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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 487 records · Page 27
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

A solid-state controllable power supply for a magnetic suspension wind tunnel

The NASA Langley 6-inch Magnetic Suspension and Balance System (6-in. MSBS) requires an independently controlled bidirectional dc power source for each of six positioning electromagnets. These electromagnets provide five-degree-of-freedom control over a suspended aerodynamic test model. Existing power equipment, which employs resistance-coupled thyratron-controlled rectifiers as well as ac to dc motor-generator converters, is obsolete, inefficient, and unreliable. A replacement six-phase bidirectional controlled bridge rectifier is proposed, which employs power MOSFET switches sequenced by hybrid analog/digital circuits. Full-load efficiency is 80 percent compared with 25 percent for the resistance-coupled thyratron system. Current feedback provides high control linearity, adjustable current limiting, and current overload protection. A quenching circuit suppresses inductive voltage impulses. It is shown that 20-kHz interference from positioning magnet power into MSBS electromagnetic model position sensors results predominantly from capacitively coupled electric fields. Hence, proper shielding and grounding techniques are necessary. Inductively coupled magnetic interference is negligible.

Daniels, Taumi S.↗
NASA NTRS1960

Aerodynamic Performance and Static Stability at Mach Number 3.3 of an Aircraft Configuration Employing Three Triangular Wing Panels and a Body Equal Length

An aircraft configuration, previously conceived as a means to achieve favorable aerodynamic stability characteristics., high lift-drag ratio, and low heating rates at high supersonic speeds., was modified in an attempt to increase further the lift-drag ratio without adversely affecting the other desirable characteristics. The original configuration consisted of three identical triangular wing panels symmetrically disposed about an ogive-cylinder body equal in length to the root chord of the panels. This configuration was modified by altering the angular disposition of the wing panels, by reducing the area of the panel forming the vertical fin, and by reshaping the body to produce interference lift. Six-component force and moment tests of the modified configuration at combined angles of attack and sideslip were made at a Mach number of 3.3 and a Reynolds number of 5.46 million. A maximum lift-drag ratio of 6.65 (excluding base drag) was measured at a lift coefficient of 0.100 and an angle of attack of 3.60. The lift-drag ratio remained greater than 3 up to lift coefficient of 0.35. Performance estimates, which predicted a maximum lift-drag ratio for the modified configuration 27 percent greater than that of the original configuration, agreed well with experiment. The modified configuration exhibited favorable static stability characteristics within the test range. Longitudinal and directional centers of pressure were slightly aft of the respective centroids of projected plan-form and side area.

James, Carlton S.↗
NASA NTRS1951

Damping in a Roll of a Missile Configuration with a Modified Triangular Wing and a Cruciform Tail at a Mach Number of 1.52

The damping-in-Toll stability derivatives of a missile configuration and its components were determined both experimentally and theoretically. The tests were conducted at a Mach number of 1.52 and at a Reynolds number, based on the mean aerodynamic chord of the wing, of 0.82 x 10(exp 6). The experimental damping derivative of the wing-body combination was 67 percent of the theoretical value. The difference is believed to have resulted mainly from the fact that the theory is not strictly applicable when the Mach number normal to the leading edge is almost unity, which was the case in the present investigation. For the tail-body combination the damping derivative was 86 percent of the theoretical value. In this case, the difference is believed to have been caused partially by mutual interference between the tail surfaces and partially by the low Reynolds number of the flow over the tail. It was found that the damping of the complete configuration was not equal to the sum of the damping derivatives of the components because of the effect of the wing downwash on the damping of the tail.

Scherrer, Richard↗
NASA NTRS1976

Calculation of the longitudinal aerodynamic characteristics of wing-flap configurations with externally blown flaps

An analytical method for predicting the longitudinal aerodynamic characteristics of externally blown flap configurations is described. Two potential flow models make up the prediction method: a wing and flap lifting-surface model and a turbofan engine wake model. A vortex-lattice lifting-surface method is used to represent the wing and multiple-slotted trailing-edge flaps. The jet wake is represented by a series of closely spaced vortex rings normal to a centerline which is free to move to conform to the local flow field. The two potential models are combined in an iterative fashion to predict the jet wake interference effects on a typical EBF configuration. Comparisons of measured and predicted span-load distributions, individual surface forces, forces and moments on the complete configuration, and flow fields are included.

Mendenhall, M. R.↗
NASA NTRS1977

Results of tests using a 0.0125-scale model (70-QT) of the space shuttle vehicle orbiter in the AEDC VKF tunnel B (IA22), volume 2

Tabulated data of an experimental investigation are presented which was conducted in the AEDC/VKF Tunnel B to obtain interaction effects of RCS thruster jet plumes on SSV aerodynamics during staging to simulate RTLS abort. Interaction effects of the orbiter RCS thruster jet plumes on the orbiter and ET aerodynamics were investigated. RCS thruster jet plumes were simulated using both air and a 15 percent argon 85 percent helium gas mixture. The ET angle of attack range was -40 to +25 deg at sideslip angles of 0, 3, and 6 degrees. Orbiter angle of attack was varied from -15 to +10 degrees at sideslip angles of 0 and 3 deg. External tank full scale separation distances simulated were 0 to 1400 in. axially; 0 to 54 in. laterally; and a range of -100 to 1000 in. vertically. Data were also obtained on the ET in the interference-free flow field. Quiescent (no tunnel flow) thruster plume interaction data were obtained on the orbiter and orbiter-ET combination. Tests were conducted at Mach number 6 and a Reynolds number of 0.86 million per foot.

Daileda, J. J.↗
NASA NTRS1971

Static longitudinal aerodynamic characteristics of close-coupled wing-canard configurations at Mach numbers from 1.60 to 2.86

An experimental investigation was made in the Mach number range from 1.60 to 2.86 to determine the static longitudinal aerodynamic characteristics of close-coupled wing-canard configurations. Three canards, ranging in exposed planform area from 17.5 to 30.0 percent of the wing reference area, were employed in this investigation. The canards were either located in the plane of the wing or in a position 18.5 percent of the wing mean geometric chord above the wing plane. Most data obtained were for a model with a 60 deg leading-edge-sweep wing; however, a small amount of data were obtained for a 44 deg leading-edge-sweep wing. The model utilized two balances to isolate interference effects between wing and canard. In general, it was determined that at angle of attack for all configurations investigated with the canard in the plane of the wing an unfavorable interference exists which causes the additional lift on the canard generated by a canard deflection to be lost on the wing due to an increased downwash at the wing from the canard. Further, this interference decreased somewhat with increasing Mach number. Raising the canard above the plane of the wing also greatly decreased the interference of the canard deflection on the wing lift. However, at Mach 2.86 the presence of the canard in the high position had a greater unfavorable interference effect at high angles of attack than the canard in the wing plane. This interference resulted in the in-plane canard having better trimmed performance at Mach 2.86 for the same center-of-gravity location.

Dollyhigh, S. M.↗
NASA NTRS1979

Critical considerations for wind-tunnel testing V/STOL aircraft models

Low-speed wind-tunnel testing of V/STOL aircraft concepts to determine the aerodynamic-propulsion interaction effects during the transition between hover and wingborne flight is a necessary step in the development cycle of this type of aircraft. The paper examines factors which must be dealt with to assure that the information obtained in experiments is accurate and representative of the full-scale aircraft modeled. Areas of discussion include: (1) Proper engine simulation selection and model size constraints as a result of the selection; (2) selection of reference power-off condition for determining propulsion interaction effects; (3) modeling of realistic flight conditions to obtain a comprehensive evaluation of the entire transition performance; and (4) wind-tunnel boundary interferences and how these are highly configuration dependent.

Margason, R. J.↗
NASA NTRS1924

The Resistance of Spheres in Wind Tunnels and In Air

To supplement the standardization tests now in progress at several laboratories, a broad investigation of the resistance of spheres in wind tunnels and free air has been carried out by the National Advisory Committee for Aeronautics. The subject has been classed in aerodynamic research, and in consequence there is available a great mass of data from previous investigations. This material was given careful consideration in laying out the research, and explanation of practically all the disagreement between former experiments has resulted. A satisfactory confirmation of Reynolds law has been accomplished, the effect of means of support determined, the range of experiment greatly extended by work in the new variable density wind tunnel, and the effects of turbulence investigated by work in the tunnels and by towing and dropping tests in free air. It is concluded that the erratic nature of most of the previous work is due to support interference and differing turbulence conditions. While the question of support has been investigated thoroughly, a systematic and comprehensive study of the effects of scale and quality of turbulence will be necessary to complete the problem, as this phase was given only general treatment.

Bacon, D L↗
NASA NTRS1981

Propfan installation aerodynamics of a supercritical swept wing transport configuration

A semispan model with a powered propeller has been tested to provide data on the installation drag penalty of advanced propfan-powered transports designed to cruise at a Mach number of 0.8. These tests, conducted in 14-foot and 11-foot transonic wind tunnels, are a part of a NASA program to develop efficient, high-speed propellers for more fuel-efficient commercial transports for the 1990s and beyond. The model is instrumented for measuring propeller forces, wing/nacelle forces and moments, and pressure distributions over the wing and nacelle. The body in these tests was nonmetric, being connected to the wing by an RTV seal at the wing/body juncture. Tests were run at angles of attack from -3 to +5 deg over the Mach number range 0.6 to 0.85 at a Reynolds number of about 9,000,000. Results of these tests indicate that the nacelle interference drag can be quite large relative to an uninstalled nacelle. However, the losses due to the nacelle were reduced to acceptable levels by changes to the wing leading edge and nacelle intersection. The propeller slipstream causes substantial changes in the wing span load distribution indicating that twist modifications are needed to recover a more favorable span load distribution.

Smith, R. C.↗
NASA NTRS1976

Upper Surface Nacelle Influence on SCAR Aerodynamic Characteristics at Transonic Speeds

The arrow-wing transport configuration with detached engines located over the wing to produce upper surface exhaust flow effects was tested at angles of attack from -4 deg to 8 deg and jet total-pressure ratios from 1 (Jet off) to approximately 10. Wing tip leading edge flap deflections of -10 deg to 10 deg were tested with the wing-body configuration only (no nacelles). Tests were made with various nacelle chordwise, spanwise, and vertical height locations over the Mach number, angle of attack, and jet total-pressure ratio ranges. Deflecting the wing tip leading edge flap from 0 deg to -10 deg increased maximum lift to drag ratio by 1.0 at subsonic speeds. Installation of upper surface nacelles (no wing/nacelle pylons) increased the wing-body pitching moment at all Mach numbers and decreased the drag of the wing-body configuration at subsonic Mach numbers. Jet exhaust interference effects were negligible.

Mercer, C. E.↗
NASA NTRS1982

Theoretical evaluation of a V/STOL fighter model utilizing the PAN AIR code

The PAN AIR computer code was investigated as a tool for predicting closely coupled aerodynamic and propulsive flowfields of arbitrary configurations. The NASA/Ames V/STOL fighter model, a configuration of complex geometry, was analyzed with the PAN AIR code. A successful solution for this configuration was obtained when the nozzle exit was treated as an impermeable surface and no wakes were included around the nozzle exit. When separated flow was simulated from the end of the nacelle, requiring the use of wake networks emanating from the nozzle exit, a number of problems were encountered. A circular body nacelle model was used to investigate various techniques for simulating the exhaust plume in PAN AIR. Several approaches were tested and eliminated because they could not correctly simulate the interference effects. Only one plume modeling technique gave good results. A PAN AIR computation that used a plume shape and inflow velocities obtained from the Navier-Stokes solution for the plume produced results for the effects of power that compared well with experimental data.

Howell, G. A.↗
NASA NTRS1986

Installation aerodynamics of wing-mounted, single-rotation propfans

The installation of advanced (M = 0.8) turboprop propulsion systems on transport aircraft represents a challenging task to design engineers. The installation aerodynamics of wing-mounted, single-rotation, tractor turboprop systems have been extensively investigated by NASA Ames using a large powered semispan wind tunnel model. Two configurations have been studied: a straight under-the-wing (UTW) nacelle and a contoured over-the-wing (OTW) configuration. The installation characteristics of these two configurations are presented in terms of installed drag, wing pressure distributions, and surface oil flows. Through the use of wing leading-edge modifications, the installed drag of the UTW nacelle was reduced to less than isolated nacelle drag at the cruise condition of M = 0.8 with a wing C(L) of 0.5. At this condition, the favorable interference is attributed to the recovery of a portion of the swirl in the slipstream generated by the single rotation propeller. The OTW installation represented a much more difficult task. For the configuration tested, the installed drag was unacceptably high and demonstrated the sensitive nature of installing a nacelle/slipstream combination in the transonic flow regime.

Bencze, Daniel P.↗
NASA NTRS2006

Space Shuttle Body Flap Actuator Bearing Testing for NASA Return to Flight

The Space Shuttle body flap (BF) is located beneath the main engine nozzles and is required for proper aerodynamic control during orbital descent. The body flap is controlled by four actuators connected by a common shaft and driven by the hydraulic power drive unit. Inspection of the actuators during refurbishment revealed three shaft bearings with unexpected damage. One was coated with black oxide on the balls and race wear surfaces, a second contained a relatively deep wear scar, and the third with scratches and an aluminum particle in the wear track. A shaft bearing life test program was initiated to measure the wear life and explain the 5.08-micrometer wear scar. A tribological analysis was conducted to demonstrate that the black oxide coated wear surfaces did not damage the bearing, interfere with the lubrication, or cause severe bearing wear. Pre-damaged (equivalent of 30 missions), commercial equivalent bearings and previously flown shaft bearings were tested at axial loads, speeds, and temperatures seen during flight operations. These bearing were successfully life tested at 60 C for 24 hours or 90 flights. With a safety factor of 4X, the bearings were qualified for 22 flights when only a maximum of 12 flights are expected. Additional testing at 23 C was performed to determine the lubricant life and to further understand the mechanism that caused the blackened balls. Test results indicating bearing life was shortened at a lower temperature surprised the investigators. Start\Stop bearing testing that closely simulates mission profile was conducted at 23 C. Results of this testing showed lubricant life of 12 flights including a safety factor of four. Additional testing with bearings that have the equivalent of 30 missions of damage is being tested at 23 C. These tests are being performed over the Shuttle load profile to demonstrate the residual bearing life in the actuators exceeds 12 missions. Testing showed that the end of the shaft bearing life was characterized by bearing temperature rise, preload drop, and the onset of a severe wear bearing failure mechanism. The severe wear failure mechanism is characterized by rough wear scars, extensive bearing wear and steel transfer between the balls and the races.

Jet, Timothy R.↗
NASA NTRS1974

Experimental investigation at Mach 8 of the effects of projections and cavities on heat transfer to a model of the Viking aeroshell

An experimental investigation into the aerodynamic heating on a Mars entry vehicle shape with several types of local surface distortion is presented. The configurations tested were 0.033-scale models of a spherically blunted 70 deg half-angle cone with two protuberances of different length, representing the tube leading to the gas chromograph mass spectrometer, and two aeroshell-bioshield attachment points of different size. These models were tested at free-stream Reynolds numbers per meter of 3.7 million and 17 million over an angle-of-attack range from 0 to 18 deg in the Langley Mach 8 variable density hypersonic tunnel. The phase change-coating technique was used to measure heat transfer coefficient. The long protuberance caused more severe interference heating than the short protuberance for the same conditions. When the short projection was located close to the edge of the aeroshell, the interference heating was greater than that on the same projection when located near the vertex. A significant increase in heat transfer coefficient was measured only on the larger aeroshell-bioshield attachment point.

Creel, T. R., Jr.↗
NASA NTRSDate not supplied

Fundamental Noise Characterization of a Ducted Propeller in Hover

Unmanned aerial vehicles (UAVs) are currently being used for reconnaissance missions, tactical surveillance, and infrastructure inspection. When legislation allows it, these devices will provide additional services close to inhabited areas, which could lead to noise complaints. On most UAVs, the propellers are the dominant source of noise. As a result, researchers have studied the impact of propeller shape and blade count on noise. Much of this work, however, has focused on isolated propellers. While different UAV concepts are equipped with ducts for aerodynamic and protection reasons, few studies focus on the acoustic benefit of ducts, as is achieved, for example, on turbofan aircraft. The objectives of this paper are: first, to simulate the noise radiation of a UAV propeller in static conditions based on its location in a hard wall duct; second, to analyze the contribution of the different acoustic source components (i.e., thrust, torque, and thickness); and third, to validate the approach with experiments conducted in the NASA Small Hover Anechoic Chamber over a range of propeller rotation rates. For the propeller and duct geometries considered in this experiment, it is shown that the best attenuation is achieved when the propeller is centered axially in the duct because of interference between upstream and downstream radiated waves.

ducted propeller↗
NASA NTRS1983

Influence of airplane components on rotational aerodynamic data for a typical single-engine airplane

The influence of airplane components, as well as wing location and tail length, on the rotational flow aerodynamics is discussed for a 1/6-scale general aviation airplane model. Examination of the individual component effects provides insight as to the source of the moment and forces that determine spin characteristics. For the subject airplane, it is seen that the presence of the horizontal tail adversely effects the yaw damping of the vertical tail such that at certain angles of attack the vertical actually becomes a propelling component in spinning motion. It is shown that the level of this adverse interaction is a function of the relative positions of the wing and the tail surfaces, e.g., relocating the wing to the high wing position or shortening the tail moment arm from the basic configuration both reduced or eliminated the adverse tail interference for certain angles of attack. The influence of wing location on the damping characteristics of the horizontal-vertical tail combination would not be addressed in any existing tail design criterion for spinning and could only be discerned through rotary balance testing.

Ralson, J.↗
NASA NTRS1989

Influence of thickness and camber on the aeroelastic stability of supersonic throughflow fans: An engineering approach

An engineering approach was used to include the nonlinear effects of thickness and camber in an analytical aeroelastic analysis of cascades in supersonic acial flow (supersonic leading-edge locus). A hybrid code using Lighthill's nonlinear piston theory and Lanes's linear potential theory was developed to include these nonlinear effects. Lighthill's theory was used to calculate the unsteady pressures on the noninterference surface regions of the airfoils in cascade. Lane's theory was used to calculate the unsteady pressures on the remaining interference surface regions. Two airfoil profiles was investigated (a supersonic throughflow fan design and a NACA 66-206 airfoil with a sharp leading edge). Results show that compared with predictions of Lane's potential theory for flat plates, the inclusion of thickness (with or without camber) may increase or decrease the aeroelastic stability, depending on the airfoil geometry and operating conditions. When thickness effects are included in the aeroelastic analysis, inclusion of camber will influence the predicted stability in proportion to the magnitude of the added camber. The critical interblade phase angle, depending on the airfoil profile and operating conditions, may also be influenced by thickness and camber. Compared with predictions of Lane's linear potential theory, the inclusion of thickness and camber decreased the aerodynamic stifness and increased the aerodynamic damping at Mach 2 and 2.95 for a cascade of supersonic throughflow fan airfoils oscillating 180 degrees out of phase at a reduced frequency of 0.1.

Ramsey, John K.↗
NASA NTRS1991

Influence of thickness and camber on the aeroelastic stability of supersonic throughflow fans

An engineering approach was used to include the nonlinear effects of thickness and camber in an analytical aeroelastic analysis of casecades in supersonic acial flow (supersonic leading-edge locus). A hybrid code using Lighthill's nonlinear piston theory and Lane's linear potential theory was developed to include these nonlinear effects. Lighthill's theory was used to calculate the unsteady pressures on the noninterference surface regions of the airfoils in cascade. Lane's theory was used to calculate the unsteady pressures on the remaining interference surface regions. Two airfoil profiles were investigated (a supersonic throughflow fan design and a NACA 66-206 airfoil with a sharp leading edge). Results show that compared with predictions of Lane's potential theory for flat plates, the inclusion of thickness (with or without camber) may increase or decrease the aeroelastic stability, depending on the airfoil geometry and operating conditions. When thickness effects are included in the aeroelastic analysis, inclusion of camber will influence the predicted stability in proportion to the magnitude of the added camber. The critical interblade phase angle, depending on the airfoil profile and operating conditions, may also be influenced by thickness and camber. Compared with predictions of Lane's linear potential theory, the inclusion of thickness and camber decreased the aerodynamic stiffness and increased the aerodynamic damping at Mach 2 and 2.95 for a cascade of supersonic throughflow fan airfoils oscillating 180 degrees out of phase at a reduced frequency of 0.1.

Ramsey, John K.↗