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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

Effect of low-stiffness closeout overwrap on rocket thrust-chamber life

Three rocket thrust chambers with copper liners and a thrust level of 20.9 kN were cyclically test fired to failure. Two of the liners were made from oxygen free, high conductivity (OFHC) copper and from annealed Amzirc. The milled coolant channels were closed out with a thin copper closeout over which a fiberglass composite was wrapped to provide hoop strength only. Experimental data are presented, along with the results of a preliminary analysis that was performed before fabrication to evaluate the life extending potential of a thin copper closeout with a fiberglass overwrap.

Kasper, H. J.↗
NASA NTRS1979

Variational description of the positive column with two-stem ionization

The ionization balance in diffusion dominated discharges which depends on both one and two step ionization processes is considered. The Spenke diffusion equation (D sq delta n + neutrino n + sq kn =0) describing such conditions is solved by the Rayleigh-Ritz variational method. Simple analytic approximations to the density profile, and the similarity relation between neutrino,k,D and the discharge dimensions, are derived for planar and cylindrical geometry, and compared with exact computations for certain limiting cases.

Crawford, F. W.↗
NASA NTRS1979

Experimental study of acoustic loads on an upper-surface-blown STOL airplane configuration

Fluctuating pressure levels were measured on the flap and fuselage of an upper-surface-blown jet-flap airplane configuration in a wind tunnel. The model tested had turbofan engines with a bypass ratio of 3 and a thrust rating of 10 kN. Rectangular nozzles were mounted flush with the upper surface at 35 percent of the wing chord. Test parameters were flap deflection angle, jet impingement angle, angle of attack, free-stream velocity, spanwise location of the engine, and jet dynamic pressure. Load levels were high throughout the jet impingement region, with the highest levels (about 159 dB) occurring on the fuselage and near the knee of the flap. The magnitude of the forward-velocity effect appeared to depend upon the ratio of free-stream and jet velocities. Good agreement was obtained between fluctuating pressure spectra measured at jet dynamic pressures of 7 and 22 kPa when the spectra were scaled by nondimensional functions of dynamic pressure, velocity, and the empirical relationship between dynamic pressure and overall fluctuating pressure level.

Willis, C. M.↗
NASA NTRS1980

Dual throat engine design for a SSTO launch vehicle

A propulsion system analysis of a dual fuel, dual throat engine for launch vehicle application was conducted. Basic dual throat engine characterization data are presented to allow vehicle optimization studies to be conducted. A preliminary baseline engine system was defined. Dual throat engine performance, envelope, and weight parametric data were generated over the parametric range of thrust from 890 to 8896 KN (200K to 2M lb-force), chamber pressure from 6.89 million to 34.5 million N/sq m (1000 to 5000 psia) thrust ratio from 1.2 to 5, and a range of mixture ratios for the two tripropellant combinations: LO2/RP-1 + LH2 and LO2/LCH4 + LH2. The results of the study indicate that the dual fuel dual throat engine is a viable single stage to orbit candidate.

Obrien, C. J.↗
NASA NTRS1979

Ultimate strength of high-load-capacity composite bolted joints

Presented are the results of a series of tests initiated to obtain baseline data on the load-carrying capacity of bolted joints designed to carry large loads, specifically up to 222 kN (50 kips). The major testing purposes were to determine the load carrying capacity as a function of the width and thickness of the joint and the diameter and number of bolts, and to observe the failure mode. A total of 100 tests were conducted on three different specimen configurations. The specimens were fabricated from a T300/5208 fiber/resin system in a quasi-isotropic lay-up. The results presented indicate that for a given ratio of specimen width to hole diameter, the specimens with the smaller holes sustained a higher net-section tensile stress before failure. In addition, for a given ratio of specimen width to hole diameter, the thinner specimens withstood a higher net-section stress. No attempt has been made to correlate the results with theoretical predictions.

Hyer, M. W.↗
NASA NTRS1980

An active control system for helicopter vibration reduction by higher harmonic pitch

An analytical study defining the basic configuration of an active control system to reduce helicopter vibrations is presented. Theoretical results for a nonlinear four-bladed single rotor helicopter simulation are discussed, showing that vibration reductions on the order of 80-90% for airspeeds up to 150 kn can be expected when using a higher harmonic pitch in an active feedback control system. The rotor performance penalty associated with this level of vibration reduction is about 1-3% and the increase in rotor blade stresses is considered to be low. The location of sensor accelerometers proved to be significant for vibration reductions, and it is noted that the RTSA controller is tolerant of sensor signal noise.

Taylor, R. B.↗
NASA NTRS1983

The influence of inlet design on the aeroacoustic performance of a JT15D turbofan engine as measured in the NASA-Ames 40 x 80 foot wind tunnel

The influence of inlet design on the aeroacoustic performance of a JT15D turbofan engine was investigated in the NASA-Ames 40 x 80 ft wind tunnel. Four inlet designs representative of those found in commercial service were evaluated at a forward velocity of 80 kn. The inlet designs considered were selected to evaluate the importance of circumferential static pressure distortion (SPD) for a diffusing sound suppressing inlet and throat Mach number (Mth) for a hardwall inlet on the acoustic characteristics of the JT15D engine. Several speed points were studied simulating the operating conditions spanning approach to take-off. Results demonstrated that the dependency of the forward radiated noise on inlet aerodynamic design parameters was significant (about 5 dB) for the range of SPD (0.1-3.0%) and Mth (0.3-0.6) tested. The performance of the bulk absorber acoustic treatment used in the diffusing sound suppressing inlet is also discussed.

Holm, R. G.↗
NASA NTRS1985

Tethered Satellite System Project Overview

The Skyhook concept is reviewed and the use of a tethered satellite system (TSS) to enable scientific investigations from the shuttle using a closed loop control system is examined. The tethered satellite system has capabilities for deployment toward or away from Earth, for multiple round trip missions, and for deployment at distances up to 100 KN from the orbiter. The deployer, which consists of an entendable boom, a reel for the tether, and the tether itself, permits deployment and retrieval at a safe distance, allows alignment of the force vector of the tether through the center of gravity of the shuttle, and gives some initial gravity gradient separation to aid in deployment and ultimate retrieval of the tethered satellite. Charts show TSS activities in terms of systems studies, key guidelines, Italian and U.S. responsibilities, user activities, and major science and applications accommodation features. Scientific objectives for TSS-1 and TSS-2 verification missions and the current status of the project are also given.

Laue, J. H.↗
NASA NTRS1989

A VLSI chip set for real time vector quantization of image sequences

The architecture and implementation of a VLSI chip set that vector quantizes (VQ) image sequences in real time is described. The chip set forms a programmable Single-Instruction, Multiple-Data (SIMD) machine which can implement various vector quantization encoding structures. Its VQ codebook may contain unlimited number of codevectors, N, having dimension up to K = 64. Under a weighted least squared error criterion, the engine locates at video rates the best code vector in full-searched or large tree searched VQ codebooks. The ability to manipulate tree structured codebooks, coupled with parallelism and pipelining, permits searches in as short as O (log N) cycles. A full codebook search results in O(N) performance, compared to O(KN) for a Single-Instruction, Single-Data (SISD) machine. With this VLSI chip set, an entire video code can be built on a single board that permits realtime experimentation with very large codebooks.

Baker, Richard L.↗
NASA NTRS1989

STOL and STOVL hot gas ingestion and airframe heating tests in the NASA Lewis 9- by 15-foot low-speed wind tunnel

Short takeoff and landing (STOL) and advanced short takeoff and vertical landing (STOVL) aircraft are being pursued for deployment near the end of this century. These concepts offer unique capabilities not seen in conventional aircraft: for example, shorter takeoff distances and the ability to operate from damaged runways and remote sites. However, special technology is critical to the development of this unique class of aircraft. Some of the real issues that are associated with these concepts are hot gas ingestion and airframe heating while in ground effects. Over the past nine years, NASA Lewis Research Center has been involved in several cooperative programs in the 9- by 15 Foot Low-Speed Wind Tunnel (LSWT) to establish a database for hot gas ingestion and airframe heating. The modifications are presented that were made in the 9- by 15-Foot LSWT, including the evolution of the ground plane, model support system, and tunnel sidewalls; and flow visualization techniques, instrumentation, test procedures, and test results. The 9- by 15-Foot LSWT tests were conducted at full scale exhaust nozzle pressure ratios. The headwind velocities varied from 8 to 120 kn depending on the concept (STOL or STOVL). Typical compressor-face distortions (pressure and temperature), ground plane contours, and model surface temperature profiles are presented.

Johns, Albert L.↗
NASA NTRS1990

Serial Escape System For Aircraft Crews

Emergency escape system for aircraft and aerospace vehicles ejects up to seven crewmembers, one by one, within 120 s. Intended for emergencies in which disabled craft still in stable flight at no more than 220 kn (113 m/s) equivalent airspeed and sinking no faster than 110 ft/s (33.5 m/s) at altitudes up to 50,000 ft (15.2 km). Ejection rockets load themselves from magazine after each crewmember ejected. Jumpmaster queues other crewmembers and helps them position themselves on egress ramp. Rockets pull crewmembers clear of aircraft structure. Provides orderly, controlled exit and avoids ditching at sea or landing in rough terrain.

Wood, Kenneth E.↗
NASA NTRS1989

Engine inlet distortion in a 9.2 percent scaled vectored thrust STOVL model in ground effect

Advanced Short Takeoff/Vertical Landing (STOVL) aircraft which can operate from remote locations, damaged runways, and small air capable ships are being pursued for deployment around the turn of the century. To achieve this goal, a cooperative program has been defined for testing in the NASA Lewis 9- by 15-foot Low Speed Wind Tunnel (LSWT) to establish a database for hot gas ingestion, one of the technologies critical to STOVL. This paper presents results showing the engine inlet distortions (both temperature and pressure) in a 9.2 percent scale Vectored Thrust STOVL model in ground effects. Results are shown for the forward nozzle splay angles of 0, -6, and 18 deg. The model support system had 4 deg of freedom, heated high pressure air for nozzle flow, and a suction system exhaust for inlet flow. The headwind (freestream) velocity was varied from 8 to 23 kn.

Johns, Albert L.↗
NASA NTRS1990

Preliminary analysis of measured sound propagation over various seasonal snow covers

Measurements of acoustic pulse propagation in the 5 to 500-Hz frequency band were conducted under various snow cover conditions during the 1989 to 1990 winter in New Hampshire. The objective was to determine the effect of snow cover thickness and other snow properties on the absorption of acoustic pulses. Blank pistol shots were used as the source of the acoustic waves, and geophones and microphones in an 80 m-long linear array served as receivers. Snow thicknesses ranged from 0.05 to 0.35 m, and densities varied from 100 to 350 kg m(sup -3) during the 10 separate measurement days. Preliminary analysis indicates that the peak pulse amplitude decayed in proportion to approx. gamma (sup -1.7) for most conditions and that the acoustic-to-seismic ratios varied from about 4 to 15 x 10(exp -6) m s(sup -1) Pa(sup -1). Theoretical waveforms were calculated for propagation in a homogeneous atmosphere using Attenborough's model of ground impedance. An automatic fitting procedure for the normalized experimental and theoretical waveforms was used to determine the effective flow resistivity of the snow covers, and gave values of 10 to 35 kN s m(sup -4), in agreement with earlier results.

Albert, Donald G.↗
NASA NTRS1990

Hot gas ingestion test results of a two-poster vectored thrust concept with flow visualization in the NASA Lewis 9- by 15-foot low speed wind tunnel

A 9.2 percent scale STOVL hot gas ingestion model was tested in the NASA Lewis 9 x 15-foot Low-Speed Wind Tunnel. Flow visualization from the Phase 1 test program, which evaluated the hot ingestion phenomena and control techniques, is covered. The Phase 2 test program evaluated the hot gas ingestion phenomena at higher temperatures and used a laser sheet to investigate the flow field. Hot gas ingestion levels were measured for the several forward nozzle splay configurations and with flow control/life improvement devices (LIDs) which reduced the hot gas ingestion. The test was conducted at full scale nozzle pressure ratios and inlet Mach numbers. Results are presented over a range of nozzle pressure ratios at a 10 kn headwind velocity. The Phase 2 program was conducted at exhaust nozzle temperatures up to 1460 R and utilized a sheet laser system for flow visualization of the model flow field in and out of ground effects. The results reported are for nozzle exhaust temperatures up to 1160 R and contain the compressor face pressure and temperature distortions, the total pressure recovery, the inlet temperature rise, and the environmental effects of the hot gas. The environmental effects include the ground plane contours, the model airframe heating, and the location of the ground flow separation.

Johns, Albert L.↗
NASA NTRS1991

Titan improvement study - Hydrogen core stages

The performance benefits and comparative cost-effectiveness of Titan/Centaur launcher improvement options involving a new LH2/LOX second stage and/or a new LH2/LOX first stage are presented. This incremental improvement approach involves higher confidence due to high components-commonality with well-tested systems. The proposed second stage's capabilities are such as to allow it to serve as an OTV. This improvement to Titan constitutes one of several justifications for a new 500-1000 kN thrust liquid-fuel engine.

Martin, James A.↗
NASA NTRS1990

ALS turbomachinery technology

Advanced Development Programs are being pursued by Rocketdyne, Aerojet, and Pratt and Whitney to define and validate design approaches toward producing low-cost, reliable liquid-hydrogen and liquid-oxygen turbopumps for a 2580 kN (580 klb) thrust Advanced Launch System. The generic approach, which is evolving after 18 months of trade studies and conceptual and preliminary design efforts, is explained. In addition, the preliminary liquid-hydrogen turbopump designs produced in parallel tasks by Rocketdyne and Aerojet and the liquid-oxygen turbopump design produced by Pratt and Whitney are described, and technology features and issues are discussed.

Csomor, A.↗
NASA NTRS1991

Heat transfer analysis of fuel assemblies in a heterogeneous gas core nuclear rocket

Heat transfer problems of a heterogeneous gaseous core nuclear rocket were studied. The reactor core consists of 1.5-m long hexagonal fuel assemblies filled with pressurized uranium tetrafluoride (UF4) gas. The fuel gas temperature ranges from 3500 to 7000 K at a nominal operating condition of 40 atm. Each fuel assembly has seven coolant tubes, through which hydrogen propellant flows. The propellant temperature is not constrained by the fuel temperature but by the maximum temperature of the graphite coolant tube. For a core achieving a fission power density of 1000 MW/cu m, the propellant core exit temperature can be as high as 3200 K. The physical size of a 1250 MW gaseous core nuclear rocket is comparable with that of a NERVA-type solid core nuclear rocket. The engine can deliver a specific impulse of 1020 seconds and a thrust of 330 kN.

Watanabe, Yoichi↗
NASA NTRS1992

Effect of model selection on combustor performance and stability predictions using ROCCID

The ROCket Combustor Interactive Design (ROCCID) methodology is an interactive computer program that combines previously developed combustion analysis models to calculate the combustion performance and stability of liquid rocket engines. Test data from 213 kN (48,000 lbf) Liquid Oxygen (LOX)/RP-1 combustor with an O-F-O (oxidizer-fuel-oxidizer) triplet injector were used to characterize the predictive capabilities of the ROCCID analysis models for this injector/propellant configuration. Thirteen combustion performance and stability models were incorporated into ROCCID, and ten of them, which have options for triplet injectors, were examined. Calculations using different combinations of analysis models, with little or no anchoring, were carried out on a test matrix of operating combinations matching those of the test program. Results of the computer analyses were compared to test data, and the ability of the model combinations to correctly predict combustion stability or instability was determined. For the best model combination(s), sensitivity of the calculations to fuel drop size and mixing efficiency was examined. Error in the stability calculations due to uncertainty in the pressure interaction index (N) was examined. The recommended model combinations for this O-F-O triplet LOX/RP-1 configuration are proposed.

Giuliani, James E.↗