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NASA NTRS1989

Verification of spatial and temporal pressure distributions in segmented solid rocket motors

A wide variety of analytical tools are in use today to predict the history and spatial distributions of pressure in the combustion chambers of solid rocket motors (SRMs). Experimental and analytical methods are presented here that allow the verification of many of these predictions. These methods are applied to the redesigned space shuttle booster (RSRM). Girth strain-gage data is compared to the predictions of various one-dimensional quasisteady analyses in order to verify the axial drop in motor static pressure during ignition transients as well as quasisteady motor operation. The results of previous modeling of radial flows in the bore, slots, and around grain overhangs are supported by approximate analytical and empirical techniques presented here. The predictions of circumferential flows induced by inhibitor asymmetries, nozzle vectoring, and propellant slump are compared to each other and to subscale cold air and water tunnel measurements to ascertain their validity.

Salita, Mark↗

NASA NTRS1992

Acoustic properties of supersonic helium/air jets at low Reynolds numbers

Experiments have been performed with the objective of developing a greater understanding of the physics of hot supersonic jet noise. Cold helium/air jets are used to easily and inexpensively simulate the low densities of hot air jets. The experiments are conducted at low Reynolds numbers in order to facilitate study of the large-scale turbulent structures (instability waves) that cause most of the radiated noise. Experiments have been performed on Mach 1.5 and 2.1 jets of pure air, pure helium and 10 percent helium by mass. Helium/air jets are shown to radiate more noise than pure air jets due to the increased exit velocity. Microphone spectra are usually dominated by a single spectral component at a predictable frequency. Increasing the jet's helium concentration is shown to increase the dominant frequency. The helium concentration in the test chamber is determined by calculating the speed of sound from the measured phase difference between two microphone signals. Bleeding outside air into the test chamber controls the accumulation of helium so that the hot jet simulation remains valid. The measured variation in the peak radiated noise frequency is in good agreement with the predictions of the hot jet noise theory of Tam et al.

Mclaughlin, Dennis K.↗

NASA NTRS2015

Design and Test of a Liquid Oxygen / Liquid Methane Thruster with Cold Helium Pressurization Heat Exchanger

A liquid oxygen / liquid methane 2,000 lbf thruster was designed and tested in conjuction with a nozzle heat exchanger for cold helium pressurization. Cold helium pressurization systems offer significant spacecraft vehicle dry mass savings since the pressurant tank size can be reduced as the pressurant density is increased. A heat exchanger can be incorporated into the main engine design to provide expansion of the pressurant supply to the propellant tanks. In order to study the systems integration of a cold-helium pressurization system, a 2,000 lbf thruster with a nozzle heat exchanger was designed for integration into the Project Morpheus vehicle at NASA Johnson Space Center. The testing goals were to demonstrate helium loading and initial conditioning to low temperatures, high-pressure/low temperature storage, expansion through the main engine heat exchanger, and propellant tank injection/pressurization. The helium pressurant tank was an existing 19 inch diameter composite-overwrap tank, and the targert conditions were 4500 psi and -250 F, providing a 2:1 density advantage compared to room tempatrue storage. The thruster design uses like-on-like doublets in the injector pattern largely based on Project Morpheus main engine hertiage data, and the combustion chamber was designed for an ablative chamber. The heat exchanger was installed at the ablative nozzle exit plane. Stand-alone engine testing was conducted at NASA Stennis Space Center, including copper heat-sink chambers and highly-instrumented spoolpieces in order to study engine performance, stability, and wall heat flux. A one-dimensional thermal model of the integrated system was completed. System integration into the Project Morpheus vehicle is complete, and systems demonstrations will follow.

Melcher, John C.↗

NASA NTRS1973

Noise characteristics of combustion augmented high speed jets.

Noise generation by a subsonic flow discharging from a combustion chamber is examined with regard to the relative importance of combustion as a source of noise in such a flow system. Measurements of pressure fluctuations inside the combustor are compared with far-field noise measurements by direct cross-correlations. The cross-correlations and derived cross-spectral densities verify that much of the noise originates inside the combustor. A first-order fluid-mechanical perturbation model is used to predict exit plane velocity fluctuations due to internal pressure fluctuations. Far-field noise levels estimated on this basis are in good agreement with measured values. The overall noise level from the combustor/jet is found to be 10 to 20 dB higher than for an equivalent clean, cold jet at the same exit velocity.

Abdelhamid, A. N.↗

NASA NTRS1994

Thermal balance testing of MSAT 2 spacecraft

The present work reports on the recently completed infrared thermal balance/thermal vacuum testing of a MSAT satellite, the first satellite to provide mobile communications service for all of continental North America. MSAT is a two spacecraft program, using a three-axis stabilized Hughes HS-601 series Bus as the vehicle for the Canadian designed Payload. The thermal tests which were performed at the Canadian Space Agency's David Florida Laboratory in Ottawa, Canada, lasted approximately 35 days. The infrared (IR) heating rig was designed to provide radiant heat inputs into seven spacecraft zones during Thermal Vacuum (TV) testing. The TV test was divided into multiple phases. It began with a thermal balance cold phase, followed by a thermal cold cycle and a hot balance phase, complemented by a thermal hot cycle to finish with a thermal cycle with continuous monitoring of the Bus and Payload. The spacecraft's external heat fluxes were provided by IR lamp sources. To ensure flux uniformity, highly reflective baffles and IR East and West faces; the Earth facing (Nadir); and the inside of the thrust cylinder. The aft-end panel heat fluxes were provided by a heated LN2 shroud. The radiation flux intensity on the spacecraft zones from the various rig elements was measured using Monitored Background Radiometers (MBR's) and compared with direct calculations and with pretest predictions. The temperature measurement system was based on Uniform Temperature References (UTR's) located inside the chamber such that all feedthroughs were copper-copper. This system was devised to achieve a temperature measurement accuracy of plus/minus 0.5 C for over 850 thermocouples used in the test. A PC-(QNX-based) based real-time data acquisition system was utilized to provide continuous monitoring of all channels based on a 30-second time scan. In addition, the data acquisition system was able to retrieve telemetry stream from the Satellite Test Equipments (STE) station for real-time data manipulation. Preliminary results showed the test to be successful from both the thermal balance side and the electrical testing side.

Samson, Serge↗

NASA NTRS1961

Interaction of Highly Underexpanded Jets with Simulated Lunar Surfaces

Pressure distributions and erosion patterns on simulated lunar surfaces (hard and soft) and interference effects between the surface and two representative lunar vehicles (cylindrical and spherical) were obtained with cold-air jets at various descent heights and nozzle total-pressure ratios up to 288,000. Surface pressure distributions were dependent on both nozzle area ratio and, nozzle contour. Peak pressures obtained with a sonic nozzle agreed closely with those predicted theoretically for a near-sonic jet expanding into a vacuum. Short bell-shaped nozzles gave annular pressure distributions; the low center pressure resulted from the coalescence of shocks that originated within the nozzle. The high surface pressures were contained within a circle whose diameter was about 16 throat diameters, regardless of nozzle area ratio or contour. The peak pressure increased rapidly as the vehicle approached the surface; for example, at a descent height of 40 throat diameters the peak pressure was 0.4 percent of the chamber pressure, but increased to 6 percent at 13 throat diameters. The exhaust jet eroded a circular concave hole in white sand at descent heights from about 200 to 600 throat diameters. The hole diameter was about 225 throat diameters, while the depth was approximately 60 throat diameters. The sand particles, which formed a conical sheet at a semivertex angle of 50 deg, appeared to follow a ballistic trajectory and at no time struck the vehicle. An increase in pressure was measured on the base of the cylindrical lunar vehicle when it approached to within 14 throat diameters of the hard, flat surface. No interference effects were noted between the spherical model and the surface to descent heights as low as 8 throat diameters.

Stitt, Leonard E.↗

NASA NTRSDate not supplied

Updating the Thermal Vacuum Chambers at the NASA Johnson Space Center

Chambers A and B are two large thermal vacuum chambers at Johnson Space Center which enable space simulation for unmanned and human-rated missions, respectively. With the resurgence in deep space missions for scientific research and various private commercial ventures, these chambers are expected to be used frequently for at least the next decade. For qualifying the James Webb Space Telescope, upgrades to Chamber A were performed which included the addition of a 12.5 kW refrigeration system with helium shrouds capable of simulating deep space environment and an efficient and reliable LN2 natural flow thermosiphon system for the thermal shield. Continuous improvements since then have focused on ensuring operational readiness by modernizing the data acquisition, recording, controls, and visualization systems for both chambers and clean room. These upgrades will be the focus of this paper. Controlling the Cryochambers was enhanced by moving from a 32-bit SCADA system to a 64-bit architected system. Infrastructural changes involved installing redundant power circuits, adding new servers, network switches, and including load balancing with fail-over between servers to minimize downtime. Instead of distributed servers, 3 redundant servers are used to share configurations. Configurations are now kept in a shared SQL instance making it easy to deploy and maintain. During this upgrade process many sub-systems (PLCs and NI PXI interfaces to sensors) were upgraded from the prior OPC-DA to the more secure OPC-UA protocol. Cryo-system PLCs were updated to allow Ganni cycle floating pressure calculations from any cold box to be sent to any compressor. During the project, the team recreated over 70,000 live data points, 50000 historical data points, and 4000 alarms. Finally, the graphical interfaces were upgraded to support HTML 5 in conjunction shared pages were implemented reducing the total number of webpages by over 75%. These system updates reinvigorated the previous SCADA system which had reached its end of life. The same look and feel was maintained while providing operators with an updated interface to control, troubleshoot, and record. The new system architecture is more robust and easier to maintain creating a path forward to address remaining problem points and implement additional features.

Cody Schaefer↗

NASA NTRSDate not supplied

Updating the Thermal Vacuum Chambers at the NASA Johnson Space Center

Cody Schaefer↗

NASA NTRS2007

Loop Heat Pipe Operation with Thermoelectric Converters and Coupling Blocks

This paper presents theoretical and experimental studies on using thermoelectric converters (TECs) and coupling blocks to control the operating temperature of a miniature loop heat pipes (MLHP). The MLHP has two parallel evaporators and two parallel condensers, and each evaporator has its own integral compensation chamber (CC). A TEC is attached to each CC, and connected to the evaporator via a copper thermal strap. The TEC can provide both heating and cooling to the CC, therefore extending the LHP operating temperature over a larger range of the evaporator heat load. A bi-polar power supply is used for the TEC operation. The bipolar power supply automatically changes the direction of the current to the TEC, depending on whether the CC requires heating or cooling, to maintain the CC temperature at the desired set point. The TEC can also enhance the startup success by maintaining a constant CC temperature during the start-up transient. Several aluminum coupling blocks are installed between the vapor line and liquid line. The coupling blocks serve as a heat exchanger which preheats the cold returning liquid so as to reduce the amount of liquid subcooling, and hence the power required to maintain the CC at the desired set point temperature. This paper focuses on the savings of the CC control heater power afforded by the TECs when compared to traditional electric heaters. Tests were conducted by varying the evaporator power, the condenser sink temperature, the CC set point temperature, the number of coupling blocks, and the thermal conductance of the thermal strap. Test results show that the TECs are able to control the CC temperature within k0.5K under all test conditions, and the required TEC heater power is only a fraction of the required electric heater power.

Ku, Jentung↗

NASA NTRS2013

Environmental Qualification of a Single-Crystal Silicon Mirror for Spaceflight Use

This innovation is the environmental qualification of a single-crystal silicon mirror for spaceflight use. The single-crystal silicon mirror technology is a previous innovation, but until now, a mirror of this type has not been qualified for spaceflight use. The qualification steps included mounting, gravity change measurements, vibration testing, vibration- induced change measurements, thermal cycling, and testing at the cold operational temperature of 225 K. Typical mirrors used for cold applications for spaceflight instruments include aluminum, beryllium, glasses, and glass-like ceramics. These materials show less than ideal behavior after cooldown. Single-crystal silicon has been demonstrated to have the smallest change due to temperature change, but has not been spaceflight-qualified for use. The advantage of using a silicon substrate is with temperature stability, since it is formed from a stress-free single crystal. This has been shown in previous testing. Mounting and environmental qualification have not been shown until this testing.

Hagopian, John↗

NASA NTRS1993

Experimental Determination of in Situ Utilization of Lunar Regolith for Thermal Energy Storage

A Lunar Thermal Energy from Regolith (LUTHER) experiment has been designed and fabricated at the NASA Lewis Research Center to determine the feasibility of using lunar soil as thermal energy storage media. The experimental apparatus includes an alumina ceramic canister (25.4 cm diameter by 45.7 cm length) which contains simulated lunar regolith, a heater (either radiative or conductive), 9 heat shields, a heat transfer cold jacket, and 19 type B platinum rhodium thermocouples. The simulated lunar regolith is a basalt, mined and processed by the University of Minnesota, that closely resembles the lunar basalt returned to earth by the Apollo missions. The experiment will test the effects of vacuum, particle size, and density on the thermophysical properties of the regolith. The properties include melt temperature (range), specific heat, thermal conductivity, and latent heat of storage. Two separate tests, using two different heaters, will be performed to study the effect of heating the system using radiative and conductive heat transfer. The physical characteristics of the melt pattern, material compatibility of the molten regolith, and the volatile gas emission will be investigated by heating a portion of the lunar regolith to its melting temperature (1435 K) in a 10(exp -4) pascal vacuum chamber, equipped with a gas spectrum analyzer. A finite differencing SINDA model was developed at NASA Lewis Research Center to predict the performance of the LUTHER experiment. The analytical results of the code will be compared with the experimental data generated by the LUTHER experiment. The code will predict the effects of vacuum, particle size, and density has on the heat transfer to the simulated regolith.

Richter, Scott W.↗

NASA NTRS1993

CFD Modeling of Turbulent Duct Flows for Coolant Channel Analysis

The design of modern liquid rocket engines requires the analysis of chamber coolant channels to maximize the heat transfer while minimizing the coolant flow. Coolant channels often do not remain at a constant cross section or at uniform curvature. New designs require higher aspect ratio coolant channels than previously used. To broaden the analysis capability and to complement standard analysis tools an investigation on the accuracy of CFD predictions for coolant channel flow has been initiated. Validation of CFD capabilities for coolant channel analysis will enhance the capabilities for optimizing design parameters without resorting to extensive experimental testing. The eventual goal is to use CFD to determine the flow fields of unique coolant channel designs and therefore determine critical heat transfer coefficients. In this presentation the accuracy of a particular CFD code is evaluated for turbulent flows. The first part of the presentation is a comparison of numerical results to existing cold flow data for square curved ducts (NASA CR-3367, 'Measurements of Laminar and Turbulent Flow in a Curved Duct with Thin Inlet Boundary Layers'). The results of this comparison show good agreement with the relatively coarse experimental data. The second part of the presentation compares two cases of higher aspect ratio channels (AR=2.5,10) to show changes in axial and secondary flow strength. These cases match experimental work presently in progress and will be used for future validation. The comparison shows increased secondary flow strength of the higher aspect ratio case due to the change in radius of curvature. The presentation includes a test case with a heated wall to demonstrate the program's capability. The presentation concludes with an outline of the procedure used to validate the CFD code for future design analysis.

Ungewitter, Ronald J.↗

NASA NTRS1999

A Design Tool for Liquid Rocket Engine Injectors

A practical design tool for the analysis of flowfields near the injector face has been developed and used to analyze the Fastrac engine. The objective was to produce a computational design tool which was detailed enough to predict the interactive effects of injector element impingement angles and points and the momenta of the individual orifice flows. To obtain a model which could be used to simulate a significant number of individual orifices, a homogeneous computational fluid dynamics model was developed. To describe liquid and vapor sub- and super-critical flows, the model included thermal and caloric equations of state which were valid over a wide range of pressures and temperatures. A homogeneous model was constructed such that the local state of the flow was determined directly, i.e. the quality of the flow was calculated. Such a model does not identify drops or their distribution, but it does allow the flow along the injector face and into the acoustic cavity to be predicted. It also allows the film coolant flow to be accurately described. The initial evaluation of the injector code was made by simulating cold flow from an unlike injector element and from a like-on-like overlapping fan (LOL) injector element. The predicted mass flux distributions of these injector elements compared well to cold flow test results. These are the same cold flow tests which serve as the data base for the JANNAF performance prediction codes. The flux distributions 1 inch downstream of the injector face are very similar; the differences were somewhat larger at further distances from the faceplate. Since the cold flow testing did not achieve good mass balances when integrations across the entire fan were made, the CFD simulation appears to be reasonable alternative to future cold flow testing. To simulate the Fastrac, an RP-1/LOX combustion model must be chosen. This submodel must be relatively simple to accomplish three-dimensional, multiphase flow simulations. Single RP-1 pyrolysis and partial oxidation steps were chosen and the combustion was completed with the wet CO mechanism. Soot was also formed with a single global reaction. To validate the combustion submodel, global data from gas generator tests and from subscale motor test were used to predict qualitatively correct mean molecular weights, temperature, and soot levels. Because such tests do not provide general kinetics rates, the methodology is not necessarily appropriate for other than rocket type flows conditions. Soot predictions were made so that radiation heating to the motor walls can be made. These initial studies of the Fastrac were for a small region close to the injector face and chamber wall which included a segment of the acoustic cavity. The region analyzed includes 11 individual orifice holes to represent the LOL elements and the H2 film coolant holes. Typical results of this simulation are shown in Figure 1. At this point the only available test data to verify the predictions are temperatures measured in the acoustic cavity. These temperatures are in reasonable agreement at about 2000R (1111 K). Future work is expected to include improving the computational efficiency or the CFD model and/or using more computer capacity than the single Pentium PC with which these simulations were made.

Farmer, Richard C.↗

NASA NTRS2002

NASA Tech Briefs, April 2002

The contents include: 1) Application Briefs; 2) Sneak Preview of Sensors Expo; 3) The Complexity of the Diagnosis Problem; 4) Design Concepts for the ISS TransHab Module; 5) Characteristics of Supercritical Transitional Mixing Layers; 6) Electrometer for Triboelectric Evaluation of Materials; 7) Infrared CO2 Sensor With Built-In Calibration Chambers; 8) Solid-State Potentiometric CO Sensor; 9) Planetary Rover Absolute Heading Detection Using a Sun Sensor; 10) Concept for Utilizing Full Areas of STJ Photodetector Arrays; 11) Development of Cognitive Sensors; 12) Enabling Higher-Voltage Operation of SOl CMOS Transistors; 13) Estimating Antenna-Pointing Errors From Beam Squints; 14) Advanced-Fatigue-Crack-Growth and Fracture- Mechanics Program; 15) Software for Sequencing Spacecraft Actions; 16) Program Distributes and Tracks Organizational Memoranda; 16) Flat Membrane Device for Dehumidification of Air; 17) Inverted Hindle Mount Reduces Sag of a Large, Precise Mirror; 18) Heart-Pump-Outlet/Cannula Coupling; 19) Externally Triggered Microcapsules Release Drugs In Situ; 20) Combinatorial Drug Design Augmented by Information Theory; 21) Multiple-Path-Length Optical Absorbance Cell; 22) Model of a Fluidized Bed Containing a Mixture of Particles; 23) Refractive Secondary Concentrators for Solar Thermal Systems; 24) Cold Flow Calorimeter; 25) Methodology for Tracking Hazards and Predicting Failures; 26) Estimating Heterodyne-Interferometer Polarization Leakage; 27) An Efficient Algorithm for Propagation of Temporal- Constraint Networks; 28) Software for Continuous Replanning During Execution; 29) Surface-Launched Explorers for Reconnaissance/Scouting; 30) Firmware for a Small Motion-Control Processor; 31) Gear Bearings and Gear-Bearing Transmissions; and 32) Linear Dynamometer With Variable Stroke and Frequency.

Source record↗

NASA NTRS2009

Assessment of SOI AND Gate, Type CHT-7408, for Operation in Extreme Temperature Environments

Electronic parts based on silicon-on-insulator (SOI) technology are finding widespread applications due to their ability to operate in harsh environments and the benefits they offer as compared to their silicon counterparts. Due to their construction, they are tailored for high temperature operation and show good tolerance to radiation events. In addition, their inherent design lessens the formation of parasitic junctions, thereby reducing leakage currents, decreasing power consumption, and enhancing speed. These devices are typically rated in temperature capability from -55 C to about +225 C, and their characteristics over this temperature range are documented in data sheets. Since electronics in some of NASA space exploration missions are required to operate under extreme temperature conditions, both cold and hot, their characteristic behavior within the full temperature spectrum must be determined to establish suitability for use in space applications. The effects of extreme temperature exposure on the performance of a new commercial-off-the-shelf (COTS) SOI AND gate device were evaluated in this work. The high temperature, quad 2-inputs AND gate device, which was recently introduced by CISSOID, is fabricated using a CMOS SOI process. Some of the specifications of the CHT-7408 chip are listed in a table. By supplying a constant DC voltage to one gate input and a 10 kHz square wave into the other associated gate input, the chip was evaluated in terms of output response, output rise (t(sub r)) and fall times (tf), and propagation delays (using a 50% level between input and output during low to high (tPLH) and high to low (tPHL) transitions). The supply current of the gate circuit was also obtained. These parameters were recorded at various test temperatures between -195 C and +250 C using a Sun Systems environmental chamber programmed at a temperature rate of change of 10 C/min. In addition, the effects of thermal cycling on this chip were determined by exposing it to a total of 12 cycles between -195 C and +250 C. Following the cycling activity, measurements were performed again at the test temperatures of -195 C, +21 C, and +250 C.

Patterson, Richard↗

NASA NTRS2015

The Observed State of the Water Cycle in the Early Twenty-First Century

This study quantifies mean annual and monthly fluxes of Earth's water cycle over continents and ocean basins during the first decade of the millennium. To the extent possible, the flux estimates are based on satellite measurements first and data-integrating models second. A careful accounting of uncertainty in the estimates is included. It is applied within a routine that enforces multiple water and energy budget constraints simultaneously in a variational framework in order to produce objectively determined optimized flux estimates. In the majority of cases, the observed annual surface and atmospheric water budgets over the continents and oceans close with much less than 10% residual. Observed residuals and optimized uncertainty estimates are considerably larger for monthly surface and atmospheric water budget closure, often nearing or exceeding 20% in North America, Eurasia, Australia and neighboring islands, and the Arctic and South Atlantic Oceans. The residuals in South America and Africa tend to be smaller, possibly because cold land processes are negligible. Fluxes were poorly observed over the Arctic Ocean, certain seas, Antarctica, and the Australasian and Indonesian islands, leading to reliance on atmospheric analysis estimates. Many of the satellite systems that contributed data have been or will soon be lost or replaced. Models that integrate ground-based and remote observations will be critical for ameliorating gaps and discontinuities in the data records caused by these transitions. Continued development of such models is essential for maximizing the value of the observations. Next-generation observing systems are the best hope for significantly improving global water budget accounting.

Numerical analysis/modeling↗

NASA NTRS2006

Materials Needs for Future In-Space Propulsion Systems

NASA's In-Space Propulsion Technology Project is developing the next generation of in-space propulsion systems in support of robotic exploration missions throughout the solar system. The propulsion technologies being developed are non-traditional and have stressing materials performance requirements. Earth-storable bipropellant performance is constrained by temperature limitations of the columbium used in the chamber. Iridium/rhenium (Ir/Re) is now available and has been implemented in initial versions of Earth- Storable rockets with specific impulses about 10 seconds higher than columbium rocket chambers. New chamber fabrication methods that improve process and performance of Ir/Re and other promising material systems are needed. The solar sail is a propellantless propulsion system that gains momentum by reflecting sunlight. The sails need to be very large in area (from 10000 sq m up to 62500 sq m) yet be very lightweight in order to achieve adequate accelerations for realistic mission times. Lightweight materials that can be manufactured in thicknesses of less than 1 micron and that are not harmed by the space environment are desired. Blunt Body Aerocapture uses aerodynamic drag to slow an approaching spacecraft and insert it into a science orbit around any planet or moon with an atmosphere. The spacecraft is enclosed by a rigid aeroshell that protects it from the entry heating and aerodynamic environment. Lightweight, high-temperature structural systems, adhesives, insulators, and ablatives are key components for improving aeroshell efficiencies at heating rates of 1000-2000 W/sq cm and beyond. Inflatable decelerators in the forms of ballutes and inflatable aeroshells will use flexible polymeric thin film materials, high temperature fabrics, and structural adhesives. The inflatable systems will be tightly packaged during cruise and will be inflated prior to entry interface at the destination. Materials must maintain strength and flexibility while packaged at cold temperatures (-100 C) for up to 10 years and then withstand the high temperatures (500 C) encountered during aerocapture.

Johnson, Les↗

NASA NTRS2008

Materials Needs for Future In-space Propulsion Systems

NASA is developing the next generation of in-space propulsion systems in support of robotic exploration missions throughout the solar system. The propulsion technologies being developed are non-traditional and have stressing materials performance requirements. (Chemical Propulsion) Earth-storable chemical bipropellant performance is constrained by temperature limitations of the columbium used in the chamber. Iridium/rhenium (Ir/Re) is now available and has been implemented in initial versions of Earth-Storable rockets with specific impulses (Isp) about 10 seconds higher than columbium rocket chambers. New chamber fabrication methods that improve process and performance of Ir/Re and other promising material systems are needed. (Solar Sail Propulsion) The solar sail is a propellantless propulsion system that gains momentum by reflecting sunlight. The sails need to be very large in area (from 10000 m2 up to 62500 m2) yet be very lightweight in order to achieve adequate accelerations for realistic mission times. Lightweight materials that can be manufactured in thicknesses of less than 1 micron and that are not harmed by the space environment are desired. (Aerocapture) Blunt Body Aerocapture uses aerodynamic drag to slow an approaching spacecraft and insert it into a science orbit around any planet or moon with an atmosphere. The spacecraft is enclosed by a rigid aeroshell that protects it from the entry heating and aerodynamic environment. Lightweight, high-temperature structural systems, adhesives, insulators, and ablatives are key components for improving aeroshell efficiencies at heating rates of 1000-2000 W/cu cm and beyond. Inflatable decelerators in the forms of ballutes and inflatable aeroshells will use flexible polymeric thin film materials, high temperature fabrics, and structural adhesives. The inflatable systems will be tightly packaged during cruise and will be inflated prior to entry interface at the destination. Materials must maintain strength and flexibility while packaged at cold temperatures (_100oC) for up to 10 years and then withstand the high temperatures (500oC) encountered during aerocapture. The presentation will describe the status of each propulsion technology and summarize the materials needed for their implementation.

Johnson, Charles Les↗