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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 145 records · Page 8
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NASA NTRS1976

IUS capability for earth orbital spacecraft

A basic element of the Space Transportation System (STS) is the Interim Upper Stage (IUS) currently being planned by the USAF. The prime use of the basic two-stage IUS vehicle by both NASA and DoD will be to place satellites in geosynchronous orbit. High energy missions such as planetary will be accomplished by three- and four-stage vehicles. Various other configurations such as tandem vehicles in a single Shuttle and multiple spacecraft are being planned. Also under study is a smaller stage that can be flown with other payloads on a space available basis. The various concepts being studied offer economical transportation alternatives to potential users. This paper describes details of the IUS program status, design requirements, concepts and applications to Earth orbital spacecraft.

Huber, W. G.↗
NASA NTRS1976

Advanced radial inflow turbine rotor program: Design and dynamic testing

The advancement of small, cooled, radial inflow turbine technology in the area of operation at higher turbine inlet temperature is discussed. The first step was accomplished by designing, fabricating, and subjecting to limited mechanical testing an advanced gas generator rotating assembly comprising a radial inflow turbine and two-stage centrifugal compressor. The radial inflow turbine and second-stage compressor were designed as an integrally machined monorotor with turbine cooling taking place basically by conduction to the compressor. Design turbine inlet rotor gas temperature, rotational speed, and overall gas generator compressor pressure ratio were 1422 K (2560 R), 71,222 rpm, and 10/1 respectively. Mechanical testing on a fabricated rotating assembly and bearing system covered 1,000 cold start/stop cycles and three spins to 120 percent design speed (85,466 rpm).

Rodgers, C.↗
NASA NTRS1977

Ceramics for the advanced automotive gas turbine engine: A look at a single shaft design

The results of a preliminary analysis of a single shaft regenerative design with a single stage radial turbine are presented to show the fuel economy that can be achieved at high turbine inlet temperatures, with this particular advanced design, if the turbine tip speed and regenerator inlet temperature are not limited. The engine size was 100 hp for application to a 3500 lb auto. The fuel economy was analyzed by coupling the engine to the auto through a continuously variable speed-ratio transmission and operating the system at constant turbine inlet temperature over the Composite Driving Cycle. The fuel was gasoline and the analysis was for a 85 F day. With a turbine inlet temperature of 2500 F the fuel economy was 26.2 mpg, an improvement of 18 percent over that of 22.3 mpg with a turbine inlet temperature of 1900 F. The turbine tip speed needed for best economy with the 2500 F engine was 2530 ft/sec. The regenerator temperature was approximately 2200 F at idle. Disk stresses were estimated for one single stage radial turbine and two two-stage radial-axial turbines and compared with maximum allowable stress curves estimated for a current ceramic material. Results show a need for higher Weibull Modulus, higher strength ceramics.

Nosek, S. M.↗
NASA NTRS1976

Remote sensing as a tool for watershed-wide estimation of net solar radiation and water loss to the atmosphere

Results are presented for a study intended to develop a general remote sensing-aided cost-effective procedure to estimate watershed-wide water loss to the atmosphere via evapotranspiration and to estimate net solar radiation over the watershed. Evapotranspiration estimation employs a basic two-stage two-phase sample of three information resolution levels. Net solar radiation is taken as one of the variables at each level of evapotranspiration modeling. The input information for models requiring spatial information will be provided by Landsat digital data, environmental satellite data, ground meteorological data, ground sample unit information, and topographic data. The outputs of the sampling-estimation/data bank system will be in-place maps of evapotranspiration on a data resolution element basis, watershed-wide evapotranspiration isopleths, and estimates of watershed and subbasin total evapotranspiration with associated statistical confidence bounds. The methodology developed is being tested primarily on the Spanish Creek Watershed Plumas County, California.

Khorram, S.↗
NASA NTRS1977

Evolution of mare basalts - The complexity of the U-Th-Pb system

An attempt has been made to gain more insight into mare-basalt evolution by performing a very detailed leaching and mineral-separation U-Th-Pb systematics study on mare basalt 15085. It is found that about 20-50% of the U, Th, and Pb reside on the grain boundaries or in the mesostasis and that the Pb-207/Pb-206 ratios of the grain boundaries and crystal interiors are distinctly different. These distinct trends appear to represent either continuous or episodic postcrystallizational disturbances to the U-Th-Pb system of this rock. Using U and Pb partition coefficients, it is concluded that existing two- and three-stage U-Pb evolution models do not accurately describe mare-basalt genesis. An alternative two-stage + KREEP mixing model is proposed as a simple approximation to U-Pb evolution in lunar rocks. Most Rb-Sr and Sm-Nd data are compatible with this model.

Unruh, D. M.↗
NASA NTRS1980

Looking toward single-stage launch systems

The prospects for utilization of single-stage-to-orbit launch vehicles to meet requirements for improved space transportation economics near the turn of the century are addressed. Vehicle concepts based on incremental or derivative advancements in technology are described. Comparative sizing and economic features of single and two-stage concepts are shown.

Eldred, C. H.↗
NASA NTRS1986

Analysis of separation of a two-stage winged launch vehicle

The feasibility of separating a two-stage winged, parallel configured launch vehicle at Mach 3 has been determined. The two-stage fully-reusable launch vehicle is comprised of an orbiter, which can be used alone as a single-stage-to-orbit vehicle, and a booster which is identical to the baseline orbiter design. A three-degree-of-freedom analysis was conducted on the ascent trajectory of the two-stage concept to determine the optimum staging conditions. The booster provides all propellant for both stages using a crossfeed system until staging occurs. After staging, the empty booster glides back to the launch site for a horizontal landing. To analyze the separation maneuver, the six-degree-of-freedom equations of motion for both stages were numerically integrated using a computer program which incorporates analytically derived interference aerodynamic data. The necessary conditions for a successful separation were determined and required the use of active controls on both the orbiter and booster. Once the booster and orbiter were separated, a three-degree-of-freedom analysis was conducted on the booster's glideback trajectory.

Naftel, J. C.↗
NASA NTRS1988

Force-Balance Dynamic Display

Balance dynamic display unit (BDDU) is compact system conditioning six dynamic analog signals so they are monitored simultaneously in real time on single-trace oscilloscope. Typical BDDU oscilloscope display in scan mode shows each channel occupying one-sixth of total trace. System features two display modes usable with conventional, single-channel oscilloscope: multiplexed six-channel "bar-graph" format and single-channel display. Two-stage visual and audible limit alarm provided for each channel.

Ferris, Alice T.↗
NASA NTRS1988

HASA: Hypersonic Aerospace Sizing Analysis for the Preliminary Design of Aerospace Vehicles

A review of the hypersonic literature indicated that a general weight and sizing analysis was not available for hypersonic orbital, transport, and fighter vehicles. The objective here is to develop such a method for the preliminary design of aerospace vehicles. This report describes the developed methodology and provides examples to illustrate the model, entitled the Hypersonic Aerospace Sizing Analysis (HASA). It can be used to predict the size and weight of hypersonic single-stage and two-stage-to-orbit vehicles and transports, and is also relevant for supersonic transports. HASA is a sizing analysis that determines vehicle length and volume, consistent with body, fuel, structural, and payload weights. The vehicle component weights are obtained from statistical equations for the body, wing, tail, thermal protection system, landing gear, thrust structure, engine, fuel tank, hydraulic system, avionics, electral system, equipment payload, and propellant. Sample size and weight predictions are given for the Space Shuttle orbiter and other proposed vehicles, including four hypersonic transports, a Mach 6 fighter, a supersonic transport (SST), a single-stage-to-orbit (SSTO) vehicle, a two-stage Space Shuttle with a booster and an orbiter, and two methane-fueled vehicles.

Harloff, Gary J.↗
NASA NTRS1989

Two stage sorption type cryogenic refrigerator including heat regeneration system

A lower stage chemisorption refrigeration system physically and functionally coupled to an upper stage physical adsorption refrigeration system is disclosed. Waste heat generated by the lower stage cycle is regenerated to fuel the upper stage cycle thereby greatly improving the energy efficiency of a two-stage sorption refrigerator. The two stages are joined by disposing a first pressurization chamber providing a high pressure flow of a first refrigerant for the lower stage refrigeration cycle within a second pressurization chamber providing a high pressure flow of a second refrigerant for the upper stage refrigeration cycle. The first pressurization chamber is separated from the second pressurization chamber by a gas-gap thermal switch which at times is filled with a thermoconductive fluid to allow conduction of heat from the first pressurization chamber to the second pressurization chamber.

Jones, Jack A.↗
NASA NTRS1989

Thermal re-design of the Galileo spacecraft for a Venus-earth-earth-gravity assist (VEEGA) trajectory

The cancellation of the Centaur upper stage program in the aftermath of the Challenger tragedy forced a redesign of the flight trajectory of the Galileo spacecraft to Jupiter, i.e., from a direct trajectory to the Venus-earth-earth-gravity-assist (VEEGA) trajectory on the lower energy two-stage inertial upper stage (IUS), with the result that the spacecraft would be exposed to more than twofold increase in peak solar irradiance. This paper describes the general system-level thermal redesign effort for the Galileo spacecraft, from the start of feasibility studies to its final implementation. Results indicate that the addition of sunshades and the generous utilization of second-surface aluminized Kapton surface material for reflecting high percentages of incident solar irradiation would 'harden' the spacecraft's existing thermal protection system adequately, provided that sun-pointing at the relatively higher solar irradiance levels could be maintained. The final miximum flight temperature predictions for the spacecraft's subsystem thermal designs are given.

Reeve, R.↗
NASA NTRS1989

Design and test of an oxygen turbopump for a dual expander cycle rocket engine

A liquid oxygen (LOX) turbopump with an 860 R gaseous oxygen (GOX) turbine drive was designed for a 3750 lb thrust dual expander cycle rocket engine. This turbopump, which requires no interpropellant seals or system purges, features a 156 hp, single stage, full admission, impulse turbine; an axial flow inducer; a two-stage centrifugal pump with unshrouded impellers; long-life, LOX-lubricated, self-aligning, hydrostatic bearings; and a subcritical rotor design. It is constructed of Monel, a nickel-copper alloy, which has low ignition potential in oxygen. The pump was designed to deliver 34.7 gpm of 4655 psia liquid oxygen at a shaft speed of 75,000 rpm. The dual expander cycle rocket engine and the performance it requires of the LOX turbopump will be discussed as well as the design of the pump, turbine, bearings, and the turbopump rotordynamics. The test program and preliminary test results will also be presented.

Buckmann, P. S.↗
NASA NTRS1989

Dual-fuel versus single-fuel propulsion systems for AMLS applications

The results of using a computerized preliminary design system to integrate propulsion systems examined as a part of the Space Transportation Main Engine (STME) and Space Transportation Booster Engine (STBE) studies with reference vehicle concepts from the Advanced Manned Launch System (AMLS) study are presented. The major trade study presented is an analysis of the effect of using a single fuel for both stages of two-stage AMLS reference vehicles as opposed to using a separate fuel for the boosters. Other trade studies presented examine the effect of varying relevant engine parameters in an attempt to optimize the reference engines for use with the AMLS launch vehicles. In each propulsion trade discussed, special attention is given to the major vehicle performance and operational issues involved.

Stanley, Douglas O.↗
NASA NTRS1989

Application of a hybrid generation/utility assessment heuristic to a class of scheduling problems

A two-stage heuristic solution approach for a class of multiobjective, n-job, 1-machine scheduling problems is described. Minimization of job-to-job interference for n jobs is sought. The first stage generates alternative schedule sequences by interchanging pairs of schedule elements. The set of alternative sequences can represent nodes of a decision tree; each node is reached via decision to interchange job elements. The second stage selects the parent node for the next generation of alternative sequences through automated paired comparison of objective performance for all current nodes. An application of the heuristic approach to communications satellite systems planning is presented.

Heyward, Ann O.↗
NASA NTRS1990

Advanced Launch Vehicle Upper Stages Using Liquid Propulsion and Metallized Propellants

Metallized propellants are liquid propellants with a metal additive suspended in a gelled fuel or oxidizer. Typically, aluminum (Al) particles are the metal additive. These propellants provide increase in the density and/or the specific impulse of the propulsion system. Using metallized propellant for volume-and mass-constrained upper stages can deliver modest increases in performance for low earth orbit to geosynchronous earth orbit (LEO-GEO) and other earth orbital transfer missions. Metallized propellants, however, can enable very fast planetary missions with a single-stage upper stage system. Trade studies comparing metallized propellant stage performance with non-metallized upper stages and the Inertial Upper Stage (IUS) are presented. These upper stages are both one- and two-stage vehicles that provide the added energy to send payloads to altitudes and onto trajectories that are unattainable with only the launch vehicle. The stage designs are controlled by the volume and the mass constraints of the Space Transportation System (STS) and Space Transportation System-Cargo (STS-C) launch vehicles. The influences of the density and specific impulse increases enabled by metallized propellants are examined for a variety of different stage and propellant combinations.

Palaszewski, Bryan A.↗
NASA NTRS1959

1958 NASA/USAF Space Probes (Able-1): Summary - Volume 1

Early in calendar year 1958 Space Technology Laboratories, Inc. (STL) (then Space Technology Laboratories, a division of the Ramo-Wooldridge Corp.) developed for the Air Force Ballistic Missile Division (AFBMD) an Advanced Re-entry Test Vehicle (ARTV) for the purpose of testing ballistic missile nose cones at the full range of 5500 nautical miles. The two-stage ARTV utilized the Thor ballistic missile and the second stage propulsion system developed for the Vanguard program. In late 1957 and early 1958, STL/AFBMD prepared studies of various missile combinations which could be utilized for space testing. The Thor, in combination with the Vanguard second and third stages, was one of the vehicles considered which offered a very early capability of placing a reasonable payload in a lunar orbit. These STL/AFBMD studies were presented to various appropriate groups including the Killian, Millikan, H. J . Stewart Committees; Headquarters, Air Research and Development Command, and ARDC Centers. Subsequently the Advanced Research Projects Agency (ARPA) contacted STL relative to the availability of hardware for an early lunar shot. By utilizing existing spares already purchased for the ARTV, and by making use of the ARTV contractors already in being, it appeared feasible to launch by the third quarter of calendar year 1958 a payload which would be captured by the moon's gravitational force. On 27 March 1958, ARPA directed STL to proceed with a program of three lunar shots. As much as possible, these shots were to utilize existing ARTV spare hardware and impose no interference with the ballistic missile programs. In September this program was transferred to the direction of the National Aeronautics and Space Administration (NASA). On 17 August 1958 the first launching of the Able-1 vehicle was attempted, but the flight was terminated by a propulsion failure of the first stage. Subsequent launchings were attempted on 13 October and 8 November 1958. Of these launchirigs the October attempt was the most successful. Although the payload did not reach the vicinity of the moon, a maximum altitude of 71,700 was attained, and useful scientific data was obtained from the instrumentation.

Source record↗
NASA NTRS1969

Biosatellite II mission

Biosatellite B was launched from Cape Kennedy, Florida, on a two-stage DELTA launch vehicle at 6:04 p.m. on 7 September, 1967. Approximately nine minutes later the 435 kg spacecraft biological laboratory was placed into a satisfactory 315 km near-circular earth orbit, successfully separated from the launch vehicle's second stage and was designated Biosatellite II. The scientific payload consisting of thirteen selected general biology and radiation experiments were subjected to planned, carefully controlled environmental conditions during 45 hours of earth-orbital flight. The decision was made to abbreviate the scheduled 3-day mission by approximately one day because of a threatening tropical storm in the recovery area, and a problem of communication with the spacecraft from the tracking stations. Highest priority was placed on recovery which was essential to obtain the scientific results on all the experiments. The operational phase of the mission came to a successful conclusion with the deorbit of the recovery capsule, deployment of the parachute system and air recovery by the United States Air Force. The 127 kg recovery capsule was returned to biology laboratories at Hickam Air Force Base, Hawaii, for disassembly and immediate inspection and analysis of the biological materials by the experimenters. It was evident immediately that the quality of the biology was excellent and this fact gave promise of a high return of scientific data. The environmental conditions provided to the experimental material in the spacecraft, provisions for experimental controls, and operational considerations are presented as they relate to interpretation of the experimental results.

Biosatellite 2 Project↗
NASA NTRS1989

Two-statge sorption type cryogenic refrigerator including heat regeneration system

A lower stage chemisorption refrigeration system physically and functionally coupled to an upper stage physical adsorption refrigeration system. Waste heat generated by the lower stage cycle is regenerated to fuel the upper stage cycle thereby greatly improving the energy efficiency of a two-stage sorption refrigerator. The two stages are joined by disposing a first pressurization chamber providing a high pressure flow of a first refrigerant for the lower stage refrigeration cycle within a second pressurization chamber providing a high pressure flow of a second refrigerant for the upper stage refrigeration cycle. The first pressurization chamber is separated from the second pressurization chamber by a gas-gap thermal switch which at times is filled with a thermoconductive fluid to allow conduction of heat from the first pressurization chamber to the second pressurization chamber.

Jones, Jack A.↗