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At least 649 records · Page 36

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NASA NTRS1990

Forebody vortex control as a complement to thrust vectoring

The desire to enhance the controllability of fighter aircraft at high angles of attack, particularly yaw control, has fostered an interest in both vectored thrust and active control of forebody vortices. This paper reviews several methods of forebody vortex control that have been investigated with water and wind tunnel tests of both generic and actual fighter configurations. The methods investigated include pneumatic or blowing techniques using surface-mounted jets and slots, surface suction, variable-height deployable strakes, and rotatable tip strakes. Flow visualization, and force and moment measurements have shown that all of the methods are effective in manipulating the forebody vortices over a wide range of angles of attack and sideslip, primarily through control over flow separation on the surface of the forebody. All are most effective when applied near the forebody tip. The advantages and limitations of the various methods are reviewed.

Malcolm, G. N.↗

NASA NTRS1986

Effects of the law for nongravitational forces on the precession model of comet Encke

The precession model for P/Encke formulated by Whipple and Sekanina (1979) is reexamined in the light of the recent measurements of the OH and H production rates. An optimized precession solution derived using the production law of A'Hearn et al. (1985) is shown to require unacceptably small nucleus dimensions and to offer an inferior fit to observed variations in the orientation of the comet's perihelion fan-shaped coma. Observed changes in the emission fan suggest that both the thrust on the nucleus exerted by unit mass of sublimating ice and the average lag in outgassing are substantially greater before than after perihelion. It is concluded that at this time the Whipple-Sekanina model remains the best available precession solution for this comet.

Sekanina, Z.↗

NASA NTRS2004

Conical Magnetic Bearing Development and Magnetic Bearing Testing for Extreme Temperature Environments

The main proposed research of this grant were: to design a high-temperature, conical magnetic bearing facility, to test the high-temperature, radial magnetic bearing facility to higher speeds, to investigate different backup bearing designs and materials, to retrofit the high-temperature test facility with a magnetic thrust bearing, to evaluate test bearings at various conditions, and test several lubricants using a spiral orbit tribometer. A high-temperature, conical magnetic bearing facility has been fully developed using Solidworks. The facility can reuse many of the parts of the current high-temperature, radial magnetic bearing, helping to reduce overall build costs. The facility has the ability to measure bearing force capacity in the X, Y, and Z directions through a novel bearing mounting design. The high temperature coils and laminations, a main component of the facility, are based upon the current radial design and can be fabricated at Texas A&M University. The coil design was highly successful in the radial magnetic bearing. Vendors were contacted about fabrication of the high temperature lamination stack. Stress analysis was done on the laminations. Some of the components were procured, but due to budget cuts, the facility build up was stopped.

Keith, Theo G., Jr.↗

NASA NTRS1972

Structural evaluations and dynamic testing of solar electric propulsion components.

Results of an experimental and analytical study of selected Solar Electric Propulsion (SEP) thrust subsystem components are presented. A mass and center-of-mass mockup of an experimental power conditioning (PC) panel was used for the PC structure evaluation. Test environment was based on that specified for prooftesting the Viking spacecraft electronics. Modal frequencies were surveyed over the range of 0-2000 Hz using both conventional test equipment and the holographic interferometry technique. Results of the forced vibration tests point to the need for some additional stiffening to make the structure qualifiable.

Womack, J. R.↗

NASA NTRS1984

Proposed system design for a 20 kW pulsed electrothermal thruster

A conceptual design is presented for a Pulsed Electrothermal (PET) propulsion system for the Air Force Space Based Radar satellite, which has a mass of 7000 kg. The proposed system boosts the SBR satellite from 150 n.m. to 600 n.m. with a 4 deg plane change, for a total mission Delta v of 1 km/sec. Satellite power available is 50 kW, and 45 kW are used to drive two water-injected 20 kW PET thrusters, delivering 5.6 N thrust to the SBR at 1000 seconds specific impulse. The predicted mission trip time is 15 days. The proposed system consumes 850 kg of water propellant, stored in a central tank and injected with pressurized helium. Component mass estimates based on space-qualified hardware are presented for the propellant handling, power conditioning and thruster subsystems. The estimated total mass is 400 kg and the propulsion system specific mass is alpha = 10 kg/kW. The proposed system efficiency of 0.62 at 1000 seconds specific impulse is supported by experimental performance measurements.

Burton, R. L.↗

NASA NTRS2003

High-Flow Jet Exit Rig Designed and Fabricated

The High-Flow Jet Exit Rig at the NASA Glenn Research Center is designed to test single flow jet nozzles and to measure the appropriate thrust and noise levels. The rig has been designed for the maximum hot condition of 16 lbm/sec of combustion air at 1960 R (maximum) and to produce a maximum thrust of 2000 lb. It was designed for cold flow of 29.1 lbm/sec of air at 530 R. In addition, it can test dual-flow nozzles (nozzles with bypass flow in addition to core flow) with independent control of each flow. The High- Flow Jet Exit Rig was successfully fabricated in late 2001 and is being readied for checkout tests. The rig will be installed in Glenn's Aeroacoustic Propulsion Laboratory. The High-Flow Jet Exit Rig consists of the following major components: a single component force balance, the natural-gas-fueled J-79 combustor assembly, the plenum and manifold assembly, an acoustic/instrumentation/seeding (A/I/S) section, a table, and the research nozzles. The rig will be unique in that it is designed to operate uncooled. The structure survives the 1960 R test condition because it uses carefully selected high temperature alloy materials such as Hastelloy-X. The lower plenum assembly was designed to operate at pressures to 450 psig at 1960 R, in accordance with the ASME B31.3 piping code. The natural gas-fueled combustor fires directly into the lower manifold. The hot air is directed through eight 1-1/2-in. supply pipes that supply the upper plenum. The flow is conditioned in the upper plenum prior to flowing to the research nozzle. The 1-1/2-in. supply lines are arranged in a U-shaped design to provide for a flexible piping system. The combustor assembly checkout was successfully conducted in Glenn's Engine Component Research Laboratory in the spring of 2001. The combustor is a low-smoke version of the J79 combustor used to power the F4 Phantom military aircraft. The natural gas-fueled combustor demonstrated high-efficiency combustion over a wide range of operating conditions. This wide operating envelope is required to support the testing of both single- and dual-flow nozzles. Key research goals include providing simultaneous, highly accurate acoustic, flow, and thrust measurements on jet nozzle models in realistic flight conditions, as well as providing scaleable acoustic results. The High-Flow Jet Exit Rig is a second-generation high-flow test rig. Improvements include cleaner flow with reduced levels of particulate, soot, and odor. Choked-flow metering is required with plus or minus 0.25-percent accuracy. Thrust measurements from 0 to 2000 lbf are required with plus or minus 0.25-percent accuracy. Improved acoustics will be achieved by minimizing noise through large pipe bend radii, lower internal flow velocities, and microdrilled choke plates with thousands of 0.040-in.- diameter holes.

Buehrle, Robert J.↗

NASA NTRS1977

Space shuttle orbiter rear mounted reaction control system jet interaction study

The effect of interaction between the reaction control system (RCS) jets and the flow over the space shuttle orbiter in the atmosphere was investigated in the NASA Langley 31-inch continuous flow hypersonic tunnel at a nominal Mach number of 10.3 and in the AEDC continuous flow hypersonic tunnel B at a nominal Mach number of 6, using 0.01 and .0125 scale force models with aft RCS nozzles mounted both on the model and on the sting of the force model balance. The data show that RCS nozzle exit momentum ratio is the primary correlating parameter for effects where the plume impinges on an adjacent surface and mass flow ratio is the parameter when the plume interaction is primarily with the external stream. An analytic model of aft mounted RCS units was developed in which the total reaction control moments are the sum of thrust, impingement, interaction, and cross-coupling terms.

Rausch, J. R.↗

NASA NTRS1983

An aerodynamic analysis computer program and design notes for low speed wing flap systems

The expanded capabilities for analysis and design of low speed flap systems afforded by recent modifications of an existing computer program is described. The program provides for the simultaneous analysis of up to 25 pairs of leading-edge and trailing-edge flap deflection schedules. Among other new features of the program are a revised attainable thrust estimation method to provide more accurate predictions for low Mach numbers, and a choice of three options for estimation of leading-edge separation vortex flow effects. Comparison of program results with low speed experimental data for an arrow wing supersonic cruise configuration with leading-edge and trailing-edge flaps showed good agreement over most of the range of flap deflections. Other force data comparisons and an independent study of airfoil and wing pressure distributions indicated that wind-tunnel measurements of the aerodynamic performance of twisted and cambered wings and wings with leading-edge flaps can be very sensitive to Reynolds number effects.

Carlson, H. W.↗

NASA NTRS1954

A Wind-Tunnel Investigation of the Effects of Thrust-Axis Inclination on Propeller First-Order Vibration

Gray, W H↗

NASA NTRS1980

DOD low-thrust mission studies

The space transportation system (STS) will be the principal means of launching USAF spacecraft beginning in the 1980's. Since it is manned and reusable it provides new opportunities for unique approaches for cost effective utilization of its capabilities. The STS also places additional requirements and constraints on advanced spacecraft deployment systems that did not previously exist for expandable launch vehicles. To fully utilize these new capabilities designers must be prepared by having cost effective technologies available. Advanced propulsion technology that would provide flexibility, performance, and economic benefits to future Air Force missions was identified. Both electric and chemical propulsion systems are discussed. An LO2/LH2 stage with a torus LO2 tank and 500 lbf pumb fed engine is high on the list of propulsion technology.

Pipes, W. E.↗

NASA NTRS1992

Flap-lag damping in hover and forward flight with a three-dimensional wake

Prediction of lag damping is difficult owing to the delicate balance of drag, induced drag and Coriolis forces in the in-plane direction. Moreover, induced drag is sensitive to dynamic wake, both shed and trailing components, and thus its prediction requires adequate unsteady-wake representation. Accordingly, rigid-blade flap-lag equations are coupled with a three-dimensional finite-state wake model; three isolated rotor configurations with three, four and five blades are treated over a range of thrust levels, Lock numbers, lag frequencies and advance ratios. The investigation includes convergence characteristics of damping with respect to the number of shape functions and harmonics of the wake model for multiblade modes of low frequency (less than 1/rev.) to high frequency (greater than l/rev.). Predicted flap and lag damping levels are then compared with similar predictions with (1) rigid wake (no unsteady induced flow), (2) Loewy lift deficiency, and (3) dynamic inflow.

Manjunath, A. R.↗

NASA NTRS1950

A wind-tunnel investigation of the effects of thrust-axis inclination on propeller first-order vibration

Data on the aerodynamic excitation of first-order vibration occurring in a representative three-blade propeller having its thrust axis inclined to the air stream at angles of 0 degrees, 4.55 degrees, and 9.8 degrees are included in this paper. For several representative conditions the aerodynamic excitation has been computed and compared with the measured values. Blade stresses also were measured to permit the evaluation of the blade stress resulting from a given blade aerodynamic excitation. It was concluded that the section aerodynamic exciting force of a pitched propeller may be computed accurately at low rotational speeds. As section velocities approach the speed of sound, the accuracy of computation of section aerodynamic exciting force is not always so satisfactory. First-order blade vibratory stresses were computed with satisfactory accuracy from untilted-propeller loading data. A stress prediction which assumes a linear relation between first-order vibratory stress and the product of pitch angle and dynamic pressure and which is based on stresses at low rotational speeds will be conservative when the outer portions of the blade are in the transonic and low supersonic speed range.

Gray, W H↗

NASA NTRS1993

Flap-lag damping in hover and forward flight with a three-dimensional wake

Prediction of lag damping is difficult owing to the delicate balance of drag, induced drag and Coriolis forces in the in-plane direction. Moreover, induced drag is sensitive to dynamic wake, both shed and trailing components, and thus its prediction requires adequate unsteady-wake representation. Accordingly, rigid-blade flap-lag equations are coupled with a three-dimensional finite-state wake model; three isolated rotor configurations with three, four and five blades are treated over a range of thrust levels, Lock numbers, lag frequencies and advance ratios. The investigation includes convergence characteristics of damping with respect to the number of radial shape functions and harmonics of the wake model for multiblade modes of low frequency (less than 1/rev.) to high frequency (greater than 1/rev.). Predicted flap and lag damping levels are then compared with similar predictions with (1) rigid wake (no unsteady induced flow, (2) Loewy lift defficiency and (3) dynamic inflow. The coverage also includes correlations with the measured lag regressive-mode damping in hover and forward flight and comparisons with similar correlations with dynamic wake model are consistently higher than the predictions with the dynamic inflow model; even for the low frequency lag regressive mode, the number of wake harmonics should at least be equal to twice the number of blades.

Manjunath, A. R.↗

NASA NTRS1999

Upper-Stage Flight Experiment

For propulsion applications that require that the propellants are storable for long periods, have a high density impulse, and are environmentally clean and non-toxic, the best choice is a combination of high-concentration hydrogen peroxide (High Test Peroxide, or HTP) and a liquid hydrocarbon (LHC) fuel. The HTP/LHC combination is suitable for low-cost launch vehicles, space taxi and space maneuvering vehicles, and kick stages. Orbital Sciences Corporation is under contract with the NASA Marshall Space Flight Center in cooperation with the Air Force Research Lab to design, develop and demonstrate a new low-cost liquid upper stage based on HTP and JP-8. The Upper Stage Flight Experiment (USFE) focuses on key technologies necessary to demonstrate the operation of an inherently simple propulsion system with an innovative, state-of-the-art structure. Two key low-cost vehicle elements will be demonstrated - a 10,000 lbf thrust engine and an integrated composite tank structure. The suborbital flight test of the USFE is scheduled for 2001. Preceding the flight tests are two major series of ground tests at NASA Stennis Space Center and a subscale tank development program to identify compatible composite materials and to verify their compatibility over long periods of time. The ground tests include a thrust chamber development test series and an integrated stage test. This paper summarizes the results from the first phase of the thrust chamber development tests and the results to date from the tank material compatibility tests. Engine and tank configurations that meet the goals of the program are described.

Anderson, W. E.↗

NASA NTRS1994

Design and flight test of the Propulsion Controlled Aircraft (PCA) flight control system on the NASA F-15 test aircraft

This report describes the design, development and flight testing of the Propulsion Controlled Aircraft (PCA) flight control system performed at McDonnell Douglas Aerospace (MDA), St. Louis, Missouri and at the NASA Dryden Flight Research Facility, Edwards Air Force Base, California. This research and development program was conducted by MDA and directed by NASA through the Dryden Flight Research Facility for the period beginning January 1991 and ending December 1993. A propulsion steering backup to the aircraft conventional flight control system has been developed and flight demonstrated on a NASA F-15 test aircraft. The Propulsion Controlled Aircraft (PCA) flight system utilizes collective and differential thrust changes to steer an aircraft that experiences partial or complete failure of the hydraulically actuated control surfaces. The PCA flight control research has shown that propulsion steering is a viable backup flight control mode and can assist the pilot in safe landing recovery of a fighter aircraft that has damage to or loss of the flight control surfaces. NASA, USAF and Navy evaluation test pilots stated that the F-15 PCA design provided the control necessary to land the aircraft. Moreover, the feasibility study showed that PCA technology can be directly applied to transport aircraft and provide a major improvement in the survivability of passengers and crew of controls damaged aircraft.

Wells, Edward A.↗

NASA NTRS1994

Comparison of X-31 flight, wind-tunnel, and water-tunnel yawing moment asymmetries at high angles of attack

The X-31 aircraft are being used in the enhanced fighter maneuverability (EFM) research program, which is jointly funded by the (U.S.) Advanced Research Projects Agency (ARPA) and Germany's Federal Ministry of Defense (FMOD). The flight test portion of the program, which involves two aircraft, is being conducted by an International Test Organization (ITO) comprising the National Aeronautics and Space Administration (NASA), the U.S. Navy, the U.S. Air Force, Rockwell International, and Deutsche Aerospace (DASA). The goals of the flight program are to demonstrate EFM technologies, investigate close-in-combat exchange ratios, develop design requirements, build a database for application to future fighter aircraft, and develop and validate low-cost prototype concepts. For longitudinal control the X-31 uses canards, symmetrical movement of the trailing-edge flaps, and pitch deflection of the thrust vectoring system. The trim, inertial coupling, and engine gyroscopic coupling compensation tasks are performed primarily by the trailing-edge flaps. For lateral-directional control the aircraft uses differential deflection of the trailing-edge flaps for roll coordination and a conventional rudder combined with the thrust vectoring system to provide yaw control. The rudder is only effective up to about 40 deg angle of attack (alpha), after which the thrust vectoring becomes the primary yaw control effector. Both the leading-edge flaps and the inlet lip are scheduled with the angle of attack to provide best performance.

Brent R Cobleigh↗

NASA NTRS1997

Enhanced Control Effector Designs for Airbreathing Transatmospheric Vehicles

A study was conducted to evaluate the potential effectiveness of a moveable cowl-trailing-edge design for airbreathing hypersonic single-stage-to-orbit (SSTO) configurations, which can be extended or deflected from the nominal position in order to provide additional pitching moment capability. This additional pitching moment capability may reduce the necessary deflection angle from conventional control surfaces and the associated trim drag penalty. Calculations for a generic SSTO configuration with baseline and modified cowl trailing edge geometries were performed at Mach 6 and 10 initially with an engineering analysis code in order to examine several design parametrics. In order to more accurately model geometries and flow physics, 2-D viscous computational fluid dynamics (CFD) predictions were obtained. FInally, a limited set of 3D CFD predictions were obtained at Mach 6 in order to show the effects of modeling 3D flow fields as well as the full 3D vehicle geometry. Comparisons of aftbody surface pressures and force and moment predictions show differences between initial predictions and 2-D CFD solutions due to geometry modeling and calculation method differences. The 3-D CFD predictions confirm the trends observed in the 2-D solutions and provide additional information on 3D effects. These analyses shows that cowl-trailing-edge extensions were effective in providing additional (nose-down) pitching moment increments as well as increased thrust compared to the baseline geometry. These effects reduce the control surface deflection angle required to trim the vehicle and the associated trim drag. Cowl-trailingedge deflections were not as effective, generating more nose-up pitching moment and decreased thrust compared to non-deflected cases.

Cockrell, Charles E.↗

NASA NTRS2005

AOCS Performance and Stability Validation for a 160-m Solar Sail with Control-Structure Interactions

Future solar sail missions, such as NASA's Solar Polar Imager Vision, will require sails with dimensions on the order of 50-500 m. We are examining a square sail design with moving mass (trim control mass, TCM) and quadrant rotation primary actuators plus pulsed plasma thrusters (PPTs) at the mast tips for backup attitude control. Quadrant rotation is achieved via roll stabilizer bars (RSB) at the mast tips. At these sizes, given the gossamer nature of the sail supporting structures, flexible modes may be low enough to interact with the control system, especially as these actuators are located on the flexible structure itself and not on the rigid core. This paper develops a practical analysis of the flexible interactions using state-space systems and modal data from finite element models of the system. Torsion and bending of the masts during maneuvers could significantly affect the function of the actuators while activation of the membrane modes could adversely affect the thrust vector direction and magnitude. Analysis of the RSB and TCM dynamics for developing high-fidelity simulations is included. For control analysis of the flexible system, standard finite-element models of the flexible sail body are loaded and the modal data is used to create a modal coordinate state-space system. Key parameters include which modes to include, which nodes are of interest for force inputs and displacement outputs, connecting nodes through which external forces and torques are applied from the flex body to the core, any nominal momentum in the system, and any steady rates. The system is linearized about the nominal attitude and rate. The state-space plant can then be analyzed with a state-space controller, and Bode, Nyquist, step and impulse responses generated. The approach is general for any rigid core with a flexible appendage. This paper develops a compensator for a simple two-mass flex system and extrapolates the results to the solar sail. A finite element model of the 20 m solar sail by ATK Space Systems, recently validated in ground tests, is used to demonstrate the sail analysis approach.

Wie, Bong↗