Cryogeneic insulation and TPS system demonstration for RLV reusable composite hydrogen tank

Because of the low coefficient of thermal expansion (CTE) of graphite/epoxy laminates, traditional polyurethane foams used for aluminum or steel cryo-tankage are not compatible with composite tank materials. Insulation of such tanks requires development of an entirely new insulation system which can withstand the thermal strain of cryogenic conditions while providing enough support at higher temperatures to sustain TPS materials during reentry. This paper presents the results of the development, screening, selection, evaluation, and characterization of cryogenic insulation materials in conjunction with TPS system materials and composite substrate for use as a composite liquid hydrogen tank system. Data are presented for commercially available insulation materials as well as a customized bilayer foam-filled honeycomb core insulation system designed to meet the stringent requirements of the tank cylindrical surface under TPS materials. Necessary elements of a suitable cryogenic insulation/TPS system are identified for a reusable composite hydrogen tank, and this paper shows that these elements have been demonstrated for the insulation/TPS configurations selected for the Rockwell RLV design. Introduction and Background Successful design and flight of a reusable launch vehicle demands the use of materials which are of the lightest weight available, including extensive use of composite materials. The largest single structural element of the RLV is the liquid hydrogen tank, and it is imperative that the vehicle incorporate a composite hydrogen tank i f at all possible. This paper describes materials and structures testing that was performed by Rockwell Space Systems Division in Downey, California under a cooperative agreement with 925 Copyright © 1996 by Rockwell International. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. NASA-MSFC' 2 in order to demonstrate that a composite hydrogen tank can be successfully fabricated and utilized for a reusable launch vehicle application. In the design of a single-stage-to-orbit launch vehicle, it is desirable to use a composite hydrogen tank for a lightweight, rigid structure with improved payload margin. A design trade study comparison between welded aluminum and graphite/epoxy tank structures for Rockwell's reusable launch vehicle (RLV) resulted in a weight savings of 15 to 20 percent for the composite tank, depending upon design parameters used. In order to fully realize this weight savings, however, the hydrogen tank must be usable as a system, including cryogenic insulation, thermal protection system materials, and associated cryogenic system hardware such as feedlines and seals. Traditional expendable aluminum cryogenic systems have used closed-cell polyurethane foams for cryogenic insulation and to prevent cryopumping of liquefied air onto cryogenic surfaces. The repeated cyclic use of polyurethane foam on metals was not investigated in this project, but it has traditionally been limited to expendable cryogenic hardware. In the few cases where polyurethane foam is used for repeated applications (such as ground test hardware), the foam is repaired and often removed and reapplied because of cracking and disbonds. This would not be acceptable for an RLV system, which must operate within strict cost and schedule constraints. Regardless of the experience with metallic systems, polyurethane foams are not compatible with low-CTE composite tank materials, and insulation of such a tank requires development of an entirely new insulation concept which can withstand the thermal strain and load of cryogenic conditions while providing enough support to sustain TPS materials under reentry conditions. System Requirements for RLV Cryogenic Insulation RLV Configuration Drives Insulation Design Insulation of a composite hydrogen tank requires special materials because of the inherently low coefficient of thermal expansion, (CTE) of the composite material, as compared to the CTE of the plastic foam materials or the CTE of aluminum or steel. When the temperature drops in a composite cryotank, the insulation material tries to shrink but is stopped by the composite which does not shrink. This difference in thermal shrinkage is a major design challenge for the RLV system. The lightest-weight vehicle configuration for the RLV uses an "integral" tank , which requires that the tank wall also function as the outer fuselage structure of the vehicle. For improved operability, and in order to simplify installation of the insulation and TPS, the stringers and frames that stiffen the tank wall are located inside the tank. This provides a smooth outer tank wall mold line, which greatly facilitates attachment of insulation and TPS to satisfy aerodynamic smoothness requirements for reentry performance. The bare inner tank wall also permits ready access to frames, stringers and instrumentation between flights, avoids problems of insulation particles in the propulsion system, and allows easy structural inspection via removal of a manhole cover at one end of the tank. The tank system design requires that the cryoinsulation provide sufficient support to allow attachment of a high-temperature thermal protection system material (such as ceramic tiles and/or blankets). The system also must be resistant to moisture, frost, cryopumping, and be durable enough to survive hundreds of missions of fueling, liftoff and reentry without significant inspection, repair or rework. Cylinder and Dome Have Different Requirements Two distinctly different regions are defined over the hydrogen tank surface for application of cryogenic insulation. The tank cylinder section comprises most of the acreage of the tank, and i t has the more severe design requirements because it is exposed to the conditions of reentry heating. The second region is the protected dome bulkheads of the tank at either end, consisting of elliptically curved surfaces, protected by surrounding fuselage sections, not exposed to heating beyond that of the structure, and possibly further protected by a helium purge system. The dome environment also applies generally to composite feedlines and other composite propulsion elements which are located inside the structure. Both areas will experience the same strain compatibility problems a t cryogenic temperatures, but the dome areas do not need to satisfy the 400°F TPS bondline upper limit for the cylinder. Strain Compatibility is Key The strain compatibility requirement for the cryogenic insulation system proved to be the most difficult to meet with currently-available foam materials. For most plastic closed-cell foams, the thermal mismatch (CTE difference) between the foam and the composite substrate amounted to 14 x Iff inches per inch of strain per degree Fahrenheit of temperature change. For a typical hydrogen fill, the temperature change is approximately 500°F, resulting in a relative strain of 7000 microinches per inch, or 7000 microstrain, purely due to temperature change. This effect is exacerbated by the mechanical strain due to pressurization of the tank, since the two both result in growth of the tank wall relative to the foam surrounding the tank. Since the tank wall is thin for minimum weight, the pressurization strain is quite significant, nearly 6000 microstrain for the current design. The co-incidence of this combined strain level with an extremely cold operating environment puts a severe test on any potential insulation material or system under consideration for this application. An additional problem for the designer is that foam insulation materials are not generally very strong to begin with, and their strength at cryogenic temperatures must increase dramatically to compensate for huge increases in modulus. The modulus increases, combined with a high strain environment and relatively low strength result in foam failure in every case we have tested.