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Large Space Systems materials requirements are discussed in terms of types of materials, critical properties, and environmental stability. An outline is given of the materials technology development that will be needed to meet these requirements.
Tandem solar cells, where multiple single-junction cells are combined optically in series, provide a path to making cells with high areal efficiencies, with multiple material systems capable of achieving greater than 30% efficiency under 1-sun conditions. However, there are many different material combinations and configurations used to make a tandem, and it can be challenging to understand how advances in one material system will impact the performance of a tandem device. Here, we have built an open-source calculator based on the spectral efficiency metric proposed by Yu et al. to easily enable calculation of spectral efficiency for single junctions and predicted maximum efficiency of tandem pairs, accounting for different optical and electrical coupling between the top and bottom junctions.
A study was conducted to determine the effects of long term flight and ground exposure on three commercially available graphite-epoxy material systems: T300/5208, T300/5209, and T300/934. Sets of specimens were exposed on commercial aircraft and ground racks for 1, 2, 3, 5, and 10 years. Inflight specimen sites included both the interior and exterior of aircraft based in Hawaii, Texas, and New Zealand. Ground racks were located at NASA-Dryden and the above mentioned states. Similar specimens were exposed to controlled lab conditions for up to 2 years. After each exposure, specimens were tested for residual strength and a dryout procedure was used to measure moisture content. Both room and high temperature residual strengths were measured and expressed as a pct. of the unexposed strength. Lab exposures included the effects of time alone, moisture, time on moist specimens, weatherometer, and simulated ground-air-ground cycling. Residual strengths of the long term specimens were compared with residual strengths of the lab specimens. Strength retention depended on the exposure condition and the material system. Results showed that composite materials can be successfully used on commercial aircraft if environmental effects are considered.
Ceramic material systems are being considered for potential use as turbine blade tip gas path seals at temperatures up to 1370 1/4 C. Silicon carbide and silicon nitride structures were selected for study since an initial analysis of the problem gave these materials the greatest potential for development into a successful materials system. Segments of silicon nitride and silicon carbide materials over a range of densities, processed by various methods, a honeycomb structure of silicon nitride and ceramic blade tip inserts fabricated from both materials by hot pressing were tested singly and in combination. The evaluations included wear under simulated engine blade tip rub conditions, thermal stability, impact resistance, machinability, hot gas erosion and feasibility of fabrication into engine components. The silicon nitride honeycomb and low-density silicon carbide using a selected grain size distribution gave the most promising results as rub-tolerant shroud liners. Ceramic blade tip inserts made from hot-pressed silicon nitride gave excellent test results. Their behavior closely simulated metal tips. Wear was similar to that of metals but reduced by a factor of six.
Results are presented of a program to determine the impact resistance of composite fan blades subjected to foreign object damage (FOD) while operating under conditions simulating a short take-off and landing (STOL) engine at takeoff. The full-scale TF39 first-stage fan blade was chosen as the base design for the demonstration component since its configuration and operating tip speeds are similar to a typical STOL fan blade several composite configurations had already been designed and evaluated under previous programs. The first portion of the program was devoted toward fabricating and testing high impact resistant, aerodynamically acceptable composite blades which utilized only a single material system in any given blade. In order to increase the blade impact capability beyond this point, several mixed material (hybrid) designs were investigated using S-glass and Kevlar as well as boron and graphite fibers. These hybrid composite blades showed a marked improvement in resistance to bird impact over those blades made of a single composite material. The work conducted under this program has demonstrated substantial improvement in composite fan blades with respect to FOD resistance and has indicated that the hybrid design concept, which utilizes different types of fibers in various portions of a fan blade design depending on the particular requirements of the different areas and the characteristics of the different fibers involved, shows a significant improvement over those designs utilizing only one material system.
NASA Dryden is directing a program to test a Carbon-Silicon Carbide (C/SiC) Ruddervator Subcomponent Test Article (RSTA). The RSTA is a truncated version of the full-scale X-37 hot-structure control surface incorporating all major features including the metallic spindle and five major C/SiC quasi isotropic lay-up components secured with C/SiC fasteners. As part of NASA's Aeronautics Research Mission Directorate Hypersonics program, the RSTA will undergo thermal-structural testing to develop an extensive database for future structural design and analysis methodology validation. Ground Vibration Tests (GVTs) are routinely conducted for model validation in supporting flutter analysis for subsonic and supersonic vehicles; however, for hypersonic vehicle applications, GVT techniques are not well-established. New fabrication technologies, high-temperature materials systems, and sensors offer new opportunities to develop techniques for performing GVTs at elevated temperatures. The RSTA is comprised of materials whose stiffness both increases and decreases with increasing temperature. The impact of this type of material system interaction must be understood as it will ultimately affect hypersonic flutter analysis. The test objectives are to perform room-temperature GVTs, develop the capability to conduct high-temperature GVTs and to compare and generate an understanding of the modal characteristics which capture RSTA's material interaction when subjected to temperature varying conditions.
Thermal Protection Systems (TPS) are critical for enabling NASA missions involving high-speed atmospheric flight where the entries usually include descending into the atmosphere followed by a trajectory that aims to burn off energy and result in a controlled landing. NASA Ames focuses on qualifying and certifying TPS for current missions, sustaining TPS for future missions, and developing new TPS for upcoming missions where a heritage solution is not viable. More recently there is also a focus on advancing and transferring technologies that can benefit both commercial and government space needs. Developing mature thermal protection systems is a lengthy process involving advanced tools, extensive research, and testing. Design and analysis tools are used to predict aerothermal environments, aid the design of test and flight hardware, and support the testing for the thermal/mechanical response of thermal protection systems. More recently, advances in computational methods help reduce the time and cost of technological advances, aid in optimized material architecture design, and improve material properties and performance. While high-enthalpy testing that simulates the conditions of space flight remains essential for the evaluation and development of TPS materials, computational tools are already showing promise in reducing the need for widespread testing and can help fast-track the design cycle. With the exploration of new destinations EDL instrumentation remains an important element of the heatshield and NASA Ames and partners have developed and delivered instrumentation flight hardware in support of recent Mars missions (MSL and M2020) as well as Artemis Orion. Sensors installed on the heatshield and backshell of spacecraft provide coveted information about the aerodynamic and aerothermal environment during entry. Over the years NASA Ames has brought several reusable and ablative TPS materials to a level of readiness to hand off to missions and the branch continues to serve as a TPS steward for the agency. This presentation will cover current developments in the above areas that have enabled recent missions, and look to future TPS needs for missions such as Ice Giants.
Thermal Protection Systems (TPS) are critical for enabling NASA missions involving high-speed atmospheric flight where the entries usually include descending into the atmosphere followed by a trajectory that aims to burn off energy and result in a controlled landing. NASA Ames focuses on qualifying and certifying TPS for current missions, sustaining TPS for future missions, and developing new TPS for upcoming missions where a heritage solution is not viable. More recently there is also a focus on advancing and transferring technologies that can benefit both commercial and government space needs.Developing mature thermal protection systems is a lengthy process involving advanced tools, extensive research, and testing. Design and analysis tools are used to predict aerothermal environments, aid the design of test and flight hardware, and support the testing for the thermal/mechanical response of thermal protection systems. More recently, advances in computational methods help reduce the time and cost of technological advances, aid in optimized material architecture design, and improve material properties and performance. While high-enthalpy testing that simulates the conditions of space flight remains essential for the evaluation and development of TPS materials, computational tools are already showing promise in reducing the need for widespread testing and can help fast-trackthe design cycle.With the exploration of new destinations EDL instrumentation remains an important element of the heatshield and NASA Ames and partners have developed and delivered instrumentation flight hardware in support of recent Mars missions,including Mars Science Lab (MSL) and Mars 2020 (M2020),as well as Artemis Orion. Sensors installed on the heatshield and backshell of spacecraft provide coveted information about the aerodynamic and aerothermal environment during entry.Overthe years NASA Ames has brought several reusable and ablative TPS materials to a level of readiness to hand off to missions and the Thermal Protection Materials branch continues to serve as a TPS steward for the agency.
Hot structures are multifunctional material systems that integrate the functions of thermal protection systems (TPS) and primary structure to increase mass efficiency in aerospace vehicles. The goal of this project was to improve our capabilities for developing lightweight, reusable, and reliable hot structures concepts to enable advanced space exploration missions. Specifically, current hot structure carbon/carbon (C/C) composite material systems have low interlaminar strength, which can lead to material failures at low load levels, and large material property variability in as-fabricated parts. Two parallel paths were followed to address these shortcomings. First, analysis models were developed to identify the conditions under which excessive interlaminar stresses occur and to evaluate the potential of stitching the C/C laminae to improve the interlaminar thermal-structural response. This effort included the development of high-fidelity, fracture-based thermal-structural finite element analysis models to evaluate critical features needed to produce efficient hot structure concepts. The second path focused on developing atomistic and mesoscale simulation tools to model the fabrication process in an effort to reduce the time and cost of production while improving the quality of the resulting parts. Aspects of the process currently being investigated include the infusion of the matrix precursor into the fiber preform, the cure of the precursor molecules into their polymeric form, and the carbonization and graphitization of the polymer to form the carbon matrix material. Results of the 6-month study are presented herein.
The internal coupling of magnetic excitations (magnons) with themselves has created a new research sub-field in hybrid magnonics, i.e., magnon-magnon coupling, which focuses on materials discovery and engineering for probing and controlling magnons in a coherent manner. This is enabled by, one, the abundant mechanisms of introducing magnetic interactions, with examples of exchange coupling, dipolar coupling, Ruderman–Kittel–Kasuya–Yosida (RKKY) coupling, and Dzyaloshinskii–Moriya interaction (DMI) coupling, and two, the vast knowledge of how to control magnon band structure, including field and wavelength dependences of frequencies, for determining the degeneracy of magnon modes with different symmetries. In particular, we discuss how magnon-magnon coupling is implemented in various materials systems, with examples of magnetic bilayers, synthetic antiferromagnets, nanomagnetic arrays, layered van der Waals magnets, and (DMI spin-orbit torque materials) in magnetic multilayers. Here, we then introduce new concept of applications for these hybrid magnonic materials systems, with examples of frequency up/down conversion and magnon-exciton coupling, and discuss what properties are desired for achieving those applications.
- The Mars Sample Return Earth Entry System is being designed to withstand entry environments of: - Peak Heat Flux of > 3000 W/cm2 - Peak Dynamic Pressure of > 2 atmospheres - Peak Shear of > 3000 Pa. - The TPS Material System Selected to support MSR EES is “3MDCP”, 3D Mid Density Carbon Phenolic. 3MDCP is a derivative of the HEEET material system, using the “Insulation Layer” only. - Implementation of 3MDCP on EES is monolithic instead of a tile architecture in order to meet MSR reliability demands. - The EES TPS Team has recently completed a number of major engineering development milestones in support of delivering the flight 3MDCP TPS hardware: - Weaving Scale-Up completed, > 80” width woven preform production underway - Full Scale Forming & Manufacturing Attrition completed to verify delivered thickness capabilities. - Recent Design Change: EES design has been changed from a 45 degree sphere cone to a 52.5 degree sphere cone to allow for additional system mass.
Hydrogen is pivotal in the transition to sustainable energy systems, playing major roles in power generation and industrial applications. Metal–organic frameworks (MOFs) have emerged as promising mediums for efficient hydrogen storage. However, identifying potential candidates for deployment is challenging due to the vast number of currently available synthesized MOFs. This study integrates molecular simulations, machine learning, and techno-economic analysis to evaluate the performance of MOFs across broad operation conditions for hydrogen storage applications. While previous screenings of MOF databases have predominantly emphasized high hydrogen capacities under cryogenic conditions, this study reveals that optimal temperatures and pressures for cost minimization depend on the raw price of the MOF. Specifically, when MOFs are priced at $15/kg, among the 9720 MOFs tested, 9692 MOFs achieve the lowest cost at temperatures between 170 K and 250 K and a pressure of 150 bar. Under these optimal conditions, 362 MOFs deliver a lower levelized cost of storage than 350 bar compressed gas hydrogen storage. Furthermore, this study reveals key material properties that result in low system cost, such as high surface areas (>3000 m2/g), large void fractions (>0.78), and large pore volumes (>1.1 cm3/g).
A new research program was initiated as a preliminary phase. The following three objectives are being pursued for the overall program: to elucidate the mechanisms of microcracking for graphite fiber-reinforced semi-IPN polyimide matrix composites under mechanical and thermal cyclic loading; to devise material engineering solutions for possible improvement of fatigue damage resistance (or the increase of fatigue endurance strength) of semi-IPN matrix composites by tailoring of modulus and toughness of fiber-resin interface region; and to assess processing characteristics of the composites and their roles in controlling the resistance of composites to microcracking and the effectiveness of interface toughening. The main emphasis was placed upon the initial screening of material systems and optimization of processing conditions for semi-IPN matrix composites with tailored interface. As a first set of control material systems to study, the composites were prepared with unsized Celion 6000 graphite fiber reinforcement and the following resin matrices of varied fracture toughness: PMR-15 thermoset polyimide, semi-IPN of PNR-15 thermoset polyimide and NR150B2 thermoplastic polyimide in 75/25 ratio, and semi-IPN of PMR-15 and NR150B2 in 50/50 ratio. For the composites with the resin matrix of semi-IPN in 75/25 ratio, interface tailoring was attempted by using graphite fibers coated with the resins of systematically varied fracture toughness. In the continuing work, a broad range of interlayer toughness will be achieved by coating the fibers with reactants of semi-IPN having lower or higher content of thermoplastic constituent in comparison with the composition of surrounding resin matrix. In pursuing the objectives of the overall research program, the respective roles and interaction of critical parameters were defined.
Enhanced, lightweight material systems, such as 17-4PH stainless steel sandwich panels are being developed for use as fan blades and fan containment material systems for next generation engines. In order to improve the production for these systems, nondestructive evaluation (NDE) techniques, such as ultrasonic spectroscopy, are being utilized to evaluate the brazing quality between the 17-4PH stainless steel face plates and the 17-4PH stainless steel foam core. Based on NDE data, shear tests are performed on sections representing various levels of brazing quality from an initial batch of these sandwich structures. Metallographic characterization of brazing is done to corroborate NDE findings and the observed shear failure mechanisms.
Ground vibration tests or modal surveys are routinely conducted for supporting flutter analysis for subsonic and supersonic vehicles; however, for hypersonic vehicle applications, thermoelastic vibration testing techniques are not well established and are not routinely performed for supporting hypersonic flutter analysis. New high-temperature material systems, fabrication technologies and high-temperature sensors expand the opportunities to develop advanced techniques for performing ground vibration tests at elevated temperatures. High-temperature materials have the unique property of increasing in stiffness when heated. When these materials are incorporated into a hot-structure, which includes metallic components that decrease in stiffness with increasing temperature, the interaction between the two materials systems needs to be understood because that interaction could ultimately affect the hypersonic flutter analysis. Performing a high-temperature modal survey will expand the research database for hypersonics and will help build upon the understanding of the dual material interaction. This paper will discuss the vibration testing of the Carbon-Silicon Carbide Ruddervator Subcomponent Test Article which is a truncated version of the full-scale X-37 hot-structure control surface. In order to define the modal characteristics of the test article during the elevated-temperature modal survey, two series of room-temperature modal test configurations had to be performed. The room-temperature test series included one with the test article suspended from a bungee cord (free-free) and the second with it mounted on the strongback (fixed boundary condition) in NASA Dryden's Flight Loads Lab large nitrogen test chamber.
An object within a chamber is acoustically moved by applying wavelengths of different modes to the chamber to move the object between pressure wells formed by the modes. In one system, the object is placed in one end of the chamber while a resonant mode, applied along the length of the chamber, produces a pressure well at the location. The frequency is then switched to a second mode that produces a pressure well at the center of the chamber, to draw the object. When the object reaches the second pressure well and is still traveling towards the second end of the chamber, the acoustic frequency is again shifted to a third mode (which may equal the first model) that has a pressure well in the second end portion of the chamber, to draw the object. A heat source may be located near the second end of the chamber to heat the sample, and after the sample is heated it can be cooled by moving it in a corresponding manner back to the first end of the chamber. The transducers for levitating and moving the object may be all located at the cool first end of the chamber.
Four current viscoplastic models are compared experimentally for Inconel 718 at 593 C. This material system responds with apparent negative strain rate sensitivity, undergoes cyclic work softening, and is susceptible to low cycle fatigue. A series of tests were performed to create a data base from which to evaluate material constants. A method to evaluate the constants is developed which draws on common assumptions for this type of material, recent advances by other researchers, and iterative techniques. A complex history test, not used in calculating the constants, is then used to compare the predictive capabilities of the models. The combination of exponentially based inelastic strain rate equations and dynamic recovery is shown to model this material system with the greatest success. The method of constant calculation developed was successfully applied to the complex material response encountered. Backstress measuring tests were found to be invaluable and to warrant further development.
The Solar Array Module Plasma Interactions Experiment (SAMPIE) is a flight experiment that flew on the Space Shuttle Columbia (STS-62) in March 1994, as part of the OAST-2 mission. The overall objective of SAMPIE was to determine the adverse environmental interactions within the space plasma of low earth orbit (LEO) on modern solar cells and space power system materials which are artificially biased to high positive and negative direct current (DC) voltages. The two environmental interactions of interest included high voltage arcing from the samples to the space plasma and parasitic current losses. High voltage arcing can cause physical damage to power system materials and shorten expected hardware life. parasitic current losses can reduce power system efficiency because electric currents generated in a power system drain into the surrounding plasma via parasitic resistance. The flight electronics included two programmable high voltage DC power supplies to bias the experiment samples, instruments to measure the surrounding plasma environment in the STS cargo bay, and the on-board data acquisition system (DAS). The DAS provided in-flight experiment control, data storage, and communications through the Goddard Space Flight Center (GSFC) Hitchhiker flight avionics to the GSFC Payload Operations Control Center (POCC). The DAS and the SAMPIE POCC computer systems were designed for telescience operations; this paper will focus on the experiences of the SAMPIE team regarding telescience development and operations from the GSFC POCC during STS-62. The SAMPIE conceptual development, hardware design, and system verification testing were accomplished at the NASA Lewis Research Center (LeRC). SAMPIE was developed under the In-Space Technology Experiment Program (IN-STEP), which sponsors NASA, industry, and university flight experiments designed to enable and enhance space flight technology. The IN-STEP Program is sponsored by the Office of Space Access and Technology (OSAT).