Frontiers of Engineering: Reports on Leading-Edge Engineering from the 2015 Symposium (2016)

Realizing Large Structures in Space

JEREMY BANIK
Air Force Research Laboratory

Since the dawn of space access in 1957 the size of spacecraft payloads and solar arrays has steadily grown. Yet the demand for larger antenna arrays and telescope mirrors and higher power on spacecraft continues to outpace availability.

BACKGROUND

Large structures in space have one simple purpose: to support payloads that manipulate the electromagnetic spectrum. Telescopes operate at short wavelengths for astronomy and surveillance missions, from ultraviolet to near infrared. Star-shades block visible sunlight for direct imaging of Earth-like planets. Sunshields reject the shortwave spectrum to keep instruments cool. Solar sails reflect photons in the short wave, generating constant propulsion for interplanetary exploration and station-keeping maneuvers.

For practical Earth-based applications, communication and radar antennas transmit and receive microwaves for communications, coarse imaging, and object tracking. The larger the antenna or radar in space, the smaller the antenna on a soldier’s back and the greater number of objects tracked. Photovoltaic solar arrays convert photons to electrons for charging satellite batteries. Radio frequency missions require larger antennas to generate higher-resolution radar imaging for Earth science, and to keep up with the higher data throughput demands of modern hand-held devices.

In each of these examples the larger the space structure, the more effective the mission data return. Exoplanet discovery is no exception. Larger optical mirrors and bigger starshade occulters are needed to detect both a larger number of

stars, thus increasing Earth-like exoplanet discovery opportunities, and a broad range of the electromagnetic spectrum for exoplanet spectral characterization.

But the design, construction, and deployment of larger antennas and mirrors are fraught with highly specialized challenges. The largest commercial radio frequency antenna on-orbit is 22 meters in diameter, and the largest space telescope mirror is 6.5 meters. These structures must be designed to hold extreme dimensional precision and stability tolerances.

Optical and radio signal quality are directly related to the precision of the surface from which the signal emanates. The larger the structure, the more difficult it is to achieve a given figure precision. Radio frequency missions operate on long wavelengths so precision requirements are not as stringent as for optical missions, but because signal gain scales inversely with the square of wavelength, radio antennas require larger apertures than optical.

Once unfolded in space, these structures face extreme temperature swings that can cause large static and dynamic dimensional changes. Spacecraft in a geosynchronous orbit endure daily temperature swings from −200°C to +200°C over a typical 15-year lifetime. Materials must exhibit a low coefficient of thermal expansion, and assembly interfaces are extensively tested to control the characteristics of such expansion.

CURRENT CAPABILITIES AND CONSTRAINTS

Figure 1 illustrates the indirect relationship between structure size and dimensional precision. As this ratio grows, payload cost escalates. Some of the highest-performing space structures to date are represented on this chart, yet they are restricted to relatively low diameter-to-precision ratios when compared to future needs in the tens to hundreds of meters.

Aside from dimensional precision challenges, large space structures must also be folded compactly for a violent trip to orbit. Once designed, built, and stowed in a 5 meter launch vehicle fairing, these payloads endure 10 to 70 g peak accelerations during the 10-minute trip to low Earth orbit, reaching a velocity of 7 km/s. Once on-orbit, the structure must unfold precisely and reliably.

As an example, an exoplanet starshade must unfurl from 5 meters to 34 meters into a shape that is within 0.1 mm precision, approximately the width of a human hair, across a span similar to an eight-lane interstate overpass. Figure 2 and Table 1 show the relative scale of the largest launch vehicles, the typical large structures that must stow in them, and the respective packaging ratios.

The mechanical approach thus far for realizing large communication antennas, telescopes, and radar antennas has generally been to fold a deep truss structure and self-deploy using multiple pin-clevis joints, motors, torsion springs, and dampers. This approach has led to incremental improvements in size, weight, and power over the past 50 years, but these heritage mechanisms and structural support schemes are reaching size and mass limits. Adding hinges to package

FIGURE 1 Size and surface figure precision are indirectly related in space structure design. Boxes indicate the current art of the possible. Comm = communications; JWST = James Webb Space Telescope; MOIRE = Membrane Optical Imager for Real-Time Exploitation; RMS = root mean square.

larger payloads into these limited launch volumes is causing reliability concerns and cost escalation.

EMERGING APPROACHES

Two structural design techniques show high payoff potential: (1) tension-aligned antennas and optics and (2) high-strain composite mechanisms. These methods have been used successfully for decades, but in very limited form because of the absence of high-strength, high-stiffness carbon fiber composites and robust analytical and test tools.¹

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¹ Two historical space structures that used lower-stiffness glass and aramid fiber composites are the Continuous Longeron Mast of the 1960s and the Wrap-Rib reflector of the 1970s.

FIGURE 2 Scale of the largest launch vehicles (left) relative to representative large space structures (right). Images reprinted with permission.

TABLE 1 Packaged Size of Typical Precision Space Mirror, Shade, and Antenna

JWST = James Webb Space Telescope; m = meter.

Tension-Aligned Antennas and Optics

As more sophisticated computerized testing and analysis tools became available and the use of high-strength carbon fibers in aircraft became prevalent, feasible new space architectures began to surface, enabling the ground testing of the Innovative Space-based Radar Antenna truss in 2007 (Lane et al. 2011), the flexible unfurlable and refurlable lightweight (FURL) solar sail in 2010 (Banik and Ardelean 2010), and in 2013 the Membrane Optical Imager for Real-Time Exploitation (MOIRE) brassboard telescope (Domber et al. 2014). Most recently, the roll-out solar array (ROSA; Spence et al. 2015) was manifested for a spaceflight experiment to the International Space Station in 2016.

Other advanced concepts currently under development by government and industry all share high-strain composite features for folding functionality and/or tension as the means of structural stability: a low-cost multiarm radial composite

FIGURE 3 Ground deployment sequence of a starshade demonstration model shows a perimeter truss collapsed against a central drum (left). As the truss articulates radially (center) petals begin to rotate into position. Finally the perimeter truss is fully locked out (right), providing full support to the four petals. SOURCE: NASA (2015).

radio-frequency reflector (Footdale and Banik 2016), an isogrid column (Jeon et al. 2016), a starshade occulter (Figure 3; NASA 2015), an extremely high expansion deployable structure (Warren et al. 2007), a triangular rollable and collapsible mast (Banik and Murphey 2010), and a tensioned planar membrane antenna (Warren et al. 2015).

High-Strain Composite Mechanisms

High-strain composites are defined as thin carbon and glass fiber polymer matrix laminate materials used to construct shell structures that undergo large elastic deformations during folding and then release the stored strain energy to enforce deployment.

Architectures constructed from these materials have 7× greater deployment force, 20× greater dimensional stability, and 4× higher stiffness compared to traditional metallic flexure mechanisms (Murphey et al. 2013, 2015; Welsh et al. 2007). Moreover, when compared to traditional pin-clevis-type hinges, the payoff is a reduced mechanism part count, less susceptibility to binding, and more robust deployments. The hinges have greater lateral and torsion compliance during the transition from folded to deployed, a critical transition when there is a high risk of binding (e.g., when asymmetric solar heating and resulting expansion induce side loads on hinges). High-strain composite hinges can operate through this state and achieve a repeatable, dimensionally stable locked-out condition due to the near-zero coefficient of thermal expansion of carbon fibers.

Combined with the kinematic determinacy of a tensioned antenna or optic, these technologies are cracking open the door to a new era of space structures where 50-meter telescopes, 100-meter antennas, and megawatt-class solar arrays are all feasible.

LONG-TERM POSSIBILITIES AND CONSIDERATIONS

Despite the potential of tension-aligned payloads and high-strain composite mechanisms, even these will eventually reach limitations in size scaling, mass efficiency, and dimensional stability. It will be necessary to push beyond these limits to meet long-term civil and military space needs.

As the promise of robotic assembly and in-space additive manufacturing technologies is only beginning to emerge, the best tack is not yet clear. Certainly the unique realities and constraints of space flight must remain front and center in any pursuit of exciting new technologies.

Few industries are more risk-averse than those involved in space flight. NASA and Department of Defense program managers regularly spend hundreds of millions (sometimes billions) of dollars on a single spacecraft to try to ensure mission success. For example, the price tag on the 6.5 meter James Webb Space Telescope has reportedly reached $8.7 billion during the 16 years from inception to launch (Leone 2011). If this paradigm holds, then a 20-meter space telescope will remain in development for 87 years and cost $47.5 billion (Arenberg et al. 2014).

It is a spiraling effort. As more money is spent, additional testing and analysis are required to try to increase the certainty of spacecraft success. Schedules are then drawn out, further adding to the costs.

Spaceflight is a one-shot business. Hundreds of critical systems must work together flawlessly the first time or the mission is lost. Deployable structures are notorious as one of the highest sources of failure. Mission managers therefore expend great effort, time, and resources on testing in space simulation chambers. But even then the effects of gravity and the lack of a truly representative space environment always raise questions about the validity of these tests despite decades of experience and evidence.

Of course, the pursuit of innovation should not be deterred by these challenges. Rather, they should motivate the continued evaluation of new architectures against all measures of success—not only structural performance but also cost factors such as ease of ground testing and validation, simplicity of analysis methods, and reduced quantity of mechanical interfaces and unique parts.

It is difficult to quantify each key cost factor in the early conceptual design phase, but this is precisely when critical design decisions that most affect cost are made. Cost evaluation metrics are therefore essential. Until they are available, rational comparison of competing structural architectures must rely on structural performance metrics along with subjective cost assessments. A common list of such metrics is provided in Table 2; note the highly sophisticated telescope mission cost metric (bottom row).

TABLE 2 Common Metrics for Evaluation of Large Space Structure Performance

Metric	Description	Equation
Packaging ratio	deployed length/stowed length	Ld/Ls
Linear packaging density	deployed size/stowed volume	D/V
Areal packaging density	deployed area/stowed volume	A]V
Beam performance index (Murphey 2006)	strength moment, bending stiffness, linear mass density
Solar array scaling index (Banik and Carpenter 2015)	acceleration load, frequency, boom quantity, length, area, blanket areal mass density, total mass
Mass efficiency	diameter/mass	D/m
Surface precision	diameter/root mean square (RMS) figure error	D/RMS
Dimensional stability	coefficient of thermal expansion
Telescope mission cost (Arenberg et al. 2014; Stahl et al. 2012)	diameter, wavelength, temperature of operation

CONCLUSION

The challenges to realize large space structures are great, but if they are successfully addressed the opportunities will be well worth the effort and cost. No doubt, one of the most exciting possibilities is discovery of Earth-like planets that might have sustained life—or perhaps still do. . . .

REFERENCES

Arenberg J, Atkinson C, Breckinridge J, Conti A, Feinberg L, Lillie C, MacEwen H, Polidan R, Postman M, Matthews G, Smith E. 2014. A new paradigm for space astrophysics mission design. SPIE Astronomical Telescopes and Instrumentation, June 22–27, Montreal.

Banik J, Ardelean E. 2010. Verification of a retractable solar sail in a thermal-vacuum environment. 51st AIAA Conference on Structures, Structural Dynamics, and Materials, April 12–15, Orlando, FL.

Banik J, Carpenter B. 2015. Analytic representation of space solar array scaling performance. 42nd IEEE Photovoltaic Specialist Conference, June 14 –19, New Orleans.

Banik J, Murphey T. 2010. Performance validation of the triangular rollable and collapsible mast. 24th AIAA/USU Conference on Small Satellites, August 9 –12, Logan, UT.

Domber J, Atcheson P, Kommers J. 2014. MOIRE: Ground test bed results for a large membrane telescope. 1st AIAA Spacecraft Structures Conference, January 13 –17, National Harbor, MD.

Footdale J, Banik J. 2016. Design and deployment testing of the Multi-Arm Radial Composite (MARCO) Reflector Antenna. 3rd AIAA Spacecraft Structures Conference, January 4 – 8, 2016, San Diego.

Jeon S, Banik J, Peterson M. 2016. Design and validation of a deployable isogrid. 3rd AIAA Spacecraft Structures Conference, January 4 – 8, 2016, San Diego.

Lane S, Murphey T, Zatman M. 2011. Overview of the Innovative Space-based Radar Antenna Technology Program. AIAA Journal of Spacecraft and Rockets 48(1):135 –145.

Leone D. 2011. NASA: James Webb Space Telescope to now cost $8.7 billion. Space News, 27 August. Available at http://www.space.com/12759-james-webb-space-telescope-nasa-cost-increase.html.

Murphey T. 2006. Booms and trusses. In: Recent Advances in Gossamer Spacecraft, ed. Jenkins CHM. Progress in Astronautics and Aeronautics Series, vol. 212, pp. 1– 43. Reston, VA: American Institute of Aeronautics and Astronautics.

Murphey T, Peterson M, Grigoriev M. 2013. Four point bending of thin unidirectional composite laminas. 54th AIAA Conference on Structures, Structural Dynamics, and Materials, April 9, Boston.

Murphey T, Francis W, Davis B. 2015. High strain composites. 2nd AIAA Spacecraft Structures Conference, January 5–9, Kissimmee, FL.

NASA JPL [National Aeronautics and Space Administration Jet Propulsion Laboratory]. 2015. Exo-S Starshade Probe-Class: Exoplanet Direct Imaging Mission Concept Final Report. Pasadena. Available at http://exep.jpl.nasa.gov/stdt/Exo-S_Starshade_Probe_Class_Final_Report_150312_URS250118.pdf.

Spence B, White S, Takeda R, LaPointe M, Schmid K, Kiefer S, Carter G, Paskin A, Allmandinger T, Douglass M, LaCorte P. 2015. Flexible blanket solar arrays: Next-generation game-changing technology. 42nd IEEE Photovoltaic Specialist Conference, June 17, New Orleans.

Stahl HP, Henrichs T, Luedtke A, West M. 2012. Update on multivariable parametric cost models for ground and space telescopes. SPIE Proceedings 8442: Space Telescopes and Instrumentation 2012: Optical, Infrared, and Millimeter Wave, September 21.

Warren P, Dobson B, Silver M. 2007. Experimental characterization of an extremely high expansion deployable gossamer structure. 48th AIAA Conference on Structures, Structural Dynamics, and Materials, April 23 –26, Honolulu.

Warren P, Steinbeck J, Minelli R, Mueller C. 2015. Large, deployable S-band antenna for a 6U Cubesat. 29th AIAA/USU Conference on Small Satellites, August 8–13, Logan, UT.

Welsh J, Mayes J, Biskner A. 2007. Experimental and numerical failure predictions of biaxially-loaded quasi-isotropic carbon composites. 16th International Conference on Composite Materials (ICCM), July 8 –13, Kyoto.