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ABSTRACT Due to their inherent lightweight and low packaged volume, inflatable/rigidizable structures have shown great promise for use in space. Utilizing these types of structures will result in considerable savings, particularly for deployment of large structures such as antennas, solar panels, solar sails, etc. in space. These systems can also be utilized for other non-conventional uses as well such as large optical devices for calibration of space related but ground-based equipment (e.g. lasers, etc.). One such device is an Inflatable Optical Sphere System (IOSS), which L’Garde designed and built. The objective of the IOSS program was to build, test and deploy in LEO a 160-in inflatable highly reflective and specular sphere that could survive the space environment for at least one year. The total system weight (including canister, inflation system and electronics) had to be kept under 50 pounds. IOSS was built out of an inflatable rigidizable membrane of only 1.2 mil. thick. L’Garde used one of its proprietary solid inflatants to maintain the internal pressure against dynamic forces after the initial rigidization by Nitrogen gas. An engineering IOSS unit was deployed and rigidized in a vacuum chamber in a ground-based experiment prior to the actual flight. The flight system was successfully deployed, inflated and rigidized in LEO on January 26, 2000 using a Minotaur launcher at Vandenberg, California. The IOSS could be seen from earth with naked eye and was used by AFRL for laser experiments. Figure 1 shows the 160-in IOSS engineering unit inflated and rigidized at 1 atmosphere. The rigidization method was based on aluminum foil laminate membrane in which the latter rigidizes when overstressed leaving a very stable rigid structure. As a result of the rigidization process, all packaging wrinkles are removed providing a highly specular surface. The main advantage of using the aluminum foil laminate rigidization method is its testability, i.e., the rigidizable flight elements can be ground-tested repeatedly prior to flight. ______________________________

This method has numerous other advantages such as unlimited shelf life, no need for power and is the most mature rigidization system for space use.

Figure 1. Rigidized IOSS Sphere

INTRODUCTION L’Garde has been investigating and manufacturing space rigidizable hardware for over 17 years. Based on L’Garde studies [Ref. 1] there are five different rigidization methods with the potential for future utilization in space. These are:

• Pressure rigidized aluminum foil • Sub Tg rigidizable thermoplastic and

thermoelastomeric composites • Hydrogel rigidization • Thermoset rigidization • UV rigidization Rigidizable structures are significantly more resistant to the hazardous space environment than constantly inflated structures. With increasing mission duration, leaks can develop in the structure mainly due to micrometeoroids and space debris. Rigidizable structures do not require internal pressure and therefore do not require any makeup inflatant.

An Inflatable Rigidizable Calibration Optical Sphere

Koorosh Guidanean, Gordon Veal L’Garde, Inc.*

*koorosh_guidanean@lgarde.com, gordon_veal@lgarde.com“Copyright 2003 by L’Garde, Inc. Published by the American Institute of Aeronautics and Astronautics, Inc.

44th AIAA/ASME/ASCE/AHS Structures, Structural Dynamics, and Materials Confere7-10 April 2003, Norfolk, Virginia

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Copyright © 2003 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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Figure 2 is a picture of a typical rigidizable space structure [Ref. 2]. It is a 24-foot long inflatable/ rigidizable truss that was designed and manufactured from a new laminate at L’Garde. The truss was compression tested at NASA/ LaRC and withstood a compression load of 556 pounds, 10% above its designed compression strength. The truss weighed only 9 pounds and consisted of separate inflatable/rigidizable legs (longerons and diagonals). The rigidization method is based on a sub-Tg method in which the 14-mils thick laminate rigidizes when exposed to the low temperatures of space. The rigidization method utilized for the IOSS fabrication was based on “pressure rigidized aluminum foil” method, which is a preferred technique when operational loads are relatively small.

PRESSURE RIGIDIZED ALUMINUM FOIL We have studied and compared major rigidization methods for space use and concluded that for low stress applications “pressure rigidized aluminum foil” is preferable to other methods in

terms of overall performance. One of the most important qualities of the “pressure rigidized aluminum foil” is its ability to be ground tested. Further, contrary to thermoset and UV based systems, the “pressure rigidized aluminum foil” is in its final cure stage and stable before launch. Therefore these composites have unlimited shelf life. The following summarizes the main attributes of “pressure rigidized aluminum foil” laminates: “Pressure Rigidized Aluminum Foil” Advantages:

• High modulus

• Mature concept, space qualified

• No power requirements

• Ground testable before flight

• No auxiliary equipment and hardware

• Positive rigidization control

• Unlimited pre deployment life time Stable material It must be noted that, none of the existing space rigidization methods is perfect and meet all requirements for all applications. As a result, “pressure rigidized aluminum foil” has its own shortcomings as follow: “Pressure Rigidized Aluminum Foil” Disadvantages• Thickness limited, limited scalability • 120 LB. Max compression • Requires high pressures for rigidization • High CTE

INFLATABE OPTICAL CALIBRATION SPHERE SYSTEM (IOSS)

The IOSS mission called for an optical calibration unit that could be deployed and operated in space with the following requirements:

• Highly reflective surface (~95%)

• Highly specular surface

• Retains geometric and optical properties for a minimum of 1 year in low earth orbit (LEO).

• Maximum weight of sphere, inflation tank, and inflatant mass to be not more than 12.5 lbs. (excl. the canister)

Figure 2. Sub Tg Rigidizable Truss

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• Total weight of payload (incl. Canister, electronics and inflation system) of 50 pounds

• 160 -in diameter deployed

• Dimensions of payload package: Maximum 20-in (L) x 14-in (D)

• Retains its optical, thermal and structural characteristics in LEO over the IOSS operational temperature range.

In order to insure that the optical balloon retains its geometry over a period of at least a year, we either had to have sufficient make up inflation gas to maintain the internal pressure or rigidize the balloon. Our initial analysis determined that the space debris and micrometeoroid environment in the orbit altitudes of interest is of a high enough flux to preclude the use of strictly inflatable balloons. The relatively small radius of curvature, coupled to the high film stress requirement for reflection in the visible portion of the spectrum, pushed the make up gas requirement for a 360-day mission considerably above the 50 lb weight limit. Figure 3 shows the make-up gas weight over the life of the optical sphere.

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Figure 3. Make-Up Gas Needs for Purely Inflatable Sphere

Thus, we concluded that the optimum balloon design would involve rigidization of the balloon surface, in precisely the same fashion as the Echo satellites of the early sixties. The balloon skin would be constructed of a laminate comprised of a thin polyimide layer and a thin aluminum foil layer. Initial inflation strains the aluminum, thus forming a thin spherical shell, which is strong

enough to sustain its shape and surface accuracy even after total loss of inflatant. This is a similar approach as the one used in our highly successful Inflatable Solar Array Program (ITSAT) to form the tubular deployment and support elements of the solar array blanket.

RIGIDIZABLE LAMINATE

Figure 4 shows the cross section of the rigidizable Kapton-aluminum laminate that was utilized to fabricate the OCS balloons for this program. Note that the main portion of the laminate which carries the load after rigidization is the inner aluminum foil. After the rigidization takes place, the outer metallized plastic layer (aluminized Kapton) will only serve as a reflective layer In order to minimize the thermal gradient across the balloon surface, we used a high emissive interior. As is shown in Figure 4 the VDA (vapor-deposited aluminum) faces space and the high emissive side faces the balloon interior.

Figure 4. Cross Section of Rigidizable Laminate Used

for IOSS

THERMAL ANALYSIS

A thermal analysis using the TRASYS [Ref. 3] and SINDA [Ref. 4] codes was conducted to determine the thermal loads during orbit and the temperature-time history of the balloon. Using the material properties from Table 1, the (static) temperature profile around the balloon surface is plotted in Figure 5 where psi is the angle as measured from the equator from zero to one π. It must be noted that the temperature profile shown in Figure 4 does not include contributions from albedo and earthshine. For the transient temperature profile of the balloon as it goes around the earth in orbit, we use the results of TRASYS and SINDA. These are shown in Figure 6 where we plot the temperatures of different points on the balloon as a function of time.

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TABLE 1. THERMAL PROPERTIES USED IN THE THERMAL ANALYSIS

External Emissivity 0.05 Internal Emissivity .90 Solar Absorptivity 0.15 Material Density 0.055 lb/in3

Specific Heat 0.262 But/lb-FE

Figure 5. Temperature Profile of a Static IOSS Balloon in Space

Figure 6. Temperature-Time History of Points on the IOSS Balloon for One Complete Orbit Around the Earth

As can be seen from the above results, the temperature environment expected is quite benign. These temperatures are well within the operating temperatures for the materials (Kapton, VDA, aluminum) and adhesives used. The maximum temperature expected is about 115ECand the temperature differential between the hottest and coldest points in the balloon is about 20EC. These temperatures also dictate the kind of inflatant that is suitable for the mission as discussed in the next section

INFLATANTS

Two different inflatants were used for deployment and rigidization of the balloon. As mentioned earlier, a strictly inflated balloon is not feasible for a 360-day minimum life. If the inflatant is nitrogen, the weight of the gas and tank to make up for the gas loss thru holes caused by debris and micrometeoroid impacts exceeds, by themselves, the 50-pound total system weight restriction. This situation led us to the rigidizable design. To deploy and rigidize the balloon, the pressure was raised by Nitrogen until the aluminum foil is strained beyond its yield point. This removes the

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packaging wrinkles, forms the spherical shape and provides a smooth-surfaced sphere. In order to assure proper balloon deployment and rigidization, it was necessary to deploy the balloon while the bus is in the sunlight portion of the orbit. This assured the balloon, and its inflatant, are warm enough to achieve the 0.05 psi rigidization pressure, which occurs at 39C (102F). At this point the inflatant is no longer required. We however used a solid inflatant in addition to the N2 gas because we did not have time and resources to test and generate sufficient data to demonstrate if the rigidized sphere could resist the dynamic pressure and further maintain its specular surface. The latter could have been caused by retraction of the Kapton layer, resulting in wrinkled surface. Figure 7 .shows vapor pressure-temperature characteristic of the solid inflatant used for the above purpose.

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Figure 7. Solid Inflatant Vapor Pressure

STRUCTURAL ANALYSIS

The on-orbit loads are drag and gravity gradient. There were some dynamic loads after deployment, but the sphere was pressure-stabilized at that time. These dynamics will then quickly decay under the stabilizing influence of drag and gravity gradient. The dynamic pressure at the lowest altitude (650 km) is only -1E-8 psi. This will result in a maximum compressive stress of 0.0017 psi. This is well below the local buckling stress allowable. Gravity gradient forces attempt to elongate the sphere along the vertical NADIR line. The maximum stress was 0.00016 psi, also well below the buckling stress.

SPACE ENVIRONMENT

Atomic Oxygen (AO) Nearly all inorganic coatings (e.g. Al, ITO, SiO2,SiOx) that are customarily used as AO protective coatings crack during packaging of typical inflatables. Since the balloon is rigidized, erosion of the Kapton layer at the cracks is acceptable. The main concern is the under cutting of the VDA that may occur in the vicinity of the VDA cracks resulting in significant changes of optical properties. AO is known to sneak through these cracks and attack the under layer material. The anticipated total atomic oxygen (AO) fluence level for a year under solar standard activity at 600 KM was calculated to be 1.14E+20 AO/Cm2. Assuming no protection from 1000A VDA, this level of AO should erode 0.14 mils of Kapton film in the vicinity of the cracks less than half the Kapton thickness. We assumed that the VDA surface remained highly reflective and specular even with the above undercutting.

UV and Ionizing Radiation

These radiations did not affect the main core of the laminate (aluminum). The Kapton film was protected by the top metallic layer from the UV effect and was unaffected by it. The annual equivalent 1 Mev electron charged particle (trapped electrons and protons) at 600 KM altitude and inclination angle of 60 degrees was calculated to be 3.39E+13 electrons/sq cm-year. This can cause an absorbed dose of 7.5E+7 on the surface of the Kapton laminate. The radiation threshold of Kapton film is known to be 10E+9 to 10E+10 rads. This is the radiation absorbed dose at which the properties of Kapton film start to be affected by the radiation. Based on the radiation levels discussed above, it appears that the material is well capable of resisting the radiation environment as well.

DESIGN DESCRIPTION

The configuration of the IOSS is shown in Figures 8 and 9. The primary components are the ejector and the canister. The canister contains the inflatable which, in turn, contains the control electronics, batteries and inflatant.

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Figure 8. Packaged IOSS with its Main Elements (Note: Balloon and door removed for clarity)

Figure 9. Deployed IOSS

EjectorThe ejector has two primary functions, 1) support the canister within the spacecraft bus, and 2) eject the canister at the desired velocity via command from the bus. The canister is held within the ejector by a cable attached to the balloon end cap and the ejector base. A bolt cutter for ejection severs this cable. Once the cable has been severed, the ejector spring provides the energy to eject the canister. The ejector spring was selected to provide an ejection velocity of 3-feet per second. This velocity was selected because it is typical of ejection velocities employed by L’Garde in the past (1.5 - 15 fps).

CanisterThe canister holds the balloon in its packaged configuration and supports it during launch. The canister size is based on using a conservative packaging factor for the laminate determined in advance. Using a conservative packaging factor and providing a proper inner surface of the canister prevented damage to the highly reflective specular balloon surface. The canister doors were released by severing a cable with a cable cutter. On-board electronics and batteries initiated the cable cutter. All parts remained attached to the balloon even after deployment including the severed cable so that all hardware will de-orbit with the balloon, thus eliminating all debris. Figure 10 shows the canister.

Figure 10. Packaged IOSS in its Canister

Control Electronics and Batteries

The control electronics and batteries used for the IOSS were based on previous designs of L’Garde’s Red Tigress II targets. A time delay/firing circuit was designed to control and initiate ejection, deployment and inflation of the IOSS in a pre determined and accurate sequence. Nickel-Cadmium batteries were selected instead of alkaline batteries for reliability purposes. These

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can be discharged and re-charged with an accurate measurement of the energy stored in the battery during the charging cycle.

TESTING Because of the rapid schedule from contract award date to product delivery (total of only 4 months!), a limited number of component and subsystem checkout tests were performed. In addition, a system deployment test to verify proper balloon deployment and inflation was conducted. Prior to the system deployment test, the unit underwent “qualification” level thermal cycling and random vibration tests. For the final deliverable flight unit, a short series of acceptance tests were performed. These tests included random vibration, thermal cycling, and an electrical functional checkout test.

Component Level Tests

The main component level tests were as follow: - Spring Rate Checkout Test - Ejection Pyro Bridge Wire Resistance Test - Battery Voltage Checks - Mechanical Fit Checks

Subsystem Levels Tests

The main subsystem levels tests were: - Balloon Canister Subsystem Ejection Test (to

verify proper balloon canister ejection and velocity)

- Inflation Subsystem Test (to verify proper activation of the balloon inflation system)

System Level Qualification Tests

- Qualification Level Thermal Cycling- This test was performed between -35° F and 160° F, with a minimum ramp rate of 2.5°F/minute for 18 cycles. The dwell time was 60 minutes at each extreme. This test was performed on all hardware except the heat-sensitive solid inflatant, and balloon subsystems.

- Qualification Level Random Vibration - After

final assembly of the entire system, the unit underwent a random vibration test. The qualification test level was 12.01 g RMS, over a frequency range of 20 to 2000 Hz and

performed in each of three mutually perpendicular axes. The test duration was three minutes in each axis. The unit was again visually and electrically checked prior to proceeding to the next phase of testing.

- Balloon Deployment and Inflation Test - To

assure proper balloon deployment and inflation, a deployment test was performed in a large vacuum chamber. The test was performed in a one-g environment with the complete unit attached to the wall of the chamber so the balloon could be deployed sideways. The balloon end cap was supported via a pull rope and a carriage track mounted to the vacuum chamber ceiling. The test was documented using several high-speed film cameras and several video cameras. It was performed at the maximum altitude capability of the chamber of 110,000 feet. (0.10 PSI). Figures 11-13 .show the IOSS canister and the balloon during vacuum deployment test.

Figure 11. IOSS In Its Canister In Vacuum Chamber

Figure 12. IOSS During Full Deployment

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Figure 13. IOSS after Full Deployment and Rigidization in Vacuum Chamber

System Level Acceptance Tests

The main system level acceptance tests were identical to those performed for the qualification tests except the actual deployment of the balloon itself.

DEPLOYMENT OF IOSS INTO SPACE The IOSS flight unit was successfully deployed, inflated and rigidized in LEO on January 26, 2000 from Vandenberg, California using a Minotaur launcher. The IOSS could be seen from earth and was used by AFRL for laser experiments. Figures 14-17 show the IOSS canister during integration into the Minotaur launcher. Based on the information from AFRL, the IOSS survived in low earth (LEO) orbit for more than a year. It eventually burned during reetntry. No official test data has been published about the performance of the balloon in space. However, we were told the IOSS functioned very well and was utilized effectively in AFRL laser experiments.

Figure 15. IOSS is being Attached to the Minotaur Launcher

Figure 16. Close-Up of IOSS to the Minotaur Launcher

Figure 17. IOSS on the Minotaur Launcher

Figure 14. The MinotaurFigure 14. The Minotaur

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ACKNOWLEDEGMENTS The authors are indebted to the following individuals for their contributions: •Major Jim Lee, AFRL •LtCol Brent Richert, AFRL •Dr. Mike Wood, Weber State University •Scott Schoneman, Orbital Sciences

REFERENCES 1- K. Guidanean, “Large Radar Antenna, Comparison of Space Rigidization Techniques”, L’Garde Technical Report LTR-00-KG-012, September 2000. 2- K. Guidanean, D. Lichodziejewski , An Inflatable Rigidizable Truss Structure Based on New Sub-Tg Polyurethane Composites, AIAA-02 1593 , 43th SDM conference April 21-25, 2002. 3- TRASYS – Thermal Radiation Analyzer System. National Aeronautics and Space Administration, Johnson Space Center, JSC-22964, April 1998. 4- SINDA – Systems Improved Numerical Differencing Analyzer and Fluid Integrator, NASA-COSMIC, University of Georgia, Athens, Georgia, Cosmic Program # MSC-21528.