and detaching rigid modules via em-
2356

bedded hardpoint interfaces. This validates the rigid-to-exible assembly concept in
2357

representative (reduced) gravity conditions.
2358

4. AOCS-qualied pressure-stabilized inatable (addresses C4). A free-ying in-
2359

atable structure (310 metre scale) with onboard AOCS demonstrating three-axis at-
2360

titude control of a pressure-stabilized membrane in LEO. This validates the extended
2361

gyroelastic theory and provides the rst ight data on control-structure interaction for
2362

inatable spacecraft.
2363

Drag Power Thermal Cascade for 100 m Inflatable at 500 km

P = F·ve/2

(  = 60%)
PSA = 300 W/m2

EP 
 100%

waste 
 60%


![Table 48](paper-23-v1_images/table_48.png)
*Table 48*

Solar Array

EP Thrust

Electrical

Thermal Waste

Best case
(solar min,

Drag Force

Required

Power

Heat

Area

edge-on)

3.3 m2

0.35 N

0.35 N

1.0 kW

0.6 kW

Radiator

Area

0.2 m2


![Table 49](paper-23-v1_images/table_49.png)
*Table 49*

Solar Array

EP Thrust

Electrical

Thermal Waste

Worst case
(solar max,

Drag Force

Required

Power

Heat

Area

broadside)

165+ m2

21 N

21 N

50+ kW

30+ kW

Radiator

Area

10+ m2

Figure 14: Drag-power-thermal cascade analysis for a 100 m-class inatable structure in
LEO, illustrating how atmospheric drag drives propulsion power requirements, which in
turn drive solar array sizing and thermal dissipation budgets. The cascade quanties the
interdependence of the AOCS, power, and thermal subsystems.

13.4
The Path to Flight Demonstration
2364

Among the roadmap milestones, the most ight-ready near-term demonstrator can be iden-
2365

tied by selecting the highest-TRL components from each technology area and integrating
2366

them into a single mission concept.
The analysis in Sections 58 suggests the following
2367

combination:
2368

 Capture mechanism: Gecko adhesive gripper (TRL 45, microgravity validated,
2369

400 kg capacity) Jiang et al. [2017], noting that this is a compliant end-eector on a
2370

conventional arm rather than a fully soft system.
2371

 Arm structure: Inatable multi-link arm based on the POPUP concept (TRL 3)
2372

Palmieri et al. [2023], using Vectran fabric links with FBG-instrumented webbing.
2373

 Structural health monitoring: FBG sensors in Vectran webbing (TRL 45 ground)
2374

Bally Ribbon Mills and Luna Innovations [2020], providing both SHM and propriocep-
2375

tive shape sensing via multicore FOSS principles Galloway et al. [2019].
2376

 Deployment: SMA-based hinge deployment for arm segments (TRL 89) Costanza
2377

and Tata [2020].
2378

This combination achieves an estimated system TRL of 34, limited by the inatable
2379

arm structure. A CubeSat-class (12U16U) demonstrator could validate the complete soft
2380

capture conceptdeploy inatable arm, acquire cooperative target, demonstrate FBG-based
2381

shape sensing during capturewithin a 35 year development timeline from programme ini-
2382

tiation. The mission would produce the rst orbital dataset on: (i) inatable arm deployment
2383

dynamics, (ii) FBG sensor performance in the LEO environment on a exible structure, and
2384

(iii) compliant capture contact dynamics. These three datasets address critical gaps C2, M6,
2385

and partially C1, making this demonstrator the highest-value single mission for advancing
2386

the eld.
2387

The key technical risk is the inatable arm structure: POPUP-class arms have been
2388

demonstrated only in simulation Palmieri et al. [2023], and the transition from analytical
2389

design to space-qualied ight hardware requires a focused engineering programme. However,
2390

the constituent technologiesVectran fabric, SMA deployment mechanisms, FBG sensors
2391

each have independent space heritage that de-risks the integration challenge.
2392

A critical observation from the roadmap analysis is that the fragmentation paradox (Sec-
2393

tion 3.1) will not be resolved by the ight demonstrator alone. The proposed CubeSat mission
2394

validates soft capture mechanics but does not generate fragmentation data. Resolving gap
2395

C1 requires a parallel ground campaign: hypervelocity and low-velocity impact testing with
2396

debris surrogate materials (solar panel fragments, aluminium honeycomb, carbon bre com-
2397

posite) at representative contact forces, comparing rigid grasp, compliant grasp, and soft
2398

envelopment capture modes. Parabolic ight campaigns can provide microgravity validation
2399

of the ground results. Together, the ight demonstrator and the ground fragmentation study
2400

would establish the quantitative evidence base that the soft ADR proposition currently lacks.
2401

14
Conclusions
2402

This survey has reviewed the state of the art in soft inatable robotic systems for space
2403

applications, covering eight enabling technology areas across 14 sections and synthesizing
2404

ndings from the active debris removal, space exploration, and robotic assembly domains.
2405

Four key ndings emerge from this comprehensive analysis.
2406

Finding 1: The Fragmentation Paradox Demands Soft Capture Solutions.
The
2407

space debris environment has reached a critical state: over 54,000 tracked objects larger than
2408

10 cm, an estimated 140 million fragments between 1 mm and 1 cm, and a total orbital mass
2409

exceeding 15,800 tonnes ESA Space Debris Oce [2025]. Active debris removal at the rate of
2410

at least 5 large objects per year is required to stabilize the LEO population Liou et al. [2010].
2411

Yet the dominant ADR approachrigid robotic capture, as exemplied by ClearSpace-1
2412

carries an unquantied but non-trivial fragmentation risk for tumbling targets (Section 3.1).
2413

Rigid capture of a debris object could generate new fragments, potentially exacerbating the
2414

very problem it aims to solve. Soft and compliant capture mechanisms (Section 3.2), by ab-
2415

sorbing kinetic energy rather than transmitting contact impulses, oer a system-level safety
2416

margin that rigid capture cannot provide. The absence of a quantitative soft-versus-rigid
2417

fragmentation comparison (gap C1) is the single most important open research question
2418

identied by this survey. Until this comparison is performed, the ADR community is select-
2419

ing capture mechanisms without the fundamental dataset needed for informed technology
2420

selection.
2421

Finding 2: Inatable Habitats Are Flight-Proven, with a Clear Path to Deep-
2422

Space Application.
BEAM's 8+ years of continuous operation on the International Space
2423

Station has conclusively demonstrated that pressure-stabilized inatable modules can sur-
2424

vive the LEO environment at TRL 9 (Section 4.1). The mass eciency advantage is decisive:
2425

39 kg m−3 for TransHab versus 137205 kg m−3 for metallic ISS modules Valle et al. [2019a].
2426

Vectran-based restraint layers provide specic strengths exceeding 2300 kN m kg−1, an or-
2427

der of magnitude beyond aerospace metals (Section 5.1). Current commercial programmes
2428

(Sierra Space LIFE) have demonstrated full-scale burst pressures of 77 psi, exceeding NASA
2429

structural requirements by 27% (Section 4.2). The path from BEAM to deep-space habitats
2430

requires addressing three challenges: radiation shielding (BEAM's 810× higher SPE dose
2431

versus metallic modules; Section 4.4), autonomous deployment reliability (BEAM's 25-burst,
2432

7-hour deployment was rescued by ISS crew; Section 6.3), and the 19× volume scale-up from
2433

BEAM's 16 m3 to a 300+ m3 deep-space transit habitat. Each challenge is substantive but
2434

bounded, with identied mitigation strategies (water-wall radiation shielding, deployment
2435

sequencing control, and multi-layer restraint engineering, respectively).
2436

Finding 3: The Space Vacuum Is a Resource, Not Merely an Obstacle.
The tra-
2437

ditional framing of the space environment as hostile to soft roboticspneumatic actuation
2438

loses its working medium, elastomers outgas, lubricants evaporateis being overturned by
2439

three developments. First, vacuum-gap electrostatic actuators Sirbu et al. [2025] achieve
2440

>4 N force at 0.7 g mass with >100 Hz bandwidth by using internal vacuum gaps as func-
2441

tional elements; these actuators require vacuum and are simpler in orbit than on Earth
2442

(Section 7.2). Second, the jamming-in-vacuum principle exploits the ambient orbital vac-
2443

uum as the external low-pressure reservoir for granular or layer jamming, eliminating the
2444

vacuum pump required in terrestrial implementations (Section 7.6); this remains a logical
2445

deduction requiring experimental validation (gap M1), but the physics is straightforward.
2446

Third, the very existence of pressure-stabilized inatable structures depends on the vacuum
2447

environment providing the pressure dierential that creates structural stiness. Together,
2448

these observations suggest that soft inatable robotic systems for space constitute a distinct
2449

engineering disciplinenot merely terrestrial soft robotics adapted for space, but a eld
2450

where the space environment enables capabilities impossible on Earth.
2451

Finding 4: The Critical Barrier Is System Integration, Not Individual Technol-
2452

ogy Maturity.
Perhaps the most signicant nding of this survey is negative: no single
2453

technology gap is a showstopper for the eld. Vectran and Kevlar are ight-proven for inat-
2454

able structures (TRL 9). SMA deployment mechanisms are ight-proven (TRL 89). FBG
2455

sensors have own on Proba-2 (TRL 78). iROSA-class exible photovoltaics power the
2456

ISS (TRL 9). Loop heat pipes transport multi-kilowatt thermal loads (TRL 9). Reaction
2457

wheels provide attitude control for the largest operational spacecraft (TRL 9). The barrier
2458

is at the interfaces: no programme has integrated FBG sensors into an inatable structure
2459

for ight; no programme is developing photovoltaics on inatable substrates; no theory ad-
2460

dresses AOCS for pressure-stabilized membranes; no interface enables rigid module assembly
2461

onto exible platforms. The eld suers from a fragmentation of its ownnot of debris, but
2462

of research communities. Soft roboticists, inatable structure engineers, space power spe-
2463

cialists, and GNC researchers each advance their disciplines without the cross-disciplinary
2464

programmes needed to integrate their outputs into ight-ready systems.
2465

This survey has attempted to bridge that fragmentation by reviewing all eight enabling
2466

technology areas through a single lens: the unifying thesis that the same high-strength fabric
2467

technologies (Vectran, Kevlar, Nextel) serve both active debris removal and space exploration
2468

applications.
The cross-domain connections identied throughoutthermal management
2469

informing actuator design (Section 10), MMOD protection materials serving as actuation
2470

substrates (Section 5), FBG sensing unifying habitat SHM and robotic proprioception (Sec-
2471

tion 8.1), and the dragpowerthermal cascade governing 100-metre-class platform architec-
2472

ture (Section 11.3)are insights that emerge only from the breadth of an integrative review.
2473

They cannot be seen from within any single technology discipline.
2474

The research roadmap proposed in Section 13.3 identies concrete near-term actions:
2475

jamming-in-vacuum validation, FBG ight demonstration on inatable webbing, perovskite/CIGS
2476

fold-deploy testing, rigid-exible interface prototyping, and gyroelastic theory extension. The
2477

most ight-ready integrated demonstratora gecko-adhesive gripper on an inatable arm
2478

with FBG structural health monitoringcould y within 35 years of programme initia-
2479

tion, generating the rst orbital dataset on soft inatable robotic capture. The longer-term
2480

visiona 10-metre inatable with integrated photovoltaics, assembly robots operating on
2481

inatable platforms, and AOCS-qualied pressure-stabilized structuresdenes a 15-year
2482

trajectory toward operational capability.
2483

The space debris crisis demands action on a timescale shorter than the 15-year technology
2484

roadmap allows. ClearSpace-1 and its successors will y rigid capture missions within this
2485

decade. The soft robotics and inatable structures communities must move from component-
2486

level demonstration to system-level integration with urgency commensurate with the prob-
2487

lem. The technologies exist; the integration does not. Closing the integration gaps identied
2488

in this survey is the dening challenge for the next generation of space robotics research.
2489

References
2490

ESA Space Debris Oce. ESA space environment report 2025. Technical report, European
2491

Space Agency, 2025. URL https://www.esa.int/Space_Safety/Space_Debris. Current
2492

debris environment statistics: 54,000 tracked objects, 15,800 tonnes, Starlink avoidance
2493

manoeuvre data.
2494

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2495

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