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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 1,405 records · Page 78
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

Solar Thermal Upper Stage Liquid Hydrogen Pressure Control Testing

High-energy cryogenic propellant is an essential element in future space exploration programs. Therefore, NASA and its industrial partners are committed to an advanced development/technology program that will broaden the experience base for the entire cryogenic fluid management community. Furthermore, the high cost of microgravity experiments has motivated NASA to establish government/aerospace industry teams to aggressively explore combinations of ground testing and analytical modeling to the greatest extent possible, thereby benefitting both industry and government entities. One such team consisting of ManTech SRS, Inc., Edwards Air Force Base, and Marshall Space Flight Center (MSFC) was formed to pursue a technology project designed to demonstrate technology readiness for an SRS liquid hydrogen (LH2) in-space propellant management concept. The subject testing was cooperatively performed June 21-30, 2000, through a partially reimbursable Space Act Agreement between SRS, MSFC, and the Air Force Research Laboratory. The joint statement of work used to guide the technical activity is presented in appendix A. The key elements of the SRS concept consisted of an LH2 storage and supply system that used all of the vented H2 for solar engine thrusting, accommodated pressure control without a thermodynamic vent system (TVS), and minimized or eliminated the need for a capillary liquid acquisition device (LAD). The strategy was to balance the LH2 storage tank pressure control requirements with the engine thrusting requirements to selectively provide either liquid or vapor H2 at a controlled rate to a solar thermal engine in the low-gravity environment of space operations. The overall test objective was to verify that the proposed concept could enable simultaneous control of LH2 tank pressure and feed system flow to the thruster without necessitating a TVS and a capillary LAD. The primary program objectives were designed to demonstrate technology readiness of the SRS concept at a system level as a first step toward actual flight vehicle demonstrations. More specific objectives included testing the pressure and feed control system concept hardware for functionality, operability, and performance. Valuable LH2 thermodynamic and fluid dynamics data were obtained for application to both the SRS concept and to future missions requiring space-based cryogen propellant management.

Moore, J. D.↗
NASA NTRS1992

Pressure fed thrust chamber technology program

This is the final report for the Pressure Fed Technology Program. It details the design, fabrication and testing of subscale hardware which successfully characterized LOX/RP combustion for a low cost pressure fed design. The innovative modular injector design is described in detail as well as hot-fire test results which showed excellent performance. The program summary identifies critical LOX/RP design issues that have been resolved by this testing, and details the low risk development requirements for a low cost engine for future Expendable Launch Vehicles (ELVi).

Dunn, Glenn M.↗
DOE OSTI2022

Dynamic Rupture Modeling of the 1999 Chi-Chi, Taiwan Earthquake Using 3DFinite Element Method. Sensitivity Analysis of Slip Rate Function to Model Parameters of Crustal Weak-Zone

The main objective of this study was the dynamic rupture modeling of the M7.3 1999 Chi-Chi, Taiwan earthquake. The purpose of the numerical modeling was two folded. First, using models of rupture dynamics, derive kinematic rupture characteristics, including spatial and temporal variations of the slip rate time history and rupture speed, in the long-period ground motion generation area (LMGA), strong motion generation areas (SMGAs) and background fault area for a thrust shallow rupture. Second, investigate the sensitivity of source time function characteristics in the LMGA and SMGA areas to model parameterization of low-velocity weak zone, such as thickness and shear wave velocity.

58 GEOSCIENCES↗
NASA NTRS1985

Experimental performance of a 1-kilowatt arcjet thruster

A formerly unused cathode and anode/nozzle assembly from a flight model arcjet was tested with nitrogen, hydrogen, and nitrogen-hydrogen mixture simulating ammonia decomposition products at arc power levels from about 300 to 950 W. Two different power sources and two nozzle configurations were tested at low back-ground pressures to exclude facility effects. Increased nozzle expansion ratio improved cold flow nozzle efficiency from 0.8 to 0.9. Hydrogen thrust efficiency of 0.26 at 872 sec specific impulse matched some 1964 performance on a similar device. Simulated ammonia thrust efficiency was 0.31 at 422 sec. Spontaneously occurring voltage mode changes at constant arc current could be partially stabilized with appropriate power source characteristics. In the higher voltage mode specific impulse was higher, but thrust efficiency changed only slightly from that of the lower voltage mode. Sustained tests of up to 2 hr duration exhibited no apparent performance degradation with time.

Nakanishi, S.↗
NASA NTRS1985

Effects of velocity profile and inclination on dual-jet-induced pressures on a flat plate in a crosswind

An experimental study was conducted to determine surface pressure distributions on a flat plate with dual subsonic, circular jets exhausting from the surface into a crossflow. The jets were arranged in both side-by-side and tandem configurations and were injected at 90 deg and 60 deg angles to the plate, with jet-to-crossflow velocity ratio of 2.2 and 4. The major objective of the study was to determine the effect of a nonuniform (vs uniform) jet velocity profile, simulating the exhaust of a turbo-fan engine. Nonuniform jets with a high-velocity outer annulus and a low-velocity core induced stronger negative pressure fields than uniform jets with the same mass flow rate. However, nondimensional lift losses (lift loss/jet thrust lift) due to such nonuniform jets were lower than lift losses due to uniform jets. Changing the injection angle from 90 deg to 60 deg resulted in moderate (for tandem jets) to significant (for side-by-side jets) increases in the induced negative pressures, even though the surface area influenced by the jets tended to reduce as the angle decreased. Jets arranged in the side-by-side configuration led to significant jet-induced lift losses exceeding, in some cases, lift losses reported for single jets.

Jakubowski, A. L.↗
NASA NTRS1953

Wind-Tunnel Investigation at Low Speed of the Pitching Stability Derivatives of a 1/9-Scale Powered Model of the Convair XFY-1 Vertically Rising Airplane, TED No. NACA DE 373

An experimental investigation has been conducted in the Langley stability tunnel at low speed to determine the pitching stability derivatives of a 1/9-scale powered model of the Convair XFY-1 vertically rising airplane. Effects of thrust coefficient, control deflections, and propeller blade angle were investigated. The tests were made through an angle-of-attack range from about -4deg to 29deg, and the thrust coefficient range was from 0 to 0.7. In order to expedite distribution of these data, no analysis of the data has been prepared for this paper.

Queijo, M. J.↗
NASA NTRS1953

Wind-Tunnel Investigation at Low Speed of the Rolling Stability Derivatives of a 1/9-Scale Powered Model of the Convair XFY-1 Vertically Rising Airplane, TED No. NACA DE 373

An experimental investigation has been conducted in the Langley stability tunnel at low speed to determine the rolling stability derivatives of a 1/9-scale powered model of the Convair XFY-1 vertically rising airplane. Effects of thrust coefficient were investigated for the complete model and for certain components of the model. Effects of control deflections and of propeller blade angle were investigated for the complete model. Most of the tests were made through an angle-of-attack range from about -4deg to 29deg, and the thrust coefficient range was from 0 to 0.7. In order to expedite distribution of these data, no analysis of the data has been prepared for this paper.

Queijo, M. J.↗
NASA NTRS2015

Oxide_Oxide Ceramic Matrix Composite (CMC) Exhaust Mixer Development in the NASA Environmentally Responsible Aviation (ERA) Project

Rolls-Royce North American Technologies, Inc. (LibertyWorksLW) began considering the development of CMC exhaust forced mixers in 2008, as a means of obtaining reduced weight and hotter operating temperature capability, while minimizing shape distortion during operation, which would improve mixing efficiency and reduce fuel burn. Increased component durability, enhanced ability to fabricate complex-shaped components, and engine noise reduction are other potential advantages of CMC mixers (compared to metallic mixers). In 2010, NASA was pursuing the reduction of NOx emissions, fuel burn, and noise from turbine engines in Phase I of the Environmentally Responsible Aviation (ERA) Project. ERA subtasks, including those focused on CMC components, were formulated with the goal of maturing technology from proof of concept validation (TRL 3) to a systemsubsystem or prototype demonstration in a relevant environment (TRL 6). In April 2010, the NASA Glenn Research Center (GRC) and LibertyWorks jointly initiated a CMC Exhaust System Validation Program within the ERA Project, teaming on CMC exhaust mixer development for subsonic jet engines capable of operating with increased performance. Our initial focus was on designing, fabricating, and characterizing the thrust and acoustic performance of a roughly quarter-scale 16-lobe oxide oxide CMC mixer and tail cone along with a conventional low bypass exhaust nozzle. Support Services, LLC (Allendale, MI) and ATK COI Ceramics, Inc. (COIC, in San Diego, CA) supported the design of a subscale nozzle assembly that consisted of an oxide oxide CMC mixer and center body, with each component mounted on a metallic attachment ring. That design was based upon the operating conditions a mixer would experience in a turbofan engine. Validation of the aerodynamic and acoustic performance of the subscale mixer via testing and the achievement of TRL 4 encouraged the NASALWCOIC team to move to the next phase where a full scale CMC mixer sized for a RR AE3007 engine and a compatible attachment flange were designed, followed by CMC component fabrication by COIC, and vibration testing at GRC under conditions simulating the structural and dynamic environment encountered during engine operation. AFRL (WPAFB) supported this testing by performing 3D laser vibrometry to identify the mixer mode shapes and modal frequencies. The successful fabrication and testing of such a component has been achieved. The CMC mixer demonstrated good durability during vibration testing at room and elevated temperature (TRL5). This has cleared the article for a ground-based test on a Rolls-Royce AE3007 engine, where the performance and benefits of the component can be further assessed.

exhaust nozzles↗
NASA NTRS2017

Advanced Noise Abatement Procedures for a Supersonic Business Jet

Supersonic civil aircraft present a unique noise certification challenge. High specific thrust required for supersonic cruise results in high engine exhaust velocity and high levels of jet noise during takeoff. Aerodynamics of thin, low-aspect-ratio wings equipped with relatively simple flap systems deepen the challenge. Advanced noise abatement procedures have been proposed for supersonic aircraft. These procedures promise to reduce airport noise, but they may require departures from normal reference procedures defined in noise regulations. The subject of this report is a takeoff performance and noise assessment of a notional supersonic business jet. Analytical models of an airframe and a supersonic engine derived from a contemporary subsonic turbofan core are developed. These models are used to predict takeoff trajectories and noise. Results indicate advanced noise abatement takeoff procedures are helpful in reducing noise along lateral sidelines.

Berton, Jeffrey J.↗
NASA NTRSDate not supplied

Study of a Plug Nozzle for Supersonic Aircraft Concepts

A plug nozzle with a cylindrical cowl is studied. The geometry is adopted following commercial supersonic aircraft concepts considered in an ongoing NASA program. For the given cowl, the plug geometry is varied that includes shapes developed by a numerical optimization study as well as porous plugs. Experimental data on noise and flow field are discussed with an eye for minimum noise at low nozzle pressure ratios (NPR2) representing landing and takeoff (LTO) conditions. NPR up to about 4.7 is covered in the experiments. Limited thrust data with accompanying numerical simulation results are also presented. Results show that a plug designed for optimum thrust at cruise is noisier while a plug designed for optimum thrust at LTO has less noise, throughout the NPR range covered. The results also show that a plug nozzle can generate transonic tones and excess broadband noise (EBBN) at lower NPR, apart from well-known screech tones and broadband shock associated noise (BBSN) at higher NPR. A long porous plug is found to effectively suppress all these aberrant noise components while suffering less than 1% loss in thrust coefficient.

Jets↗
NASA NTRSDate not supplied

Study of a Plug Nozzle for Supersonic Aircraft Concepts

A plug nozzle with a cylindrical cowl is studied. The geometry is adopted following commercial supersonic aircraft concepts considered in an ongoing NASA program. For the given cowl, the plug geometry is varied that includes shapes developed by a numerical optimization study as well as porous plugs. Experimental data on noise and flow field are discussed with an eye for minimum noise at low nozzle pressure ratios (NPR2) representing landing and takeoff (LTO) conditions. NPR up to about 4.7 is covered in the experiments. Limited thrust data with accompanying numerical simulation results are also presented. Results show that a plug designed for optimum thrust at cruise is noisier while a plug designed for optimum thrust at LTO has less noise, throughout the NPR range covered. The results also show that a plug nozzle can generate transonic tones and excess broadband noise (EBBN) at lower NPR, apart from well-known screech tones and broadband shock associated noise (BBSN) at higher NPR. A long porous plug is found to effectively suppress all these aberrant noise components while suffering less than 1% loss in thrust coefficient.

Jets↗
NASA NTRS2023

Propulsion-Airframe Integration for Conceptual Redesign of a Low-Boom Supersonic Transport

A low-boom supersonic transport was generated in a previous low-boom multidisciplinary optimization (MDO) study that used computational fluid dynamics (CFD) off-body pressure to compute the undertrack sonic boom ground signature and calibrated low-fidelity aerodynamics analyses to compute the mission performance metrics. This low-boom aircraft, referred to as the Mach 1.7 40-PAX concept, can carry 40 passengers for a low-boom overland mission with cruise Mach 1.7 for the airport pairs over the continental US and has the potential to achieve an undertrack sonic boom ground noise level below 70 perceived level of decibels at the start of overland cruise (SOC). The engine for the Mach 1.7 40-PAX concept was designed using the Numerical Propulsion System Simulation (NPSS) for the mission analysis and modeled as a flow-through nacelle for CFD-based sonic boom analysis. To understand how the engine plume affects the undertrack ground signature, the Mach 1.7 40-PAX concept is redesigned in this paper after replacing the flow-through nacelles by CFD engines for sonic boom analysis using aeropropulsive CFD simulation. A process is developed for approximation of the NPSS engine at SOC by a CFD engine for aeropropulsive CFD simulation. The generated CFD engine has the identical nozzle boundary conditions and approximately the same mass flow and thrust as those of the NPSS engine at SOC. Then, the outer mold line of the configuration with the CFD engines is redesigned to approximately restore the low-boom characteristics of the Mach 1.7 40-PAX concept at SOC. Finally, the CFD simulation data for the redesigned concept with the CFD engines is used to calibrate the low-fidelity aerodynamic analyses for mission analysis of this concept. The cyclic dependency of the involved disciplinary analyses is resolved using an iteration method for a consistent coupling of the mission analysis, NPSS engine analysis, and low-boom redesign using the aeropropulsive CFD simulation. This low-boom redesign study is used as an example to demonstrate how the propulsion-airframe integration could be implemented for conceptual design of supersonic transports that satisfy both the low-boom and mission performance requirements.

Low-boom supersonic transport↗
NASA NTRS1993

Thrust augmentation options for the Beta 2 two-stage-to-orbit vehicle

NASA LeRC is continuing to study propulsion concepts for a horizontal takeoff and landing, fully reusable, two-stage-to-orbit vehicle. This will be capable of launching and returning a 10,000 pound payload to a 100 nautical mile polar orbit using low-risk technology. The vehicle, Beta 2, is a derivative of the USAF/Boeing Beta vehicle which was designed to deliver a 50,000 pound payload to a similar orbit. Beta 2 stages at Mach 6.5 and about 100,000 ft altitude. The propulsion system for the booster is an over/under turbine bypass engine/ramjet configuration. In this paper, several options for thrust augmentation were studied in order to improve the performance of this engine where there was a critical need. Options studies were turbine engine overspeed in the transonic region, water injection at a various turbine engine locations also during the transonic region, and water injection at the turbine engine face during high speed operation. The methodology, constraints, propulsion performance, and mission study results are presented.

Snyder, Christopher A.↗
NASA NTRS1981

Transfer film evaluation for shuttle engine turbopump bearing

A series of low speed experiments to evaluate the possible occurrence of transfer film lubrication and the effectiveness of burnished films in the shuttle spacecraft main engine thrust bearings were conducted. No evidence of transfer film lubrication was evident, although this could have been the result of the (used) condition of the bearing. Burnished films of either Teflon or Rulon were found to greatly enhance the performance of the bearing. Crush load experiments indicated that the bearing ultimate load capability is on the order of 489,000 N (110,000 pounds). The effect of ball (as well as race) burnishing techniques on bearing performance, different types of burnished films, and transfer film formation are suggested for further study.

Kannel, J. W.↗
NASA NTRS1988

Thrust chamber thermal barrier coating techniques

Methods for applying thermal barrier coatings to the hot-gas side wall of rocket thrust chambers in order to significantly reduce the heat transfer in high heat flux regions was the focus of technology efforts for many years. This paper describes a successful technique developed by the Lewis Research Center that starts with the coating of a mandrel and then builds the thrust chamber around it by electroforming appropriate materials. This results in a smooth coating with exceptional adherence, demonstrated in hot fire rig tests. The low cycle fatigue life of chambers with coatings applied in this manner was increased dramatically compared to uncoated chambers.

Quentmeyer, Richard J.↗
NASA NTRS1989

Hybrid propulsion technology program: Phase 1, volume 1

A number of booster propulsion system concepts are being considered for the next generation of manned and unmanned space launch vehicles. The one propulsion system concept that has potential for reducing costs with increased safety, reliability, and performance is hybrid propulsion (HP). A HP system may be thought of as a liquid propulsion system with solid fuel or a solid propulsion system with a liquid oxidizer. The liquid propulsion features that are most attractive are the higher specific impulse, clean exhaust, separated propellants, and oxidizer loading just prior to launch. The most attractive solid propulsion features includes low life cycle costs, no rotating machinery, compact size, and a robust case. In addition, a HP system has a robust LO2 tank; provides thrust control for ignition, to alleviate flight loads, and for thrust termination; and uses an inert grain that is not sensitive to anomalies such as cracks, voids, and separations. The object is to develop the technology to enable the application of HP to manned and unmanned space launch vehicles. This program will identify the necessary technology, acquire that technology, and demonstrate that technology. This volume is the executive summary.

Schuler, A. L.↗
NASA NTRS1990

Ideal efficiency of propellers - Theodorsen revisited

Theodorsen's 1948 analog evaluation of the parameters governing the ideal (friction-free) efficiency of propellers is updated and extended by computer. The results are presented both in his format and in a much more convenient one by Kramer that avoids iteration: curves of power coefficient at constant ideal efficiency are plotted vs propeller advance coefficient. The curves for a wide range of blade numbers are collapsed into just three sets (with some approximation) by use of multiple, shifted (and distorted) abscissae scales. Along with an overview of Theodorsen's theory, analytic asymptotic results at low and high advance coefficients are given. At the low end, the disagreement with actuator disk theory is given support and physical interpretation. At the high end, exact agreement is found with the thrust of a slender twisted delta propeller.

Ribner, H. S.↗