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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 541 records · Page 30
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Ares I-X Upper Stage Simulator Structural Analyses Supporting the NESC Critical Initial Flaw Size Assessment

The structural analyses described in the present report were performed in support of the NASA Engineering and Safety Center (NESC) Critical Initial Flaw Size (CIFS) assessment for the ARES I-X Upper Stage Simulator (USS) common shell segment. The structural analysis effort for the NESC assessment had three thrusts: shell buckling analyses, detailed stress analyses of the single-bolt joint test; and stress analyses of two-segment 10 degree-wedge models for the peak axial tensile running load. Elasto-plastic, large-deformation simulations were performed. Stress analysis results indicated that the stress levels were well below the material yield stress for the bounding axial tensile design load. This report also summarizes the analyses and results from parametric studies on modeling the shell-to-gusset weld, flange-surface mismatch, bolt preload, and washer-bearing-surface modeling. These analyses models were used to generate the stress levels specified for the fatigue crack growth assessment using the design load with a factor of safety.

Knight, Norman F., Jr.↗
NASA NTRS2009

Short Duration Bioastronautics Investigation 1904: Human Factors Assessment of Vibration Effects on Visual Performance during Launch

The primary objective of the Short Duration Bioastronautics Investigation (SDBI) 1904 was to determine visual performance limits during Shuttle operational vibration and g-loads, specifically through the determination of minimal usable font sizes using Orion-type display formats. Currently there is little to no data available to quantify human visual performance under the extreme g- and vibration conditions of launch. Existing data on shuttle vibration magnitude and frequency is incomplete and does not address human visual performance. There have been anecdotal reports of performance decrements from shuttle crews, but no structured data have been collected. Previous work by NASA on the effects of vibration and linear g-loads on human performance was conducted during the Gemini era, but these experiments were performed using displays and controls that are dramatically different than current concepts being considered by the Constellation Program. Recently, three investigations of visual performance under vibration have been completed at NASA Ames Research Center: the first examining whole-body vibration, the second employing whole-body vibration coupled with a sustained g-load, and a third examining the effects of peak versus extended duration vibration. However, all of these studies were conducted using only a single x-axis direction (eyeballs in/out). Estimates of thrust oscillations from the Constellation Ares-I first stage are driving the need for realistic human performance requirements. SDBI 1904 was an opportunity to address the need for requirements by conducting a highly focused and applied evaluation in a relevant spaceflight environment. The SDBI was a companion effort to Detailed Test Objective (DTO) 695, which measured shuttle seat accelerations (vibration) during ascent. Data from the SDBI will serve an important role in interpreting the DTO vibration data. Both SDBI 1904 and DTO 695 were low impact with respect to flight resources, and combined, they represent an efficient and focused problem solving approach. This project provided (a) immediate data for developing preliminary human performance vibration requirements; (b) flight validated inputs for ongoing and future ground-based research; and (c) preliminary information related to Orion display format design.

Thompson, Shelby↗
NASA NTRS2001

Integrated Pulse Detonation Propulsion and Magnetohydrodynamic Power

The prospects for realizing an integrated pulse detonation propulsion and magnetohydrodynamic (MHD) power system are examined. First, energy requirements for direct detonation initiation of various fuel-oxygen and fuel-air mixtures are deduced from available experimental data and theoretical models. Second, the pumping power requirements for effective chamber scavenging are examined through the introduction of a scavenging ratio parameter and a scavenging efficiency parameter. A series of laboratory experiments were carried out to investigate the basic engineering performance characteristics of a pulse detonation-driven MHD electric power generator. In these experiments, stoichiometric oxy-acetylene mixtures seeded with a cesium hydroxide/methanol spray were detonated at atmospheric pressure in a 1-m-long tube having an i.d. of 2.54 cm. Experiments with a plasma diagnostic channel attached to the end of the tube confirmed the attainment of detonation conditions (p2/p1 approximately 34 and D approximately 2,400 m/sec) and enabled the direct measurement of current density and electrical conductivity (approximately = 6 S/m) behind the detonation wave front, In a second set of experiments, a 30-cm-long continuous electrode Faraday channel, having a height of 2.54 cm and a width of 2 cm, was attached to the end of the tube using an area transition duct. The Faraday channel was inserted in applied magnetic fields of 0.6 and 0.95 T, and the electrodes were connected to an active loading circuit to characterize power extraction dependence on load impedance while also simulating higher effective magnetic induction. The experiments indicated peak power extraction at a load impedance between 5 and 10 Omega. The measured power density was in reasonable agreement with a simple electrodynamic model incorporating a correction for near-electrode potential losses. The time-resolved thrust characteristics of the system were also measured, and it was found that the NM interaction exerted a negligible influence on system thrust and that the measured I(sub sp) of the system (200 see) exceeded that computed for an equivalent nozzleless rocket (120 see).

Litchford, Ron J.↗
NASA NTRS2001

Integrated Pulse Detonation Propulsion and Magnetohydrodynamic Power

The prospects for realizing an integrated pulse detonation propulsion and magnetohydrodynamic (MHD) power system are examined. First, energy requirements for direct detonation initiation of various fuel-oxygen and fuel-air mixtures are deduced from available experimental data and theoretical models. Second, the pumping power requirements for effective chamber scavenging are examined through the introduction of a scavenging ratio parameter and a scavenging efficiency parameter. A series of laboratory experiments were carried out to investigate the basic engineering performance characteristics of a pulse detonation-driven MHD electric power generator. In these experiments, stoichiometric oxy-acetylene mixtures seeded with a cesium hydroxide/methanol spray were detonated at atmospheric pressure in a 1-m-long tube having an i.d. of 2.54 cm. Experiments with a plasma diagnostic channel attached to the end of the tube confirmed the attainment of detonation conditions (p(sub 2)/p(sub 1) approx. 34 and D approx. 2,400 m/sec) and enabled the direct measurement of current density and electrical conductivity (=6 S/m) behind the detonation wave front. In a second set of experiments, a 30-cm-long continuous electrode Faraday channel, having a height of 2.54 cm and a width of 2 cm, was attached to the end of the tube using an area transition duct. The Faraday channel was inserted in applied magnetic fields of 0.6 and 0.95 T. and the electrodes were connected to an active loading circuit to characterize power extraction dependence on load impedance while also simulating higher effective magnetic induction. The experiments indicated peak power extraction at a load impedance between 5 and 10 Ohm. The measured power density was in reasonable agreement with a simple electrodynamic model incorporating a correction for near-electrode potential losses. The time-resolved thrust characteristics of the system were also measured, and it was found that the MHD interaction exerted a negligible influence on system thrust and that the measured I(sub sp) of the system (200 sec) exceeded that computed for an equivalent nozzleless rocket (120 sec).

Litchford, R. J.↗
NASA NTRS2016

Inviscid and Viscous CFD Analysis of Booster Separation for the Space Launch System Vehicle

This paper presents details of Computational Fluid Dynamic (CFD) simulations of the Space Launch System during solid-rocket booster separation using the Cart3D inviscid and Overflow viscous CFD codes. The discussion addresses the use of multiple data sources of computational aerodynamics, experimental aerodynamics, and trajectory simulations for this critical phase of flight. Comparisons are shown between Cart3D simulations and a wind tunnel test performed at NASA Langley Research Center's Unitary Plan Wind Tunnel, and further comparisons are shown between Cart3D and viscous Overflow solutions for the flight vehicle. The Space Launch System (SLS) is a new exploration-class launch vehicle currently in development that includes two Solid Rocket Boosters (SRBs) modified from Space Shuttle hardware. These SRBs must separate from the SLS core during a phase of flight where aerodynamic loads are nontrivial. The main challenges for creating a separation aerodynamic database are the large number of independent variables (including orientation of the core, relative position and orientation of the boosters, and rocket thrust levels) and the complex flow caused by exhaust plumes of the booster separation motors (BSMs), which are small rockets designed to push the boosters away from the core by firing partially in the direction opposite to the motion of the vehicle.

Viscous↗
NASA NTRS2016

Preliminary Results From a Heavily Instrumented Engine Ice Crystal Icing Test in a Ground Based Altitude Test Facility

Preliminary results from the Heavily Instrumented ALF503R-5 Engine test conducted in the NASA Glenn Research Center Propulsion Systems Laboratory will be discussed. The effects of ice crystal icing on a full scale engine is examined and documented. This model engine, serial number LF01, was used during the inaugural icing test in the PSL facility. The reduction of thrust (rollback) events experienced by this engine in flight were replicated in the facility. Limited instrumentation was used to detect icing. Metal temperature on the exit guide vanes and outer shroud and the load measurement were the only indicators of ice formation. The current study features a similar engine, serial number LF11, which is instrumented to characterize the cloud entering the engine, detect characterize ice accretion, and visualize the ice accretion in the region of interest.

enigine icing↗
NASA NTRSDate not supplied

Ascent Aborts 2 Flight, Reconstruction of Structural Dynamics (Couple Loads Transients)

- This work describes the elements of AA2 post-flight data analysis, couple loads analysis, and comparisons of the flight data with the results of analytical methods transiently. - The measured transient responses ranged from rigid body frequencies (zero Hz) to 150 Hz. - The analytical methods consisted of defining and determining all forcing functions, i.e., buffet, aerodynamics, thrusts from the abort motor and the carrier lunch vehicle motor, and the initial conditions at pre-separation. - Mathematical models representing the mechanics of Orion abort is relatively complex because 1) two motors (carrier’s and LAS’s) are operating at the same time at pre-separation, and 2) the separation occurs in atmosphere. - Other challenges consist of determining the aerodynamics with plume-on condition at every millisecond of the flight ranging from transonic to supersonic (i.e., Mach 1.1 to Mach 1.6), and defining the mechanics of separation of LAS from the launch vehicle.

Launch Vehicle↗
NASA NTRS1975

Rotorcraft low-speed download drag definition and its reduction

Download drag for rotorcraft in hover and low-speed flight is a burden which significantly affects useful load, fuel, and payload. Reduction of the burden will enhance these aspects of rotorcraft and complement the forthcoming improvements in isolated rotor performance. Analyses and experimental data are available, though fragmentary, regarding gross drag, thrust recovery, and other characteristics which can be utilized to define interim rotorcraft design changes to reduce the burden. Eventually the experimental data and a comprehensive combination of rotor, rotor-wake, and potential-flow analyses can evolve to reduce the burden to an absolute minimum.

Wilson, J. C.↗
NASA NTRS1976

Design and evaluation of a 3 million DN series-hybrid thrust bearing

The design and experimental evaluation of a series-hybrid thrust bearing, consisting of a 150-mm ball bearing and a centrifugally actuated, conical, fluid-film bearing, is presented. Tests were conducted up to 16,000 rpm and at this speed an axial load of 15,600 N (3500 lb) was safely supported by the hybrid bearing system. Through the series-hybrid bearing principle, the effective ball bearing speed was reduced to approximately one-half of the shaft speed. A speed reduction of this magnitude would result in a tenfold increase in the ball bearing fatigue life. A successful evaluation of fluid-film bearing lubricant supply failure was performed repeatedly at an operating speed of 10,000 rpm. A complete and smooth changeover to full-scale ball bearing operation was effected when the oil supply to the fluid-film bearing was cut off. Reactivation of the fluid-film oil supply system produced a flawless return to the original mode of hybrid operation.

Scibbe, H. W.↗
NASA NTRS1983

A design study of a reaction control system for a V/STOL fighter/attack aircraft

Attention is given to a short takeoff vertical landing (STOVL) aircraft reaction control system (RCS) design study. The STOVL fighter/attack aircraft employs an existing turbofan engine, and its hover requirement places a premium on weight reduction, which eliminates prospective nonairbreathing RCSs. A simple engine compressor bleed RCS degrades overall performance to an unacceptable degree, and the supersonic requirement precludes the large volume alternatives of thermal or ejector thrust augmentation systems as well as the ducting of engine exhaust gases and the use of a dedicated turbojet. The only system which addressed performance criteria without requiring major engine modifications was a dedicated load compressor driven by an auxilliary power unit.

Beard, B. B.↗
NASA NTRS1984

Space Power Management and Distribution Status and Trends

An overview of space power management and distribution (PMAD) is provided which encompasses historical and current technology trends. The PMAD components discussed include power source control, energy storage control, and load power processing electronic equipment. The status of distribution equipment comprised of rotary joints and power switchgear is evaluated based on power level trends in the public, military, and commercial sectors. Component level technology thrusts, as driven by perceived system level trends, are compared to technology status of piece-parts such as power semiconductors, capacitors, and magnetics to determine critical barriers.

Reppucci, G. M.↗
NASA NTRS1986

Combustor diffuser interaction program

Advances in gas turbine engine performance are achieved by using compressor systems with high stage loading and low part count, which result in high exit Mach numbers. The diffuser and combustor systems in such engines should be optimized to reduce system pressure loss and to maximize the engine thrust-to-weight ratio and minimize length. The state-of-the-art combustor-diffuser systems do not meet these requirements. Detailed understanding of the combustor-diffuser flow field interaction is required for designing advanced gas turbine engines. An experimental study of the combustor-diffuser interaction (CDI) is being conducted to obtain data for the evaluation and improvement of analytical models applicable to a wide variety of diffuser designs. The CDI program consists of four technical phases: Literature Search; Baseline Configuration; Parametric Configurations; and Performance Configurations. Phase 2 of the program is in progress.

Srinivasan, Ram↗
NASA NTRS1994

Metallized gelled propellants: Oxygen/RP-1/aluminum combustion experiments

A series of combustion experiments using metallized gelled liquid propellants were conducted. These experiments used a small 30- to 40-lb(sub f) thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-wt percent, 5-wt percent, and 44-wt percent loadings of aluminum and gaseous oxygen and the oxidizer. Ten different injectors were used during the testing: 6 were for the baseline O2/RP-1 tests and 4 for the gelled fuels. Relatively high C-star efficiencies were obtained with gelled RP-1 (0-wt% RP-1/Al) and metallized 5-wt% RP-1/Al over the O/F range tested: 90-98%. A peak of 98 percent efficiency was reached with ungelled O2/RP-1 and up to 95% efficiency was obtained with gelled RP-1/Al (55-wt% Al). Injector erosion was evident with the 55-wt% testing, while there was little or no erosion seen with the gelled RP-1 with 0 and 5-wt% Al. A protective layer of gelled fuel formed in the firings that minimized the damage to the rocket injector face. This effect may provide a useful technique for engine cooling.

Palaszewski, Bryan↗
NASA NTRS2000

Wave Fluid Film Bearing Tests for an Aviation Gearbox

An oil-lubricated wave journal-thrust bearing assembly was successfully tested at conditions found in general aviation engine gearboxes. The bearing performed well at both steady state conditions and in start-stop tests. It ran stably under all loading conditions, including zero load, at all speeds up to 16 000 rpm. The bearing carried 25 percent more load than required for the gearbox application, supporting 8900 N (94 bars average pressure), and showed very good thermal stability. 450 start-stop cycles were also performed, including 350 cycles without oil supply during starting and stopping. Test results and numerical predictions were in good agreement.

Dimofte, Florin↗
NASA NTRS2007

Crew Exploration Vehicle Ascent Abort Trajectory Analysis and Optimization

The Orion Crew Exploration Vehicle is the first crewed capsule design to be developed by NASA since Project Apollo. Unlike Apollo, however, the CEV is being designed for service in both Lunar and International Space Station missions. Ascent aborts pose some issues that were not present for Apollo, due to its launch azimuth, nor Space Shuttle, due to its cross range capability. The requirement that a North Atlantic splashdown following an abort be avoidable, in conjunction with the requirement for overlapping abort modes to maximize crew survivability, drives the thrust level of the service module main engine. This paper summarizes 3DOF analysis conducted by NASA to aid in the determination of the appropriate propulsion system for the service module, and the appropriate propellant loading for ISS missions such that crew survivability is maximized.

Falck, Robert D.↗
NASA NTRS1974

Cyclic fatigue analysis of rocket thrust chambers. Volume 1: OFHC copper chamber low cycle fatigue

A three-dimensional finite element elasto-plastic strain analysis was performed for the throat section of a regeneratively cooled rocket combustion chamber. The analysis employed the RETSCP finite element computer program. The analysis included thermal and pressure loads, and the effects of temperature dependent material properties, to determine the strain range corresponding to the chamber operating cycle. The analysis was performed for chamber configuration and operating conditions corresponding to a hydrogen-oxygen combustion chamber which was fatigue tested to failure. The computed strain range at typical chamber operating conditions was used in conjunction with oxygen-free, high-conductivity (OHFC) copper isothermal fatigue test data to predict chamber low-cycle fatigue life.

Miller, R. W.↗
NASA NTRS1977

Load regulating latch

A load regulating mechanical latch is described that has a pivotally mounted latch element having a hook-shaped end with a strike roller-engaging laterally open hook for engaging a stationary strike roller. The latch element or hook is pivotally mounted in a clevis end of an elongated latch stem that is adapted for axial movement through an opening in a support plate or bracket mounted to a structural member. A coil spring is disposed over and around the extending latch stem and the lower end of the coil spring engages the support bracket. A thrust washer is removably attached to the other end of the latch stem and engages the other end of the coil spring and compresses the coil spring thereby preloading the spring and the latch element carried by the latch stem. The hook-shaped latch element has a limited degree of axial travel for loading caused by structural distortion which may change the relative positions of the latch element hook and the strike roller. Means are also provided to permit limited tilt of the latch element due to loading of the hook.

Appleberry, W. T.↗
NASA NTRS1976

Results of boron-aluminum thrust structure

Results are presented of testing-to-failure a two member boron-aluminum thrust structure. The structure represented one section of a more complex planar truss and was designed to test the integrity of a diffusion bonded joint. The structure failed at 107 percent of the ultimate design load in the diffusion bond region. Strain gages and displacement transducers were used to measure loads and deflections of the truss. The experimentally derived axial loads, bending moments, and torsion in the various members are presented and compared with predicted values.

Hyer, M. W.↗