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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 253 records · Page 14

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NASA NTRS1991

Turbulent boundary layer separation over a rearward facing ramp and its control through mechanical excitation

A vane which oscillates about a fixed point at the inlet to a two-dimensional, 20-deg rearward-facing ramp has been noted to effectively delay the separation of a turbulent boundary layer. A parametric study has been conducted to determine the variation of the maximum values of the static pressure recovery, measured close to the ramp exit as a function of the oscillation frequency of the vane. The resulting curves show a peak value in the maximum static pressure recovery as a function of the vane's oscillation frequency. For the excited cases, corona anemometer measurements in the inner wall regions of the boundary layer along the surface of the ramp show a large range of unsteadiness in the local velocities. Finally, a qualitative explanation of the phenomena of delay of separation is discussed which appears to be supported by the surface static pressure and flow field measurements obtained.

Mckinzie, Daniel J., Jr.↗

NASA NTRS2005

Experimental Investigation of a Morphing Nacelle Ducted Fan

The application of Circulation Control to the nacelle of a shrouded fan is proposed as a means to enhance off-design performance of the shrouded fan. Typically, a fixed geometry shroud is efficient at a single operating condition. Modifying circulation about the fixed geometry is proposed as a means to virtually morph the shroud without moving surfaces. This approach will enhance off-design-point performance with minimal complexity, weight, and cost. Termed the Morphing Nacelle, this concept provides an attractive propulsion option for Vertical Take-off and Landing (VTOL) aircraft, such conceptual Personal Air Vehicle (PAV) configurations proposed by NASA. An experimental proof of concept investigation of the Morphing Nacelle is detailed in this paper. A powered model shrouded fan model was constructed with Circulation Control (CC) devices integrated in the inlet and exit of the nacelle. Both CC devices consisted of an annular jet slot directing a jet sheet tangent to a curved surface, generally described as a Coanda surface. The model shroud was tailored for axial flight, with a diffusing inlet, but was operated off-design condition as a static lifting fan. Thrust stand experiments were conducted to determine if the CC devices could effectively improve off-design performance of the shrouded fan. Additional tests were conducted to explore the effectiveness of the CC devices a means to reduce peak static pressure on the ground below a lifting fan. Experimental results showed that off-design static thrust performance of the model was improved when the CC devices were employed under certain conditions. The exhaust CC device alone, while effective in diffusing the fan exhaust and improving weight flow into shroud inlet, tended to diminish performance of the fan with increased CC jet momentum. The inlet CC device was effective at reattaching a normally stalled inlet flow condition, proving an effective means of enhancing performance. A more dramatic improvement in static thrust was obtained when the inlet and exit CC devices were operated in unison, but only over a limited range of CC jet momentum. Operating the nacelle inlet and exit CC devices together proved very effective in reducing peak ground plane static pressure, while maintaining static thrust. The Morphing Nacelle concept proved effective at enhancing off-design performance of the model; however, additional investigation is necessary to generalize the results.

Kondor, Shayne A.↗

NASA NTRS2020

Full Aerothermal Characterization of the 6- by 9-Foot Icing Research Tunnel (2019 Test)

A full aerothermal characterization of the NASA Glenn Icing Research Tunnel (IRT) was completed in August 2019. A mandatory maintenance period required the test entry to be separated into two phases: the first occurring in April 2019 and the second in August 2019. The hardware used during this test entry consisted of the 6-foot quick-check rake, 6-foot traversing rake, RTD array, and hot-wire rake. The 6-foot quick check rake collects static pressure, total pressure, total temperature, and flow angularity data along the centerline of the test section. The data from the 6-foot quick-check rake were used to generate test section calibration relationships for total and static pressure as well as implement statistical process control. The 6-foot traversing rake collects flow uniformity data across the test section for various flow parameters, such as static pressure, total pressure, and flow angularity. Total temperature flow uniformity data is mapped by the RTD array and is used to generate the test section calibration relationship for total temperature. The hot-wire rake measures turbulence intensity at three lateral locations and characterizes airspeed unsteadiness in the test section. The test entry resulted in updated calibration relationships and a full characterization of test section flow quality and flow uniformity in the IRT.

Wind↗

NASA NTRS1953

Performance Characteristics of Plane-Wall Two-Dimensional Diffusers

Experiments have been made at Stanford University to determine the performance characteristics of plane-wall, two-dimensional diffusers which were so proportioned as to insure reasonable approximation of two-dimensional flow. All of the diffusers had identical entrance cross sections and discharged directly into a large plenum chamber; the test program included wide variations of divergence angle and length. During all tests a dynamic pressure of 60 pounds per square foOt was maintained at the diffuser entrance and the boundary layer there was thin and fully turbulent. The most interesting flow characteristics observed were the occasional appearance of steady, unseparated, asymmetric flow - which was correlated with the boundary-layer coalescence - and the rapid deterioration of flow steadiness - which occurred as soon as the divergence angle for maximum static pressure recovery was exceeded. Pressure efficiency was found to be controlled almost exclusively by divergence angle, whereas static pressure recovery was markedly influenced by area ratio (or length) as well as divergence angle. Volumetric efficiency. diminished as area ratio increased, and at a greater rate with small lengths than with large ones. Large values of the static-pressure-recovery coefficient were attained only with long diffusers of large area ratio; under these conditions pressure efficiency was high and. volumetric efficiency low. Auxiliary tests with asymmetric diffusers demonstrated that longitudinal pressure gradient, rather than wall divergence angle, controlled flow separation. Others showed that the addition of even a short exit duct of uniform section augmented pressure recovery. Finally, it was found that the installation of a thin, central, longitudinal partition suppressed flow separation in short diffusers and thereby improved pressure recovery

Reid, Elliott G↗

NASA NTRS1973

Theoretical and experimental internal flow characteristics of a 13.97-centimeter-diameter inlet at STOL takeoff and approach conditions

The theoretical and experimental internal flow characteristics of a 13.97-cm-diam inlet with centerbody retracted and extended are presented at STOL takeoff and approach operating conditions. The theoretical results were obtained from incompressible potential flow corrected for compressibility and boundary layer. Comparisons between theoretical internal surface static-pressure distributions and experimental data are presented for free-stream velocities of 0, 24, 32, and 45 m/sec for a range of inlet incidence angles from 0 to 50 deg. Surface static-pressure distributions are illustrated at circumferential locations of 0, 60, 120, and 180 deg. Surface Mach number distributions from the stagnation point to the diffuser exit are presented along with turbulent boundary-layer shape factors. In general, good agreement was found between the theoretical and experimental surface static pressure distributions.

Albers, J. A.↗

NASA NTRS1979

Mean velocity and decay characteristics of the near- and far-wake of a compressor rotor blade of moderate loading

This paper reports the experimental study of the three-dimensional characteristics of the mean velocity in the wake of a moderately loaded compressor rotor blade. The measurements were taken with a three-sensor hot-wire probe rotating with the rotor. The wake was surveyed at several radial and axial stations. The loading was found to have substantial effect and this was reflected not only in the axial and tangential components, but also in the radial component. The radial velocities were found to be high very near the trailing-edge and this exhibits the characteristics prevalent in a trailing vortex system. The static pressures across the wake were measured using a direction insensitive spherical head static-stagnation pressure probe. The static pressure was found to be higher inside the wake. These and other measurements are reported and correlated in this paper.

Ravindranath, A.↗

NASA NTRS1991

Cold flow testing of the Space Shuttle Main Engine high pressure fuel turbine model

In order to experimentally determine the performance of the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP) turbine, a 'cold' air flow turbine test program was established at NASA's Marshall Space Flight Center. As part of this test program, a baseline test of Rocketdyne's HPFTP turbine has been completed. The turbine performance and turbine diagnostics such as airfoil surface static pressure distributions, static pressure drops through the turbine, and exit swirl angles were investigated at the turbine design point, over its operating range, and at extreme off-design points. The data was compared to pretest predictions with good results. The test data has been used to improve meanline prediction codes and is now being used to validate various three-dimensional codes. The data will also be scaled to engine conditions and used to improve the SSME steady-state performance model.

Hudson, Susan T.↗

NASA NTRS1972

Comparison of Predicted and Experimental Heat-Transfer and Pressure-Drop Results for an Air-Cooled Plug Nozzle and Supporting Struts

A calculational procedure is presented to analyze the heat-transfer and fluid-flow characteristics of a convectively air -cooled plug-nozzle operating on an afterburning turbojet engine. Anderson's method was used to predict hot-gas static pressures in the supersonic stream with fully expanded flow (high nozzle-pressure ratios); the results were excellent. For low nozzle-pressure ratios, the flow was assumed to expand one-dimensionally and isentropically to the plug back pressure. Wall temperatures predicted using this latter pressure distribution agreed well with the wall temperatures predicted using the measured hot-gas pressures (maximum deviation was about 30 K (54 deg R)). Either an in tegral boundary-layer technique or a simple pipe-flow equation may be used to calculate convective heat transfer from the hot gas to the wall. The simple pipeflow equation results in the prediction of slightly higher wall temperatures than does the integral technique. Experimental wall temperatures were generally in good agreement with the two predicted wall temperature distributions. Excellent agreement was noted b etween measured and predicted coolant static-pressure distributions. The plug-coolant temperature rise was generally overpredicted by about 22.2 K (40 deg R); possible explanations are offered. Although an an alysis of the struts, which support the plug, was purposely kept simple, reasonable results were obtained. Potential flow over an ellipse was used to calculate hot-gas static pressure; the results were satisfactory.

Graber, E. J., Jr.↗

NASA NTRS1975

Shock position sensor for supersonic inlets

Static pressure taps or ports are provided in the throat of a supersonic inlet, and signals indicative of the pressure at each of the ports is fed to respective comparators. Means are also provided for directing a signal indicative of the total throat pressure to the comparators. A periodic signal is superimposed on the total throat pressure so that the signal from the static pressure tabs is compared to a varying scan signal rather than to total throat pressure only. This type of comparison causes each comparator to provide a pulse width modulated output which may vary from 0% 'time on' to 100% 'time on'. The pulse width modulated outputs of the comparators are summed, filtered, and directed to a controller which operates a bypass valve such as a door whereby air is dumped from the inlet to prevent the shock wave from being expelled out the front.

Dustin, M. O.↗

NASA NTRS1978

A simplified gross thrust computing technique for an afterburning turbofan engine

A simplified gross thrust computing technique extended to the F100-PW-100 afterburning turbofan engine is described. The technique uses measured total and static pressures in the engine tailpipe and ambient static pressure to compute gross thrust. Empirically evaluated calibration factors account for three-dimensional effects, the effects of friction and mass transfer, and the effects of simplifying assumptions for solving the equations. Instrumentation requirements and the sensitivity of computed thrust to transducer errors are presented. NASA altitude facility tests on F100 engines (computed thrust versus measured thrust) are presented, and calibration factors obtained on one engine are shown to be applicable to the second engine by comparing the computed gross thrust. It is concluded that this thrust method is potentially suitable for flight test application and engine maintenance on production engines with a minimum amount of instrumentation.

Hamer, M. J.↗

NASA NTRS1991

Progress toward synergistic hypermixing nozzles

Mean flow measurements have been obtained for air-to-air mixing downstream of swept and unswept ramp wall-mounted hypermixing nozzle configurations. Aside from the sweep of the ramps, the two nozzle configurations investigated are identical. The nozzles inject three parallel supersonic jets at a 15-deg angle (relative to the wind tunnel wall) into a supersonic freestream. Mach number and volume fraction distributions in a transverse plane 11.1 nozzle heights downstream from the nozzle exit plane were measured. Data are presented for a freestream Mach number of three at a matched static pressure condition and also at an underexpanded static pressure condition (pressure ratio equal to 5). Surface oil flow visualization was used to investigate the near-wall flow behavior. The results indicate that the swept ramp injectors produce stronger and larger vortex pairs than the unswept ramp injectors. The increased interaction between the swept ramp model's larger vortex pairs yields better mixing characteristics for this model.

Davis, D. O.↗

NASA NTRS1947

Low-Speed Wind-Tunnel Tests of a Pilotless Aircraft Having Horizontal and Vertical Wings and Cruciform Tail

Low-speed tests of a pilotless aircraft were conducted in the Langley propeller-research tunnel to provide information for the estimation of the longitudinal stability and. control, to measure the aileron effectiveness, and to calibrate the radome and the Machmeter pitot-static orifices. It was found that the model possessed a stEb.le variation of elevator angle required for trim throughout the speed range at the design angle of attack. A comparison of the airplane with and without JATO units and with an alternate rocket booster showed that a large loss in longitudinal stability and control resulting from the addition of the rocket booster to the aircraft was sufficient to make the rocket-booster assembly unsatisfactory as an alternate for the JATO units. Reversal of the aileron effectiveness was evident at positive deflections of the vertical wing flap indicating that the roll-stabilization system would produce roiling moments in a tight right turn contrary to its design purpose. Vertical-wing-flap deflections caused large errors in the static-pressure reading obtained by the original static-tube installation. A practical installation point on the fuselage was located which should yield reliable measurement of the free-stream static pressure.

Mastrocola, N↗

NASA NTRS1956

Preliminary Report on Experimental Investigation of Engine Dynamics and Controls for a 48-inch Ram-jet Engine

Engine dynamics and controls data are presented for a ram-jet engine which was operated in a free-jet facility at a Mach number of 2.76 and altitudes from 68,000 to 82,000 feet. The predominant engine dynamic characteristics was dead time, with values ranging from 0.018 to 0.053 sec. The control systems were designed to hold a constant ratio of a diffuser static pressure to an inlet-cone reference static pressure. Response times and percent overshoot to fuel-flow disturbances that would have caused subcritical operation without control are included for a wide range of control settings. For all settings, the control response was fast enough (0.08 to 0.5 sec) to prevent subcritical operation.

Vasu, George↗

NASA NTRS2017

Computational and Experimental Characterization of the Mach 6 Facility Nozzle Flow for the Enhanced Injection and Mixing Project at NASA Langley Research Center

Computational fluid dynamics analyses and experimental data are presented for the Mach 6 facility nozzle used in the Arc-Heated Scramjet Test Facility for the Enhanced Injection and Mixing Project (EIMP). This project, conducted at the NASA Langley Research Center, aims to investigate supersonic combustion ramjet (scramjet) fuel injection and mixing physics relevant to flight Mach numbers greater than 8. The EIMP experiments use a two-dimensional Mach 6 facility nozzle to provide the high-speed air simulating the combustor entrance flow of a scramjet engine. Of interest are the physical extent and the thermodynamic properties of the core flow at the nozzle exit plane. The detailed characterization of this flow is obtained from three-dimensional, viscous, Reynolds-averaged simulations. Thermodynamic nonequilibrium effects are also investigated. The simulations are compared with the available experimental data, which includes wall static pressures as well as in-stream static pressure, pitot pressure and total temperature obtained via in-stream probes positioned just downstream of the nozzle exit plane.

Drozda, Tomasz G.↗

NASA NTRS1992

Crossing shock wave turbulent boundary layer interactions - Variable angle and shock generator length geometry effects at Mach 3

By comparing the detailed wall static pressure distributions for 9 inch and 11 inch long fins generating a crossing shock configuration at M = 2.93, the high resolution results of the 9 inch fins are shown to be free of exit effects. Analysis of the static pressure profiles have delineated the limited regions where the single fin results are valid. The characteristics of the complex interaction, with varying shock wave strength, have been described. The data provide a critical test for computational fluid dynamics which, in its initial phase, has performed poorly in predicting the measured wall static pressure distributions.

Bogdonoff, S. M.↗

NASA NTRS1948

An Investigation of the McDonnell XP-85 Airplane in the Ames 40- by 80-Foot Wind Tunnel: Pressure-Distribution Tests

Pressure measurements were made during wind-tunnel tests of the McDonnell XP-85 parasite fighter. Static-pressure orifices were located over the fuselage nose, over the canopy, along the wing root, and along the upper and lower stabilizer roots. A total-pressure and static-pressure rake was located in the turbojet engine air-intake duct. It was installed at the station where the compressor face would be located. Pressure data were obtained for two airplane conditions, clean and with skyhook extended, through a range of angle of attack and a range of yaw.

Hunton, Lynn W.↗

NASA NTRS1980

System for use in conducting wake investigation for a wing in flight

A system supported by a wing in flight is described which has a reference total pressure port in spaced relation with a wake as the wake is generated by the wing, a reference static pressure port supported in spaced relation with the wake, and a probe adapted to be displaced along an accurate path through the wake including a total pressure port and static pressure ports. A differential pressure transducer and a pressure switching device are interposed between the ports and the transducer is provided for selectively connecting pairs of the ports to the transducer in opposed relation, whereby a single transducer is utilized to obtain differential pressure measurement for the wake with enhanced accuracy.

Bikle, P. F.↗

NASA NTRS1984

Fuselage and nozzle pressure distributions of a 1/12-scale F-15 propulsion model at transonic speeds. Effect of fuselage modifications and nozzle variables

Static pressure coefficient distributions on the forebody, afterbody, and nozzles of a 1/12 scale F-15 propulsion model was determined in the 16 foot transonic tunnel for Mach numbers from 0.60 to 1.20, angles of attack from -2 deg to 7 deg and ratio of jet total pressure to free stream static pressure from 1 up to about 7, depending on Mach number. The effects of nozzle geometry and horizontal tail deflection on the pressure distributions were investigated. Boundary layer total pressure profiles were determined at two locations ahead of the nozzles on the top nacelle surface. Reynolds number varied from about 1.0 x 10 to the 7th power per meter, depending on Mach number.

Pendergraft, O. C., Jr.↗