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At least 775 records · Page 43
All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

Experimental Results of Schlicher's Thrusting Antenna

Experiments were conducted to test the claims by Rex L. Schlicher, et al., (Patent 5,142,86 1) that a certain antenna geometry produces thrust greatly exceeding radiation reaction, when driven by repetitive, fast rise, and relatively slower decay current pulses. In order to test this hypothesis, the antenna was suspended by strings as a 3 in pendulum. Current pulses were fed to the antenna along the suspension path by a very flexible coaxial line constructed from loudspeaker cable and copper braid sheath. When driving the antenna via this cabling, our pulser was capable of sustaining 1200 A pulses at a rate of 30 per second up to a minute. In this way, bursts of pulses could be delivered in synch with the pendulum period in order to build up any motion. However, when using a laser beam passing through a lens attached to the antenna to amplify linear displacement by a factor of at least 25, no correlated motion of the beam spot could be detected on a distant wall. We conclude, in agreement with the momentum theorem of classical electromagnetic theory, that any thrust produced is far below practically useful levels. Hence, within classical electrodynamics, there is little hope of detecting any low level motion that cannot be explained by interactions with surrounding structural steel and the Earth's magnetic field.

Fralick, Gustave C.↗
NASA NTRS2001

Elimination of Intermediate-Frequency Combustion Instability in the Fastrac Engine Thrust Chamber

A series of tests were conducted to measure the combustion performance of the Fastrac engine thrust chamber. The thrust chamber exhibited benign, yet marginally unstable combustion. The marginally unstable combustion was characterized by chamber pressure oscillations with large amplitudes and a frequency that was too low to be identified as acoustic or high-frequency combustion instability and too high to be identified as chug or low-frequency combustion instability. The source of the buzz or intermediate-frequency combustion instability was traced to the fuel venturi whose violently noisy cavitation caused resonance in the feedline downstream. Combustion was stabilized by increasing the throat diameter of the fuel venturi such that the cavitation would occur more quietly.

Rocker, Marvin↗
NASA NTRS2001

Advanced Methods for Aircraft Engine Thrust and Noise Benefits: Nozzle-Inlet Flow Analysis

The Fluid Mechanics and Acoustics Laboratory at Hampton University (HU/FM&AL) jointly with the NASA Glenn Research Center has conducted four connected subprojects under the reporting project. Basically, the HU/FM&AL Team has been involved in joint research with the purpose of theoretical explanation of experimental facts and creation of accurate numerical simulation techniques and prediction theory for solution of current problems in propulsion systems of interest to the NAVY and NASA agencies. This work is also supported by joint research between the NASA GRC and the Institute of Mechanics at Moscow State University (IM/MSU) in Russia under a CRDF grant. The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analyses for advanced aircraft and rocket engines. The FM&AL Team uses analytical methods, numerical simulations and possible experimental tests at the Hampton University campus. The fundamental idea uniting these subprojects is to use nontraditional 3D corrugated and composite nozzle and inlet designs and additional methods for exhaust jet noise reduction without essential thrust loss and even with thrust augmentation. These subprojects are: (1) Aeroperformance and acoustics of Bluebell-shaped and Telescope-shaped designs; (2) An analysis of sharp-edged nozzle exit designs for effective fuel injection into the flow stream in air-breathing engines: triangular-round, diamond-round and other nozzles; (3) Measurement technique improvement for the HU Low Speed Wind Tunnel; a new course in the field of aerodynamics, teaching and training of HU students; experimental tests of Mobius-shaped screws: research and training; (4) Supersonic inlet shape optimization. The main outcomes during this reporting period are: (l) Publications: The AIAA Paper #00-3170 was presented at the 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 17-19 June, 2000, Huntsville, AL. The AIAA Paper #01-1893 has been accepted for the AIAA/NAL-NASDA-ISAS 10th International Space Planes and Hypersonic Systems and Technologies Conference, 24-27 April 2001, Kyoto, Japan. The AIAA Paper #01 -3204 has been accepted for presentation at the 37th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, being held on 08-11 July, in Salt Lake City, UT; (2) A U.S. patent #6,082,635 was granted on July 4, 2000; (3) Grants and proposals: The H U/ FM&AL was awarded the NASA grant NAG-3-2495 in October 2000 and the laboratory is a primary U.S. research team in a joint project under the CRDF award granted to the NASA GRC and IM/MSU (Russia) in July 2000; (4) Theory and numerical simulations: Analytical theory, numerical simulation, comparison of theoretical with experimental results, and modification of theoretical approaches, models, grids, etc., have been conducted for several complicated 2D and 3D nozzle and inlet designs using NASA, ICASE, and IM/MSU codes based on full Euler and Navier-Stokes solvers: CFL3D, FLUENT, and GODUNOV, and others; (5) Experimental Tests: (a) A new course: "Advanced Aerodynamics and Aircraft Performance" presented in spring semester, 2001; training and experimental test research using the HU LSWT. (b) Small-scale M6bius-shaped screws were tested in different conditions and their application has shown essential benefits by comparison with traditional designs; (6) Installation in the FM&AL computer system: second software TECPLOT 8.0 for the UNIX SGI workstation and free TECPLOT 7.5 for the PC Dell computer, and 2D and 3D GRIDGEN (version 9) for the UNIX SGI as well as installation of two free NASA codes, 3D MAG and VULCAN; (7) Student Research Activity: Involvement of two undergraduate students as research assistants in the current research project.

Gilinsky, Mikhail↗
NASA NTRS2003

Non-Toxic Dual Thrust Reaction Control Engine Development for On-Orbit APS Applications

A non-toxic dual thrust proof-of-concept demonstration engine was successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and Space Administration's (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the Next Generation Launch Technology (NGLT) program. The demonstration engine utilized the existing Kistler K-1 870 lbf LOX/Ethanol orbital maneuvering engine ( O m ) coupled with some special test equipment (STE) that enabled engine operation at 870 lbf in the primary mode and 25 lbf in the vernier mode. Ambient testing in primary mode varied mixture ratio (MR) from 1.28 to 1.71 and chamber pressure (P(c) from 110 to 181 psia, and evaluated electrical pulse widths (EPW) of 0.080, 0.100 and 0.250 seconds. Altitude testing in vernier mode explored igniter and thruster pulsing characteristics, long duration steady state operation (greater than 420 sec) and the impact of varying the percent fuel film cooling on vernier performance and chamber thermal response at low PC (4 psia). Data produced from the testing provided calibration of the performance and thermal models used in the design of the next version of the dual thrust Reaction Control Engine (RCE).

Robinson, Philip J.↗
NASA NTRS2004

Corrugated and Composite Nozzle-Inlets for Thrust and Noise Benefits

The following research results are based on development of an approach previously proposed and investigated in for optimum nozzle design to obtain maximum thrust. The design was denoted a Telescope nozzle. A Telescope nozzle contains one or several internal designs, which are inserted at certain locations into a divergent conical or planar main nozzle near its exit. Such a design provides additional thrust augmentation over 20% by comparison with the optimum single nozzle of equivalent lateral area, What is more, experimental acoustic tests have discovered an essential noise reduction due to application of Telescope nozzles. In this paper, some additional theoretical results are presented for Telescope nozzles and a similar approach is applied for aero-performance improvement of a supersonic inlet. Numerical simulations were conducted for supersonic flow into the divergent portion of a 2D or axisymmetric nozzle with several plane or conical designs as well as into a 2D or axisymmetric supersonic inlet with a forebody. The Kryko-Godunov marching numerical scheme for inviscid supersonic flows was used. Several cases were tested using the NASA CFL3d and IM/MSU Russian codes based on the full Navier-Stokes equations. Numerical simulations were conducted for non reacting flows (both codes) as well as for real high temperature gas flows with non-equilibrium chemical reactions (the latter code). In general, these simulations have confirmed essential benefits of Telescope design applications in propulsion system. Some preliminary numerical simulations of several typical inlet designs were conducted with the goal of inlet design optimization for maneuvering flight conditions.

Gilinsky, M.↗
NASA NTRS1998

Experimental Study of a Nozzle Using Fluidic Counterflow for Thrust Vectoring

A static experimental investigation of a counterflow thrust vectoring nozzle concept was performed. The study was conducted in the NASA Langley Research Center Jet Exit Test Facility. Internal performance characteristics were defined over a nozzle pressure ratio (jet total to ambient) range of 3.5 to 10.0. The effects of suction collar geometry and suction slot height on nozzle performance were examined. In the counterflow concept, thrust vectoring is achieved by applying a vacuum to a slot adjacent to a primary jet that is shrouded by a suction collar. Two flow phenomena work to vector the primary jet depending upon the test conditions and configuration. In one case, the vacuum source creates a secondary reverse flowing stream near the primary jet. The shear layers between the two counterflowing streams mix and entrain mass from the surrounding fluid. The presence of the collar inhibits mass entrainment and the flow near the collar accelerates, causing a drop in pressure on the collar. The second case works similarly except that the vacuum is not powerful enough to create a counterflowing stream and instead a coflowing stream is present. The primary jet is vectored if suction is applied asymmetrically on the top or bottom of the jet.

Flamm, Jeffrey D.↗
NASA NTRS1948

Altitude-Wind-Tunnel Investigation of a 4000-Pound-Thrust Axial-Flow Turbojet Engine. II - Operational Characteristics: Operational Characteristics - II

An investigation was conducted in the Cleveland altitude wind tunnel to determine the operational characteristics of an axial flow-type turbojet engine with a 4000-pound-thrust rating over a range of pressure altitudes from 5,000 to 50,OOO feet, ram pressure ratios from 1.00 to 1.86, and temperatures from 60 deg to -50 deg F. The low-flow (standard) compressor with which the engine was originally equipped was replaced by a high-flow compressor for part of the investigation. The effects of altitude and airspeed on such operating characteristics as operating range, stability of combustion, acceleration, starting, operation of fuel-control systems, and bearing cooling were investigated. With the low-flow compressor, the engine could be operated at full speed without serious burner unbalance at altitudes up to 50,000 feet. Increasing the altitude and airspeed greatly reduced the operable speed range of the engine by raising the minimum operating speed of the engine. In several runs with the high-flow compressor the maximum engine speed was limited to less than 7600 rpm by combustion blow-out, high tail-pipe temperatures, and compressor stall. Acceleration of the engine was relatively slow and the time required for acceleration increased with altitude. At maximum engine speed a sudden reduction in jet-nozzle area resulted in an immediate increase in thrust. The engine started normally and easily below 20,000 feet with each configuration. The use of a high-voltage ignition system made possible starts at a pressure altitude of 40,000 feet; but on these starts the tail-pipe temperatures were very high, a great deal of fuel burned in and behind the tail-pipe, and acceleration was very slow. Operation of the engine was similar with both fuel regulators except that the modified fuel regulator restricted the fuel flow in such a manner that the acceleration above 6000 rpm was very slow. The bearings did not cool properly at high altitudes and high engine speeds with a low-flow compressor, and bearing cooling was even poorer with a high-flow compressor.

Fleming, William A.↗
NASA NTRS1955

Free-Spinning-Tunnel Investigation to Determine the Effect of Spin-Recovery Rockets and Thrust Simulation on the Recovery Characteristics of a 1/21-Scale Model of the Chance Vought F7U-3 Airplane, TED No. NACA AD 3103

An investigation of a l/21-scale model of the Chance Vought F7U-3 airplane in the co&at-load- condition has been conducted in the Langley 20-foot free-spinning tunnel, The recovery characteristics of the model were determined by use of spin-recovery rockets for the erect and inverted spinning condition. The rockets were so placed as to provide either a yawing or rolling moment about the model center of gravity. Also included in the investigation were tests to determine the effect of simulated engine thrust on the recovery characteristics of the model. On the basis of model tests, recoveries from erect and inverted spins were satisfactory when a yawing moment of 22,200 foot-pounds (full scale) was provided against the spin by rockets attached to the wing tips; the anti-spin yawing moment was applied for approximately 9 seconds, (full scale). Satisfactory recoveries were obtained from erect spins when a rolling moment of 22,200 foot-pounds (full scale) was provided with the spin (rolls right wing down in right spin). Although the inverted spin was satisfactorily terminated when a rolling moment of equal magnitude was provided, a roll rocket was not considered to be an optimum spin-recovery device to effect recoveries from inverted spins for this airplane because of resulting gyrations during spin recovery. Simulation of engine thrust had no apparent effect on the spin recovery characteristics.

Burk, Sanger H., Jr.↗
NASA NTRS1948

Theoretical Comparison of Several Methods of Thrust Augmentation for Turbojet Engines

A theoretical investigation has been made of various methods of thrust augmentation for turbojet engines. The method investigated were tail-pipe burning, water injection at the compressor inlet, a combination of tail-pipe burning and water injection, bleedoff in conjunction with water injection at the compressor inlet, and rocket assist. The effect of ratio of augmented-to-normal total liquid consumption, flight conditions, and design compressor pressure ratio on the augmentation produced by each method were determined. A comparison was also made for a given time of operation of the weight of an augmented engine plus fuel and additional liquids to the weight of a standard engine plus fuel producing the same thrust.

Hall, Eldon W.↗
NASA NTRS1995

Cascade Optimization Strategy Maximizes Thrust for High-Speed Civil Transport Propulsion System Concept

The design of a High-Speed Civil Transport (HSCT) air-breathing propulsion system for multimission, variable-cycle operations was successfully optimized through a soft coupling of the engine performance analyzer NASA Engine Performance Program (NEPP) to a multidisciplinary optimization tool COMETBOARDS that was developed at the NASA Lewis Research Center. The design optimization of this engine was cast as a nonlinear optimization problem, with engine thrust as the merit function and the bypass ratios, r-values of fans, fuel flow, and other factors as important active design variables. Constraints were specified on factors including the maximum speed of the compressors, the positive surge margins for the compressors with specified safety factors, the discharge temperature, the pressure ratios, and the mixer extreme Mach number. Solving the problem by using the most reliable optimization algorithm available in COMETBOARDS would provide feasible optimum results only for a portion of the aircraft flight regime because of the large number of mission points (defined by altitudes, Mach numbers, flow rates, and other factors), diverse constraint types, and overall poor conditioning of the design space. Only the cascade optimization strategy of COMETBOARDS, which was devised especially for difficult multidisciplinary applications, could successfully solve a number of engine design problems for their flight regimes. Furthermore, the cascade strategy converged to the same global optimum solution even when it was initiated from different design points. Multiple optimizers in a specified sequence, pseudorandom damping, and reduction of the design space distortion via a global scaling scheme are some of the key features of the cascade strategy. HSCT engine concept, optimized solution for HSCT engine concept. A COMETBOARDS solution for an HSCT engine (Mach-2.4 mixed-flow turbofan) along with its configuration is shown. The optimum thrust is normalized with respect to NEPP results. COMETBOARDS added value in the design optimization of the HSCT engine.

Source record↗
NASA NTRS2005

Analytic Modeling of the Hydrodynamic, Thermal, and Structural Behavior of Foil Thrust Bearings

A simulation and modeling effort is conducted on gas foil thrust bearings. A foil bearing is a self acting hydrodynamic device capable of separating stationary and rotating components of rotating machinery by a film of air or other gaseous lubricant. Although simple in appearance these bearings have proven to be complicated devices in analysis. They are sensitive to fluid structure interaction, use a compressible gas as a lubricant, may not be in the fully continuum range of fluid mechanics, and operate in the range where viscous heat generation is significant. These factors provide a challenge to the simulation and modeling task. The Reynolds equation with the addition of Knudsen number effects due to thin film thicknesses is used to simulate the hydrodynamics. The energy equation is manipulated to simulate the temperature field of the lubricant film and combined with the ideal gas relationship, provides density field input to the Reynolds equation. Heat transfer between the lubricant and the surroundings is also modeled. The structural deformations of the bearing are modeled with a single partial differential equation. The equation models the top foil as a thin, bending dominated membrane whose deflections are governed by the biharmonic equation. A linear superposition of hydrodynamic load and compliant foundation reaction is included. The stiffness of the compliant foundation is modeled as a distributed stiffness that supports the top foil. The system of governing equations is solved numerically by a computer program written in the Mathematica computing environment. Representative calculations and comparisons with experimental results are included for a generation I gas foil thrust bearing.

Bruckner, Robert J.↗
NASA NTRS2005

Preliminary Assessment of Thrust Augmentation of NEP Based Missions

Science Applications International Corporation (SAIC), with support from NASA Marshall Space Flight Center, has conducted a preliminary study to compare options for augmenting the thrust of a conventional nuclear electric propulsion (NEP) system. These options include a novel nuclear propulsion system concept known as Hybrid Indirect Nuclear Propulsion (HINP) and conventional chemical propulsion. The utility and technical feasibility of the HINP concept are assessed, and features and potential of this new in-space propulsion system concept are identified. As part of the study, SAIC developed top-level design tools to model the size and performance of an HINP system, as well as for several chemical propulsion options, including liquid and gelled propellants. A mission trade study was performed to compare a representative HINP system with chemical propulsion options for thrust augmentation of NEP systems for a mission to Saturn's moon Titan. Details pertaining to the approach, features, initial demonstration results for HINP model development, and the mission trade study are presented. Key technology and design issues associated with the HINP concept and future work recommendations are also identified.

Chew, Gilbert↗
NASA NTRS2005

High-Temperature (1000 F) Magnetic Thrust Bearing Test Rig Completed and Operational

Large axial loads are induced on the rolling element bearings of a gas turbine. To extend bearing life, designers use pneumatic balance pistons to reduce the axial load on the bearings. A magnetic thrust bearing could replace the balance pistons to further reduce the axial load. To investigate this option, the U.S. Army Research Laboratory, the NASA Glenn Research Center, and Texas A&M University designed and fabricated a 7-in.- diameter magnetic thrust bearing to operate at 1000 F and 30,000 rpm, with a 1000-lb load capacity. This research was funded through a NASA Space Technology Transfer Act with Allison Advance Development Company under the Ultra-Efficient Engine Technology (UEET) Intelligent Propulsion Systems Foundation Technology project.

Montague, Gerald T.↗
NASA NTRS1990

RS-88 Pad Abort Demonstrator Thrust Chamber Assembly Testing at NASA Marshall Space Flight Center

This paper documents the effort conducted to collect hot-tire dynamic and acoustics environments data during 50,000-lb thrust lox-ethanol hot-fire rocket testing at NASA Marshall Space Flight Center (MSFC) in November-December 2003. This test program was conducted during development testing of the Boeing Rocketdyne RS-88 development engine thrust chamber assembly (TCA) in support of the Orbital Space Plane (OSP) Crew Escape System Propulsion (CESP) Program Pad Abort Demonstrator (PAD). In addition to numerous internal TCA and nozzle measurements, induced acoustics environments data were also collected. Provided here is an overview of test parameters, a discussion of the measurements, test facility systems and test operations, and a quality assessment of the data collected during this test program.

Farr, Rebecca A.↗
NASA NTRS2005

An Optimal Orthogonal Decomposition Method for Kalman Filter-Based Turbofan Engine Thrust Estimation

A new linear point design technique is presented for the determination of tuning parameters that enable the optimal estimation of unmeasured engine outputs such as thrust. The engine s performance is affected by its level of degradation, generally described in terms of unmeasurable health parameters related to each major engine component. Accurate thrust reconstruction depends upon knowledge of these health parameters, but there are usually too few sensors to be able to estimate their values. In this new technique, a set of tuning parameters is determined which accounts for degradation by representing the overall effect of the larger set of health parameters as closely as possible in a least squares sense. The technique takes advantage of the properties of the singular value decomposition of a matrix to generate a tuning parameter vector of low enough dimension that it can be estimated by a Kalman filter. A concise design procedure to generate a tuning vector that specifically takes into account the variables of interest is presented. An example demonstrates the tuning parameters ability to facilitate matching of both measured and unmeasured engine outputs, as well as state variables. Additional properties of the formulation are shown to lend themselves well to diagnostics.

Litt, Jonathan S.↗
NASA NTRS2005

Proof-of-Concept Trajectory Designs for a Multi-Spacecraft, Low-Thrust Heliocentric Solar Weather Buoy Mission

A new solar weather mission has been proposed, involving a dozen or more small spacecraft spaced at regular, constant intervals in a mutual heliocentric circular orbit between the orbits of Earth and Venus. These solar weather buoys (SWBs) would carry instrumentation to detect and measure the material in solar flares, solar energetic particle events, and coronal mass ejections as they flowed past the buoys, serving both as science probes and as a radiation early warning system for the Earth and interplanetary travelers to Mars. The baseline concept involves placing a mothercraft carrying the SWBs into a staging orbit at the Sun-Earth L1 libration point. The mothercraft departs the L1 orbit at the proper time to execute a trailing-edge lunar flyby near New Moon, injecting it into a heliocentric orbit with its perihelion interior to Earth s orbit. An alternative approach would involve the use of a Double Lunar Swingby (DLS) orbit, rather than the L1 orbit, for staging prior to this flyby. After injection into heliocentric orbit, the mothercraft releases the SWBs-all equipped with low-thrust pulsed plasma thrusters (PPTs)-whereupon each SWB executes a multi-day low-thrust finite bum around perihelion, lowering aphelion such that each achieves an elliptical phasing orbit of different orbital period from its companions. The resulting differences in angular rates of motion cause the spacecraft to separate. While the lead SWB achieves the mission orbit following an insertion burn at its second perihelion passage, the remaining SWBs must complete several revolutions in their respective phasing orbits to establish them in the mission orbit with the desired longitudinal spacing. The complete configuration for a 14 SWB scenario using a single mothercraft is achieved in about 8 years, and the spacing remains stable for at least a further 6 years. Flight operations can be simplified, and mission risk reduced, by employing two mothercraft instead of one. In this scenario: the second mothercraft stays in a libration-point or DLS staging orbit until the first mothercraft has achieved nearly 180 separation from the Earth. The timing of the second mothercraft's subsequent lunar flyby is planned such that this spacecraft will be located 180 from the first mothercraft upon completion of its heliocentric circularization maneuvers. Both groups of satellites then only have to spread out over 180 to obtain full 360 coverage around the Sun.

Muller, Ronald↗
NASA NTRS2006

A Quick Method for Estimating Vehicle Characteristics Appropriate for Continuous Thrust Round Trip Missions Within the Solar System

A simple method of estimating vehicle parameters appropriate for interplanetary travel can provide a useful tool for evaluating the suitability of particular propulsion systems to various space missions. Although detailed mission analyses for interplanetary travel can be quite complex, it is possible to derive hirly simple correlations which will provide reasonable trip time estimates to the planets. In the present work, it is assumed that a constant thrust propulsion system propels a spacecraft on a round trip mission having equidistant outbound and inbound legs in which the spacecraft accelerates during the first portion of each leg of the journey and decelerates during the last portion of each leg of the journey. Comparisons are made with numerical calculations from low thrust trajectory codes to estimate the range of applicability of the simplified correlations.

Emrich, Bill↗
NASA NTRS2006

The NASA Constellation University Institutes Project: Thrust Chamber Assembly Virtual Institute

This paper documents key aspects of the Constellation University Institutes Project (CUIP) Thrust Chamber Assembly (TCA) Virtual Institute (VI). Specifically, the paper details the TCA VI organizational and functional aspects relative to providing support for Constellation Systems. The TCA VI vision is put forth and discussed in detail. The vision provides the objective and approach for improving thrust chamber assembly design methodologies by replacing the current empirical tools with verified and validated CFD codes. The vision also sets out ignition, performance, thermal environments and combustion stability as focus areas where application of these improved tools is required. Flow physics and a study of the Space Shuttle Main Engine development program are used to conclude that the injector is the key to robust TCA design. Requirements are set out in terms of fidelity, robustness and demonstrated accuracy of the design tool. Lack of demonstrated accuracy is noted as the most significant obstacle to realizing the potential of CFD to be widely used as an injector design tool. A hierarchical decomposition process is outlined to facilitate the validation process. A simulation readiness level tool used to gauge progress toward the goal is described. Finally, there is a description of the current efforts in each focus area. The background of each focus area is discussed. The state of the art in each focus area is noted along with the TCA VI research focus in the area. Brief highlights of work in the area are also included.

Tucker, P. Kevin↗