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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 1,801 records · Page 100

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NASA NTRS1988

Variable mixture ratio performance through nitrogen augmentation

High/variable mixture ratio O2/H2 candidate engine cycles are examined for earth-to-orbit vehicle application. Engine performance and power balance information are presented for the candidate cycles relative to chamber pressure, bulk density, and mixture ratio. Included in the cycle screening are concepts where a third fluid (liquid nitrogen) is used to achieve a variable mixture ratio over the trajectory from liftoff to earth orbit. The third fluid cycles offer a very low risk, fully reusable, low operation cost alternative to high/variable mixture ratio bipropellant cycles. Variable mixture ratio engines with extendible nozzle are slightly lower performing than a single mixture ratio engine (MR = 7:1) with extendible nozzle. Dual expander engines (MR = 7:1) have slightly better performance than the single mixture ratio engine. Dual fuel dual expander engines offer a 16 percent improvement over the single mixture ratio engine.

Beichel, R.↗

NASA NTRS1988

Structural response of SSME turbine blade airfoils

Reusable space propulsion hot gas-path components are required to operate under severe thermal and mechanical loading conditions. These operating conditions produce elevated temperature and thermal transients which results in significant thermally induced inelastic strains, particularly, in the turbopump turbine blades. An inelastic analysis for this component may therefore be necessary. Anisotropic alloys such as MAR M-247 or PWA-1480 are being considered to meet the safety and durability requirements of this component. An anisotropic inelastic structural analysis for an SSME fuel turbopump turbine blade was performed. The thermal loads used resulted from a transient heat transfer analysis of a turbine blade. A comparison of preliminary results from the elastic and inelastic analyses is presented.

Arya, V. K.↗

NASA NTRS1989

Integrated launch and emergency entry vehicle concept

This paper describes the design of an integrated launch and emergency entry vehicle (EEV) that can satisfy the need for an advanced manned vehicle, a heavy-lift launch vehicle, and the Space Station. It is pointed out that, because only one small entry vehicle would need to be developed, the development costs should be relatively low. The EEV would have reasonable operating costs because its expendable tank would be smaller than that of the Space Shuttle and all expensive parts would be reusable. Schematic diagrams are included.

Martin, James A.↗

NASA NTRS1989

The case for the evolution of the Shuttle system

Examining the expendable and reusable space transportation systems currently in use, it appears that the service life of a launch system, with periodic upgrades, can be 30 to 40 years. It is also evident that both reliability and performance increase with flight experience. With a life expectancy of at least 30 years, the Shuttle system is today only in the first third of its potential service life. This paper addresses the issues of why Shuttle evolution is warranted, how the evolutionary process is managed, and how that process can improve system performance and reduce costs. Specific examples of enhancements currently under consideration are presented and discussed in terms of their relationship to the achievement of greater inherent reliability in the Shuttle system.

Cohen, Aaron↗

NASA NTRS1989

A view toward future launch vehicles - A civil perspective

Prospective NASA launch vehicle development efforts, which in addition to follow-on developments of the Space Shuttle encompass the Shuttle-C cargo version, various possible Advanced Launch System (ALS) configurations, and various Heavy Lift Launch System (HLLS) design options. Fully and partially reusable manned vehicle alternatives are also under consideration. In addition to improving on the current Space Shuttle's reliability and flexibility, ALS and HLLV development efforts are expected to concentrate on the reduction of operating costs for the given payload-launch capability.

Darwin, Charles R.↗

NASA NTRS1988

Some tether applications with small launch vehicles

This paper describes four potentially useful tether applications that are best suited to or even limited to small launch vehicles and their payloads. The first application uses a Small Expendable-tether Deployment System (SEDS) to deploy a 20-40 km tether which transfers momentum from the spent final stage to its payload. This de-orbits the stage and boosts the payload. The other concepts are endo-atmospheric applications specific to various kinds of boosters. In one, a large helicopter gently recovers a reusable booster while it is descending by parachute. Another improves the performance of a small single-stage-to-orbit launch vehicle, by lifting it by helicopter to an altitude of about 3 kilometers before ignition. The last tows an air-breathing launch vehicle to high speed behind a supersonic aircraft, to eliminate any need for low-speed propulsion on the launch vehicle.

Carroll, Joseph A.↗

NASA NTRS1990

Aerodynamic separation and glideback of a Mach 3 staged booster

A study has been conducted of the staging maneuver for a two-stage, vertical-takeoff, launch vehicle and the subsequent glideback of the booster to a launch site runway. The booster and orbiter are both winged, fully reusable, and have liquid rocket main propulsion. for the staging maneuver, which nominally occurs at Mach 3, a flight control system and separation technique were devised which produced successful separations over a wide range of staging angles of attack. A guidance algorithm was developed for the unpowered glide-back of the booster from the completion of the staging maneuver to touchdown on a runway in the vicinity of the launch site. Using this guidance technique, the booster successfully touched down on the runway while being subjected to a wide range of seasonal and perturbed atmospheric conditions with and without wids as well as errors in staging state conditions and errors in predicted aerodynamics.

Naftel, J. C.↗

NASA NTRS1990

Evaluation of abort capabilities of rocket-powered single-stage-to-orbit launch vehicles

Application of advanced technologies to future launch vehicle designs would allow the introduction of a rocket-powered, single-stage-to-orbit (SSTO) launch system early in the next century. A fully reusable SSTO vehicle would be quite desirable from an operational standpoint; however, such a vehicle cannot be designed without accompanying technological advances in structure, propulsion, and subsystems. The conceptual design of such a vehicle has recently been completed. This paper examines the abort capabilities of an advanced SSTO launch vehicle which has five main engines. In the event of a single or dual main engine shutdown it was determined when the vehicle could execute return-to-launch-site, abort-to-orbit, or down-range abort maneuvers. Throughout each abort maneuver, vehicle loads are kept within nominal ascent and entry design values.

Stanley, Douglas O.↗

NASA NTRS1990

Improvement of pump tubes for gas guns and shock tube drivers

In a pump tube, a gas is mechanically compressed, producing very high pressures and sound speeds. The intensely heated gas produced in such a tube can be used to drive light gas guns and shock tubes. Three concepts are presented that have the potential to allow substantial reductions in the size and mass of the pump tube to be achieved. The first concept involves the use of one or more diaphragms in the pump tube, thus replacing a single compression process by multiple, successive compressions. The second concept involves a radical reduction in the length-to-diameter ratio of the pump tube and the pump tube piston. The third concept involves shock heating of the working gas by high explosives in a cyclindrical geometry reusable device. Preliminary design analyses are performed on all three concepts and they appear to be quite feasible. Reductions in the length and mass of the pump tube by factors up to about 11 and about 7, respectively, are predicted, relative to a benchmark conventional pump tube.

Bogdanoff, D. W.↗

NASA NTRS1990

Design considerations for honeycomb double-bubble tanks utilized in aerospace vehicles

Future earth-to-orbit vehicles will utilize reusable internal fuel tanks requiring most of the vehicle volume. Because weight reduction and volumetric efficiency are essential, these tanks will be integral to the vehicle structure, and the vehicle cross section will dictate the size and shape of the tank. For a fuselage with a noncircular cross section, a multibubble tank configuration may be employed. Multibubble tanks offer a reduction in tank weight over a single-lobe tank or several cylindrical tanks in a noncircular cross section but sacrifice the efficiency of utilizing the entire fuselage cross section. This paper includes a general configuration analysis of a double-bubble tank, a materials and efficiency study, and a complete mission design analysis centers on sizing the tank to withstand ascent and reentry loads while minimizing the structural weight.

Lentz, Christopher A.↗

NASA NTRS1990

Aerothermodynamic environments for Mars entry, Mars return, and lunar return aerobraking missions

The aeroheating environments to vehicles undergoing Mars aerocapture, earth aerocapture from Mars, and earth aerocapture from the moon are presented. An engineering approach for the analysis of various types of vehicles and trajectories was taken, rather than performing a benchmark computation for a specific point at a selected time point in a trajectory. The radiation into Mars using the Mars Rover Sample Return (MRSR) 2-ft nose radius bionic remains a small contributor of heating for 6 to 10 km/sec; however, at 12 km/sec it becomes comparable with the convection. For earth aerocapture, returning from Mars, peak radiation for the MRSR SRC is only 25 percent of the peak convection for the 12-km/sec trajectory. However, when large vehicles are considered with this trajectory, peak radiation can become 2 to 4 times higher than the peak convection. For both Mars entry and return, a partially ablative Thermal Protection System (TPS) would be required, but for Lunar Transfer Vehicle return an all-reusable TPS can be used.

Rochelle, W. C.↗

NASA NTRS1990

U.S. liquid propulsion technology stands still

NASA is currently conducting a comprehensive reevaluation of such fundamental launch vehicle-related issues as the appropriate balance between manned and unmanned launchers, reusable vs expendable vehicles and vehicle components, and the evolutionary vs innovative new-system development courses. Two trends have emerged, both emphasizing system expense: (1) low cost/manned flight vehicles for personnel transfer, small payloads, servicing, rescue, and Space Station logistics, and (2) low cost/unit mass vehicles for large and multiple payloads, propellants, bulk cargoes, and heavy upper stages. Attention is given to the contributions made by CFD to liquid-fueled engine component design.

Mccarty, John P.↗

NASA NTRS1990

Effects of state recovery on creep buckling under variable loading

Structural alloys embody internal mechanisms that allow recovery of state with varying stress and elevated temperature, i.e., they can return to a softer state following periods of hardening. Such material behavior is known to strongly influence structural response under some important thermomechanical loadings, for example, that involving thermal ratchetting. The influence of dynamic and thermal recovery on the creep buckling of a column under variable loading is investigated. The column is taken as the idealized (Shanley) sandwich column. The constitutive model, unlike the commonly employed Norton creep model, incorporates a representation of both dynamic and thermal (state) recovery. The material parameters of the constitutive model are chosen to characterize Narloy Z, a representative copper alloy used in thrust nozzle liners of reusable rocket engines. Variable loading histories include rapid cyclic unloading/reloading sequences and intermittent reductions of load for extended periods of time; these are superimposed on a constant load. The calculated results show that state recovery significantly affects creep buckling under variable loading.

Robinson, D. N.↗

NASA NTRS1990

A candidate architecture for monitoring and control in chemical transfer propulsion systems

To support the exploration of space, a reusable space-based rocket engine must be developed. This engine must sustain superior operability and man-rated levels of reliability over several missions with limited maintenance or inspection between flights. To meet these requirements, an expander cycle engine incorporating a hghly capable control and health monitoring system is planned. Alternatives for the functional organization and the implementation architecture of the engine's monitoring and control system are discussed. On the basis of this discussion, a decentralized architecture is favored. The trade-offs between several implementation options are outlined and future work is proposed.

Binder, Michael↗

NASA NTRS1990

The simulation of fluid dynamic uncertainties in the SSME turbopump

The aerodynamic uncertainties in the reusable rocket engine turbopumps due to randomness in their cryogenic environment are investigated. The probabilistic simulation of the turbopumps' aerodynamic response was accomplished using the panel method coupled with Fast Probability Integration methods. The results are presented for the probabilistic rotor blades and splitter loading and their sensitivity to specified flow coefficient and rotor preswirl variance.

Hamed, Awatef↗

NASA NTRS1990

Potential orbital use of the Space Shuttle External Tanks

The Space Shuttle was designed so that major components would be reusable; however, it has been shown that it would not be cost-effective to reuse the external tanks in the same manner as the solid rocket boosters. Studies have been conducted to investigate using the external tanks on orbit. Utilizing an external tank on-orbit appears simple enough, since the tank obtains 98 percent of orbital velocity during the Space Shuttle's ascent phase. However, there are many requirements, issues, and difficulties that users must be aware of and satisfy. Studies for converting an external tank into a gamma-ray imaging telescope (GRIT), conducted at NASA Marshall, have identified many of these specific issues. Results of the GRIT studies and other considerations for the potential conversion of external tanks into useful space resources are summarized.

Nein, Max E.↗

NASA NTRS1990

Benefits of electric propulsion for the Space Exploration Initiative

An overview of the benefits which may be derived through the use of electric propulsion in support of the Space Exploration Initiative is presented. Lunar cargo, Mars cargo and piloted Mars vehicles using electric propulsion are considered. The high performance of electric propulsion systems is shown to offer substantial benefits for these applications, including: substantially reduced initial masses in low earth orbit, reduced round-trip times for piloted Mars vehicles, availability of large amounts of electrical power en route and at the destination, less sensitivity to launch dates and windows, reusability, and growth potential for human exploration of the solar system. Hybrid chemical/NEP and NTR/NEP vehicles are discussed for their potential to reduce piloted round-trip time to Mars even further. A brief technology assessment of the major electric propulsion system components is also presented.

Brophy, John R.↗

NASA NTRS1990

LifeSat - Program overview

LifeSat is a reusable reentry satellite for use by NASA and the international space life sciences community. It will be employed beginning in late 1994, to understand the effects of the space environment (i.e., high energy space radiation and microgravity) on biological systems. This paper describes the program for LifeSat.

Gilbreath, William P.↗

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