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NASA NTRSDate not supplied

TPSAS-NF1676L-13579-DND

During aerospace vehicle conceptual and preliminary design, empirical non-optimum factors are typically applied to predicted structural component weights to account for undefined manufacturing and design details. Non-optimum factors are developed here for 32 aluminum-lithium 2195 orthogrid panels comprising the liquid hydrogen tank barrel of the Space Shuttle External Tank using measured panel weights and manufacturing drawings. Minimum values for skin thickness, axial and circumferential blade stiffener thickness and spacing, and overall panel thickness are used to estimate individual panel weights. Panel non-optimum factors computed using a coarse weights model range from 1.21 to 1.77, and a refined weights model (including weld lands and skin and stiffener transition details) yields non-optimum factors of between 1.02 and 1.54. Acreage panels have an average 1.24 non-optimum factor using the coarse model, and 1.03 with the refined version. Non-optimum factors are also calculated for the aluminum-alloy gores that compose the ogive and ellipsoidal bulkheads of the Space Shuttle Super-Lightweight Tank propellant tanks. Minimum values for actual gore skin thicknesses and weld land dimensions are extracted from component production drawings, and are used to predict reference gore weights. These actual skin thicknesses are also compared to skin thicknesses predicted using classical structural mechanics and tank proof-test pressures. Both coarse and refined weights models are developed for the gores. The coarse model is based on the proof-test skin thicknesses, and the refined model uses the actual skin thicknesses and weld land details. Both models include manufacturing tolerance weight estimates. To determine the gore non-optimum factors, these reference weights are then compared to flight hardware weights reported in a mass properties database. The gore non-optimum factors computed using the coarse weights model range from 1.18 to 2.36, with an average non-optimum factor of 1.68. Application of the refined weights model yields non-optimum factors between 1.00 and 1.45, with an average non-optimum factor of 1.13. To demonstrate their utility, these calculated non-optimum factors are then used to predict more realistic (i.e., heavier) gore weights for a proposed heavy-lift launch vehicle's propellant tank bulkheads.

K Chauncey Wu↗
NASA NTRS2003

Vacuum Plasma Spray Forming of Copper Alloy Liners for Regeneratively Cooled Liquid Rocket Combustion Chambers

Vacuum plasma spray (VPS) has been demonstrated as a method to form combustion chambers from copper alloys NARloy-Z and GRCop-84. Vacuum plasma spray forming is of particular interest in the forming of CuCrNb alloys such as GRCop-84, developed by NASA s Glenn Research Center, because the alloy cannot be formed using conventional casting and forging methods. This limitation is related to the levels of chromium and niobium in the alloy, which exceed the solubility limit in copper. Until recently, the only forming process that maintained the required microstructure of CrNb intermetallics was powder metallurgy formation of a billet from powder stock, followed by extrusion. This severely limits its usefulness in structural applications, particularly the complex shapes required for combustion chamber liners. This paper discusses the techniques used to form combustion chambers from CuCrNb and NARloy-Z, which will be used in regeneratively cooled liquid rocket combustion chambers.

Zimmerman, Frank↗
NASA NTRS2013

Process Optimization of Bismaleimide (BMI) Resin Infused Carbon Fiber Composite

Engineers today are presented with the opportunity to design and build the next generation of space vehicles out of the lightest, strongest, and most durable materials available. Composites offer excellent structural characteristics and outstanding reliability in many forms that will be utilized in future aerospace applications including the Commercial Crew and Cargo Program and the Orion space capsule. NASA's Composites for Exploration (CoEx) project researches the various methods of manufacturing composite materials of different fiber characteristics while using proven infusion methods of different resin compositions. Development and testing on these different material combinations will provide engineers the opportunity to produce optimal material compounds for multidisciplinary applications. Through the CoEx project, engineers pursue the opportunity to research and develop repair patch procedures for damaged spacecraft. Working in conjunction with Raptor Resins Inc., NASA engineers are utilizing high flow liquid infusion molding practices to manufacture high-temperature composite parts comprised of intermediate modulus 7 (IM7) carbon fiber material. IM7 is a continuous, high-tensile strength composite with outstanding structural qualities such as high shear strength, tensile strength and modulus as well as excellent corrosion, creep, and fatigue resistance. IM7 carbon fiber, combined with existing thermoset and thermoplastic resin systems, can provide improvements in material strength reinforcement and deformation-resistant properties for high-temperature applications. Void analysis of the different layups of the IM7 material discovered the largest total void composition within the [ +45 , 90 , 90 , -45 ] composite panel. Tensile and compressional testing proved the highest mechanical strength was found in the [0 4] layup. This paper further investigates the infusion procedure of a low-cost/high-performance BMI resin into an IM7 carbon fiber material and the optical, chemical, and mechanical analyses performed.

Ehrlich, Joshua W.↗
NASA NTRS1975

Performance of a hypersonic hot fuselage structure with a carbon dioxide frost projected, nonintegral cryogenic tank

A model which consisted of a hot structure and a nonintegral tank protected by a carbon dioxide frost thermal protection system was tested under the following conditions: (1) room temperature loading and (2) heating and loading corresponding to the Mach 8 flight of an air-breathing launch vehicle. In the simulated flight tests, liquid nitrogen inside the tank was withdrawn at the rate fuel would be consumed. Prior to each simulated flight test, carbon dioxide was cryodeposited in the insulation surrounding the tank; during the tests, subliming CO2 frost absorbed heat and provided a purge gas for the space between the tank and the structure. A method of flame spraying the joints between panels with a nickel-aluminum material was developed to prevent excessive leakage of the purge gas through the outer structure. The tests indicated that the hot structure (with a joint repaired by riveting), the nonintegral tank and suspension system, and the carbon dioxide frost thermal protection system provide a workable concept with predictable performance.

Sharpe, E. L.↗
NASA NTRS2018

Chapter 12: Materials for Liquid Propulsion Systems

Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks which provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de-Laval nozzle) which forces them to accelerate; then as the nozzle flares outwards, they expand and further accelerate. It is the mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, which produce thrust according to Newton's third law: for every action there is an equal and opposite reaction. Solid rocket motors are cheaper to manufacture and offer good values for their cost. Liquid propellant engines offer higher performance, that is, they deliver greater thrust per unit weight of propellant burned. They also have a considerably higher thrust to weigh ratio. Since liquid rocket engines can be tested several times before flight, they have the capability to be more reliable, and their ability to shut down once started provides an extra margin of safety. Liquid propellant engines also can be designed with restart capability to provide orbital maneuvering capability. In some instances, liquid engines also can be designed to be reusable. On the solid side, hybrid solid motors also have been developed with the capability to stop and restart. Solid motors are covered in detail in chapter 11. Liquid rocket engine operational factors can be described in terms of extremes: temperatures ranging from that of liquid hydrogen (-423 F) to 6000 F hot gases; enormous thermal shock (7000 F/sec); large temperature differentials between contiguous components; reactive propellants; extreme acoustic environments; high rotational speeds for turbo machinery and extreme power densities. These factors place great demands on materials selection and each must be dealt with while maintaining an engine of the lightest possible weight. This chapter will describe the design considerations for the materials used in the various components of liquid rocket engines and provide examples of usage and experiences in each.

Halchak, John A.↗
NASA NTRS2024

Enantioselective Three-Dimensional High-Performance Liquid Chromatographic Determination of Amino Acids in the Hayabusa2 Returned Samples from the Asteroid Ryugu

The chirality of amino acids in extraterrestrial materials may provide an insight into the origin of the essential L-enantiopure amino acids in the terrestrial biosphere. In 2020, the Hayabusa2 mission succeeded in bringing back surface materials from the C-type asteroid (162173) Ryugu to the Earth. Amino acids were one of the targeted organic molecules to be studied in the Ryugu samples. To analyze the various structural isomers of amino acids, which were expected to be present, from the limited amount of the returned samples, the development of a highly-sensitive and selective analytical method was necessary. In the present study, a three-dimensional high-performance liquid chromatography (3D-HPLC) system has been developed for the enantioselective determination of five proteinogenic and three non-proteinogenic amino acids in the Ryugu samples, in which amino acids in the sample were separated by reversed-phase, anion-exchange and enantioselective columns after the fluorescence derivatization with 4-fluoro-7-nitro-2,1,3-benzoxadiozole. The applicability of the analytical system to the extraterrestrial samples was evaluated by analyzing several types of carbonaceous meteorites before applying the system to the Ryugu samples. In the analysis of the Ryugu samples, all of the target amino acids were successfully determined quantitatively. Non-proteinogenic amino acids including 2-amino-n-butyric acid, isovaline and norvaline, rarely present in the terrestrial environment, were found as almost racemic mixtures with 47.1 to 55.2%L.

Amino acids↗
NASA NTRS1973

Liquid rocket engine centrifugal flow turbopumps

Design criteria and recommended practices are discussed for the following configurations selected from the design sequence of a liquid rocket engine centrifugal flow turbopump: (1) pump performance including speed, efficiency, and flow range; (2) impeller; (3) housing; and (4) thrust balance system. Hydrodynamic, structural, and mechanical problems are addressed for the achievement of required pump performance within the constraints imposed by the engine/turbopump system. Materials and fabrication specifications are also discussed.

Source record↗
NASA NTRS1976

International Conference on Rapidly Quenched Metals, 2nd, Massachusetts Institute of Technology, Cambridge, Mass., November 17-19, 1975, Proceedings

The range of topics covered by the papers includes: rapid quenching (and specifically splat quenching) of metals, spin quenching, levitation melting, rapid quenching from liquid state; amorphous metal phases and metastable phases, metallic glasses, ribbons, splats, films, metal ribbon reinforced resins; crystallization of metallic glasses, and ferromagnetic amorphous alloys. Structural models for amorphous metals, splat-quenching equipment, and measurements of the properties of amorphous and noncrystalline metals and metallic glasses are also discussed, in addition to: fatigue events (crack propagation) in amorphous metals, fracture behavior in rapidly quenched metals, ductile superconducting Cu-Nb-Sn alloys, and semiconducting amorphous V205 alloy. Individual items are announced in this issue.

Source record↗
NASA NTRS1981

Low-thrust chemical rocket engine study

Engine data and information are presented to perform system studies on cargo orbit-transfer vehicles which would deliver large space structures to geosynchronous equatorial orbit. Low-thrust engine performance, weight, and envelope parametric data were established, preliminary design information was generated, and technologies for liquid rocket engines were identified. Two major engine design drivers were considered in the study: cooling and engine cycle options. Both film-cooled and regeneratively cooled engines were evaluated. The propellant combinations studied were hydrogen/oxygen, methane/oxygen, and kerosene/oxygen.

Mellish, J. A.↗
NASA NTRS1987

Performance characteristics of moving belt radiators

A design features and performance capabilities evaluation is made for a novel, 'hybrid belt radiator' spacecraft heat rejection system, which retains the excellent heat transfer capacity of liquid belt radiators but does not require the direct exposure of a free liquid surface to the space environment. Attention is given to the preliminary results of a computer model analyzing the dynamic behavior of the flexible belt structure due to spacecraft accelerations, as well as to the results of system studies determining size constraints on the radiator. Over the 300-700 K range of operating temperatures, Space Shuttle-stowable radiators with 10-200 MW thermal capacities can be designed.

Mcfadden, D.↗
NASA NTRS1990

Containerless processing of undercooled melts

All practical solidification processes involve some level of melt undercooling. Usually in bulk liquids crystallization is initiated at a heterogeneous nucleation site at low undercooling. The realization of appreciable levels of liquid undercooling requires some control over the kinetics of crystal nucleation. One level of control is available in containerless processing where the capability to melt and solidify samples without a container removes a major source of impurities and heterogeneous nucleation sites which can be effective in promoting large undercooling and in studying other potential nucleants. The microgravity environment offers a unique opportunity for containerless processing of large liquid samples with negligible melt disturbance and continuous temperature measurement throughout processing. Perhaps the most important potential for microgravity containerless processing lies in the significant structural control to develop distinct microstructures and metastable phases during solidification at high undercooling.

Perepezko, J. H.↗
NASA NTRS1991

Cryo-mechanical tests of Ames 24E2 IR-black coating

In addition to the ambient condition tape test, five mechanical tests of the IR-black coating, Ames 24E2, were performed at either liquid helium or liquid nitrogen temperatures. Tensile strain in the coating at liquid nitrogen temperature was measured up to values of 4E-3, both before and after the coating was cycled down to liquid helium temperature. When applied to an aluminum substrate which was then bent in liquid nitrogen, the aluminum substrate always failed (permanently deformed) well before the coating failed. Sinusoidal accelerations up to 45 Gs in liquid nitrogen and 25 Gs in liquid helium did not crack or otherwise visibly damage the coating. Both sinusoidal and random acceleration at about 90 K of a representative baffle vane structure, at frequencies from 10 to 2000 Hz and up to 15 Gs, did not damage the coating, even at the intersection of a baffle with the telescope tube. Thus on a macroscopic level, cryogenic cooling and various levels of acceleration and strain did not affect this coating. However, on a microscopic scale, some loose particles were found associated with several tests. Outgassing data are also given.

Smith, Sheldon M.↗
NASA NTRS1990

Cryo-mechanical tests of Ames 24E2 IR-black coating

In addition to the ambient condition tape test, five mechanical tests of the IR-black coating, Ames 24E2, were performed at either liquid helium or liquid nitrogen teperatures. Tensile strain in the coating at liquid nitrogen temperature was measured up to values of 4E-3, both before and after the coating was cycled down to liquid helium temperature. When applied to an aluminum substrate which was then bent in liquid nitrogen, the aluminum substrate always failed (permanently deformed) well before the coating failed. Sinusoidal accelerations up to 45 Gs in liquid nitrogen and 25 Gs in liquid helium did not crack or otherwise visibly damage the coating. Both sinusoidal and random acceleration at about 90 K of a representative baffle vane structure, at frequencies from 10 to 2000 Hz and up to 15 Gs, did not damage the coating, even at the intersection of a baffle with the telescope tube. Thus on a macroscopic level, cryogenic cooling and various levels of acceleration and strain did not affect this coating. However, on a microscopic scale, some loose particles were found associated with several tests. Outgassing data are also given

Smith, Sheldon M.↗
NASA NTRS1994

Thermocapillary flow with evaporation and condensation and its effect on liquid retention in low-G fluid acquisition devices

The steady motion, thermal and free surface behavior of a volatile, wetting liquid in microgravity are studied using scaling and numerical techniques. The objective is to determine whether the thermocapillary and two-phase convection arising from thermodynamic nonequilibrium along the porous surfaces of spacecraft liquid acquisition devices could cause the retention failures observed with liquid hydrogen and heated vapor pressurant. Why these devices seem immune to retention loss when pressurized with heated helium or heated directly through the porous structure was also examined. Results show that highly wetting fluids exhibit large negative and positive dynamic pressure gradients towards the meniscus interline when superheated and subcooled, respectively. With superheating, the pressure variation and recoil force arising from liquid/vapor phase change exert the same influence on surface morphology and promote retention. With subcooling, however, the pressure distribution produces a suction that degrades mechanical equilibrium of the surface. This result indicates that thermocapillary-induced deformation arising from subcooling and condensation is the likely cause for retention loss. In addition, increasing the level of nonequilibrium by reducing accommodation coefficient suppresses deformation and explains why this failure mode does not occur in instances of direct screen heating or pressurization with a heated inert gas.

Schmidt, George R.↗
NASA NTRS1999

Lockheed Martin Skunk Works Single Stage to Orbit/Reusable Launch Vehicle

Lockheed Martin Skunk Works has compiled an Annual Performance Report of the X-33/RLV Program. This report consists of individual reports from all industry team members, as well as NASA team centers. This portion of the report is comprised of a status report of Lockheed Martin's contribution to the program. The following is a summary of the Lockheed Martin Centers involved and work reviewed under their portion of the agreement: (1) Lockheed Martin Skunk Works - Vehicle Development, Operations Development, X-33 and RLV Systems Engineering, Manufacturing, Ground Operations, Reliability, Maintainability/Testability, Supportability, & Special Analysis Team, and X-33 Flight Assurance; (2) Lockheed Martin Technical Operations - Launch Support Systems, Ground Support Equipment, Flight Test Operations, and RLV Operations Development Support; (3) Lockheed Martin Space Operations - TAEM and A/L Guidance and Flight Control Design, Evaluation of Vehicle Configuration, TAEM and A/L Dispersion Analysis, Modeling and Simulations, Frequency Domain Analysis, Verification and Validation Activities, and Ancillary Support; (4) Lockheed Martin Astronautics-Denver - Systems Engineering, X-33 Development; (5) Sanders - A Lockheed Martin Company - Vehicle Health Management Subsystem Progress, GSS Progress; and (6) Lockheed Martin Michoud Space Systems - X-33 Liquid Oxygen (LOX) Tank, Key Challenges, Lessons Learned, X-33/RLV Composite Technology, Reusable Cyrogenic Insulation (RCI) and Vehicle Health Monitoring, Main Propulsion Systems (MPS), Structural Testing, X-33 System Integration and Analysis, and Cyrogenic Systems Operations.

Source record↗
NASA NTRS2001

Lightweight Chambers for Thrust Assemblies

The Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA) has successfully applied new materials and fabrication techniques to create actively cooled thrust chambers that operate 200-400 degrees hotter and weigh 50% lighter than conventional designs. In some vehicles, thrust assemblies account for as much as 20% of the engine weight. So, reducing the weight of these components and increasing their operating range will benefit many engines and vehicle designs, including Reusable Launch Vehicle (RLV) concepts. Obviously, copper and steel alloys have been used successfully for many years in the chamber components of thrust assemblies. Yet, by replacing the steel alloys with Polymer Matrix Composite (PMC) and/or Metal Matrix Composite (MMC) materials, design weights can be drastically reduced. In addition, replacing the traditional copper alloys with a Ceramic Matrix Composite (CMC) or an advanced copper alloy (Cu-8Cr-4Nb, also known as GRCop-84) significantly increases allowable operating temperatures. Several small MMC and PMC demonstration chambers have recently been fabricated with promising results. Each of these designs included GRCop-84 for the cooled chamber liner. These units successfully verified that designs over 50% lighter are feasible. New fabrication processes, including advanced casting technology and a low cost vacuum plasma spray (VPS) process, were also demonstrated with these units. Hot-fire testing at MSFC is currently being conducted on the chambers to verify increased operating temperatures available with the GRCop-84 liner. Unique CMC chamber liners were also successfully fabricated and prepared for hot-fire testing. Yet, early results indicate these CMC liners need significantly more development in order to use them in required chamber designs. Based on the successful efforts with the MMC and PMC concepts, two full size "lightweight" chambers are currently being designed and fabricated for hot-fire testing at MSFC in 2001. These "full size" chambers will be similar in size to those used on the X33 engine (RS2200). One will be fabricated with a MMC structural jacket, while the other uses a PMC jacket. Each will be designed for thrust levels of 15,000 pounds in an oxygen/hydrogen environment with liquid hydrogen coolant. Both chambers will use GRCop-84 for its channel wall liner. Each unit is expected to be at least 60% lighter than a conventional design with traditional materials. Hot-fire testing on the full size units in late 2001 will directly compare performance results between a conventional chamber design and these "lightweight" alternatives. The technology developed and demonstrated by this effort will not only benefit next generation RLV programs, but it can be applied to other existing and future engine programs, as well. Efforts were sponsored by the Advanced Space Transportation Program for RLV Focused Technologies. The task team was led by MSFC with additional members from NASA-Glenn Research Center and the Rocketdyne Division of The Boeing Company. Specific materials development and fabrication processes were provided by Aerojet, Lockheed Martin Astronautics, Composite Optics, Inc., Hyper-Therm, Ceramic Composites, Inc., MSE Technology Applications, and Plasma Processes, Inc.

Elam, Sandra K.↗
NASA NTRS2010

Phase Change Material Trade Study: A Comparison Between Wax and Water for Manned Spacecraft

Phase change material heat sinks have been recognized as an important tool in optimizing thermal control systems for space exploration vehicles and habitats that must deal with widely varying thermal loads and environments. In order to better focus technology investment in this arena, NASA has supported a trade study with the objective of identifying where the best potential pay-off can be found among identified aqueous and paraffin wax phase change materials and phase change material heat sink design approaches. The study used a representative exploration mission with well understood parameters to support the trade. Additional sensitivity studies were performed to ensure the applicability of study results across varying systems and destinations. Results from the study indicate that a water ice PCM heat sink has the potential to decrease the equivalent system mass of the mission s vehicle through a combination of a smaller heat sink and a slight 5% increase in radiator size or the addition of a lightweight heat pump. An evaluation of existing and emerging PCM heat sink technologies indicates that further significant mass savings should be achievable through continued development of those technologies. The largest mass savings may be realized by managing the location of the liquid and the solid in the heat sink to eliminate the melting and freezing pressure of wax and water, respectively, while also accommodating the high structural loads expected on future manned launch vehicles.

Quinn, Gregory↗
NASA NTRS1992

Project WISH: The Emerald City

Phase 3 of Project WISH saw the evolution of the Emerald City (E-City) from a collection of specialized independent analyses and ideas to a working structural design integrated with major support systems and analyses. Emphasis was placed on comparing and contrasting the closed and open cycle gas core nuclear rocket engines to further determine the optimum propulsive system for the E-City. Power and thermal control requirements were then defined and the question of how to meet these requirements was addressed. Software was developed to automate the mission/system/configuration analysis so changes dictated by various subsystem constraints could be managed efficiently and analyzed interactively. In addition, the liquid hydrogen propellant tank was statically designed for minimum mass and shape optimization using a finite element modeling package called SDRC I-DEAS. Spoke and shaft cross-sectional areas were optimized on ASTROS (Automated Structural Optimization System) for mass minimization. A structural dynamic analysis of the optimal structure also conducted using ASTROS enabled a study of the modes, frequencies, displacements, and accelerations of the E-City. Finally, the attitude control system design began with an initial mass moment of inertia analysis and was then designed and optimized using linear quadratic regulator control theory.

Oz, Hayrani↗