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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 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Some Observations on Transitory Stall in Conical Diffusers

Results from an experimental investigation on the flow through conical diffusers are presented. The mean and fluctuating velocity fields are compared for three diffusers with total diffusion angles of 16, 20 and 24 degrees, in the throat Mach number (M sub t) range of 0.05 to 0.95. Each of the diffusers were 14 cm long and had a 5.08 cm inlet diameter, and the flow exited into the ambient. The boundary layer at the throat was thin with the throat diameter (D sub t) to momentum thickness (O) ratio being as high as 800 at M(sub t) = 0.4. While the 16 deg diffuser flow exited with a top-hat mean velocity profile, increasing losses due to increasing separation resulted in fuller profiles for the 20 and 24 degree cases. A detailed flow field study was conducted for the 16 deg. diffuser. The u'-spectrum, measured at the exit plane, exhibited a peak apparently due to the ensuing jet column instability throughout the M(sub t) range covered. In addition, a much lower frequency spectral peak also occurred in the M(sub t) range of 0.3 to 0.7. Both of the spectral peaks were due to axisymmetric flow fluctuations. A self-sustaining flow oscillation occurred in the M(sub t) range of 0.6 to 0.85, emitting a loud tone, when the jet column instability frequency matched the resonance frequency of the diffuser. Limited data showed that artificial acoustic excitation was effective in reducing the flow fluctuations, with a resultant increase in the pressure recovery, at low M(sub t).

Zaman, K. B. M. Q.↗
NASA NTRS1990

Some observations on transitory stall in conical diffusers

Results from an experimental investigation on the flow through conical diffusers are presented. The mean and fluctuating velocity fields are compared for three diffusers with total diffusion angles of 16, 20 and 24 degrees, in the throat Mach number (M sub t) range of 0.05 to 0.95. Each of the diffusers were 14 cm long and had a 5.08 cm inlet diameter, and the flow exited into the ambient. The boundary layer at the throat was thin with the throat diameter (D sub t) to momentum thickness (O) ratio being as high as 800 at M(sub t) = 0.4. While the 16 deg diffuser flow exited with a top-hat mean velocity profile, increasing losses due to increasing separation resulted in fuller profiles for the 20 and 24 degrees cases. A detailed flow field study was conducted for the 16 deg diffuser. The u'-spectrum, measured at the exit plane, exhibited a peak apparently due to the ensuing jet column instability throughout the M(sub t) range covered. In addition, a much lower frequency spectral peak also occurred in the M(sub t) range of 0.3 to 0.7. Both of the spectral peaks were due to axisymmetric flow fluctuations. A self-sustaining flow oscillation occurred in the M(sub t) range of 0.6 to 0.85, emitting a loud tone, when the jet column instability frequency matched the resonance frequency of the diffuser. Limited data showed that artificial acoustic excitation was effective in reducing the flow fluctuations, with a resultant increase in the pressure recovery, at low M(sub t).

Zaman, K. B. M. Q.↗
NASA NTRS1990

Integrated Flight-propulsion Control Concepts for Supersonic Transport Airplanes

Integration of propulsion and flight control systems will provide significant performance improvements for supersonic transport airplanes. Increased engine thrust and reduced fuel consumption can be obtained by controlling engine stall margin as a function of flight and engine operating conditions. Improved inlet pressure recovery and decreased inlet drag can result from inlet control system integration. Using propulsion system forces and moments to augment the flight control system and airplane stability can reduce the flight control surface and tail size, weight, and drag. Special control modes may also be desirable for minimizing community noise and for emergency procedures. The overall impact of integrated controls on the takeoff gross weight for a generic high speed civil transport is presented.

Burcham, Frank W., Jr.↗
NASA NTRS1991

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

A vane oscillating about a fixed point at the inlet to a two-dimensional 20 degree rearward facing ramp has proven effective in delaying the separation of a turbulent boundary layer. Measurements of the ramp surface static pressure coefficient obtained under the condition of vane oscillation and constant inlet velocity revealed that two different effects occurred with surface distance along the ramp. In the vicinity of the oscillating vane, the pressure coefficients varied as a negative function of the vane's trailing edge rms velocity; the independent variable on which the rms velocity depends are the vane's oscillation frequency and its displacement amplitude. From a point downstream of the vane to the exit of the ramp; however, the pressure coefficient varied as a more complex function of the two independent variables. That is, it was found to vary as a function of the vane's oscillation frequency throughout the entire range of frequencies covered during the test, but over only a limited range of the trailing edge displacement amplitudes covered. More specifically, the value of the pressure coefficient was independent of increases in the vane's displacement amplitude above approximately 35 inner wall units of the boundary layer. Below this specific amplitude it varied as a function of the vane's trailing edge rms velocity. This height is close to the upper limit of the buffer layer. A parametric study was made to determine the variation of the maximum static pressure recovery as a function of the vane's oscillation frequency, for several ramp inlet velocities and a constant displacement amplitude of the vane's trailing edge. The results indicate that the phenomenon producing the optimum delay of separation may be Strouhal number dependent. Corona anemometer measurements obtained in the inner wall regions of the boundary layer for the excited case reveal a large range of unsteadiness in the local velocities. These measurements imply the existence of inflections in the profiles, which provide a mechanism for resulting inviscid flow instabilities to produce turbulence in the near wall region, thereby delaying separation of the boundary layer.

Mckinzie, Daniel J., Jr.↗
NASA NTRS1990

Hot gas ingestion test results of a two-poster vectored thrust concept with flow visualization in the NASA Lewis 9- x 15-foot low speed wind tunnel

A 9.2 percent scale Short Takeoff and Vertical Landing (STOVL) hot gas ingestion model was designed and built by McDonnell Douglas Corporation (MCAIR) and tested in the Lewis Research Center 9 x 15 foot Low Speed Wind Tunnel (LSWT). Hot gas ingestion, the entrainment of heated engine exhaust into the inlet flow field, is a key development issure for advanced short takeoff and vertical landing aircraft. Flow visualization from the Phase 1 test program, which evaluated the hot ingestion phenomena and control techniques, is covered. The Phase 2 test program evaluated the hot gas ingestion phenomena at higher temperatures and used a laser sheet to investigate the flow field. Hot gas ingestion levels were measured for the several forward nozzle splay configurations and with flow control/life improvement devices (LIDs) which reduced the hot gas ingestion. The model support system had four degrees of freedom - pitch, roll, yaw, and vertical height variation. The model support system also provided heated high-pressure air for nozzle flow and a suction system exhaust for inlet flow. The test was conducted at full scale nozzle pressure ratios and inlet Mach numbers. Test and data analysis results from Phase 2 and flow visualization from both Phase 1 and 2 are documented. A description of the model and facility modifications is also provided. Headwind velocity was varied from 10 to 23 kn. Results are presented over a range of nozzle pressure ratios at a 10 kn headwind velocity. The Phase 2 program was conducted at exhaust nozzle temperatures up to 1460 R and utilized a sheet laser system for flow visualization of the model flow field in and out of ground effects. The results reported are for nozzle exhaust temperatures up to 1160 R. These results will contain the compressor face pressure and temperature distortions, the total pressure recovery, the inlet temperature rise, and the environmental effects of the hot gas. The environmental effects include the ground plane contours, the model airframe heating, and the location of the ground flow separation.

Johns, Albert L.↗
NASA NTRS1991

Design limits of compressible NLF airfoils

There has been considerable success in the design of practical low-speed natural laminar flow (NLF) airfoils, with significant profile drag reduction from conventional turbulent flow airfoils. The favorable lowspeed results give an incentive to explore the possibilities of high-subsonic speed NLF airfoil design. The design problem at higher freestream Mach numbers is more severe than for low-speed designs because these high-speed airplanes typically fly at higher chord Reynolds numbers than low-speed NLF airplanes. As the Mach number increases, the main priority changes from the use of sufficient acceleration to achieve NLF for low-drag, to the delay of separation in the far aft pressure recovery region. In this effort, NLF airfoils have been designed for Mach numbers ranging from 0.60 to 0.80 and chord Reynolds numbers of 30 x 10 to the 6th and 40 x 10 to the 6th.

Viken, Jeff↗
NASA NTRS1990

Hot gas ingestion test results of a two-poster vectored thrust concept with flow visualization in the NASA Lewis 9- by 15-foot low speed wind tunnel

A 9.2 percent scale STOVL hot gas ingestion model was tested in the NASA Lewis 9 x 15-foot Low-Speed Wind Tunnel. Flow visualization from the Phase 1 test program, which evaluated the hot ingestion phenomena and control techniques, is covered. The Phase 2 test program evaluated the hot gas ingestion phenomena at higher temperatures and used a laser sheet to investigate the flow field. Hot gas ingestion levels were measured for the several forward nozzle splay configurations and with flow control/life improvement devices (LIDs) which reduced the hot gas ingestion. The test was conducted at full scale nozzle pressure ratios and inlet Mach numbers. Results are presented over a range of nozzle pressure ratios at a 10 kn headwind velocity. The Phase 2 program was conducted at exhaust nozzle temperatures up to 1460 R and utilized a sheet laser system for flow visualization of the model flow field in and out of ground effects. The results reported are for nozzle exhaust temperatures up to 1160 R and contain the compressor face pressure and temperature distortions, the total pressure recovery, the inlet temperature rise, and the environmental effects of the hot gas. The environmental effects include the ground plane contours, the model airframe heating, and the location of the ground flow separation.

Johns, Albert L.↗
NASA NTRS1990

Integrated flight-propulsion control concepts for supersonic transport airplanes

Increased powerplant thrust and reduced fuel consumption are obtainable by controlling engine stall margins as a function of flight and engine operating conditions. An evaluation is presently conducted of the performance improvements obtainable by these means in SSTs. It is noted that inlet pressure recovery increases and inlet drag reductions are obtainable via inlet control system integration; the use of propulsion system forces and moments to augment the flight control and aircraft stabilization system can also reduce empennage areas, weights, and drag. Special control modes may be instituted for community noise minimization and emergency procedures.

Burcham, Frank W., Jr.↗
NASA NTRS1991

Mach 6.5 air induction system design for the Beta II Two-Stage-to-Orbit booster vehicle

A preliminary, two-dimensional, mixed compression air induction system is designed for the Beta II Two-Stage-to-Orbit booster vehicle to minimize installation losses and efficiently deliver the required airflow. Design concepts, such as an external isentropic compression ramp and a bypass system, are developed and evaluated for performance benefits. The design is optimized by maximizing installed propulsion/vehicle system performance, and the resulting system design operating characteristics and performance are presented. The air induction system design has significantly lower transonic drag than similar designs, and only requires approximately 1/3 of the bleed extraction. In addition, the design efficiently provides the integrated system required airflow, while maintaining adequate levels of total pressure recovery. The excellent performance of this highly integrated air induction system is essential for the successful completion of the Beta II booster vehicle mission.

Midea, Anthony C.↗
NASA NTRS1991

Mach 6.5 air induction system design for the Beta 2 two-stage-to-orbit booster vehicle

A preliminary, two-dimensional, mixed compression air induction system is designed for the Beta II Two Stage to Orbit booster vehicle to minimize installation losses and efficiently deliver the required airflow. Design concepts, such as an external isentropic compression ramp and a bypass system were developed and evaluated for performance benefits. The design was optimized by maximizing installed propulsion/vehicle system performance. The resulting system design operating characteristics and performance are presented. The air induction system design has significantly lower transonic drag than similar designs and only requires about 1/3 of the bleed extraction. In addition, the design efficiently provides the integrated system required airflow, while maintaining adequate levels of total pressure recovery. The excellent performance of this highly integrated air induction system is essential for the successful completion of the Beta II booster vehicle mission.

Midea, Anthony C.↗
NASA NTRS1993

Computational parametric study of sidewall-compression scramjet inlet performance at Mach 10

A computational parametric study of three-dimensional, sidewall-compression scramjet inlets was performed to identify the effects of geometric parameters on inlet performance. The parameters were the leading-edge sweep angle, varied between 30 and 60 deg, and the leading-edge position of the cowl, located at the throat and at two forward positions. A laminar boundary layer with cold-wall (T(sub wall) = 300 K (540 R)) boundary conditions was imposed. The parametric study was performed for a Mach number of 10 and a unit free-stream Reynolds number of 7.06 x 10(exp 6) per meter (2.15 x 10(exp 6) per foot) at a geometric contraction ratio of 5. The performance of each configuration was evaluated in terms of the mass capture, throat Mach number, total pressure recovery, kinetic energy efficiency, and internal compression. One computation of an unswept configuration was included as a baseline to determine the effects of introducing leading-edge sweep on the flow-field parameters. The purpose of the computational parametric study was to perform a trade-off of the effects of various parameters on the global performance of the inlet. Although no single optimal configuration emerged, trade-offs among the stated performance parameters identified a leading-edge sweep angle of 45 deg as possessing the most attractive performance characteristics.

Holland, Scott D.↗
NASA NTRS1991

Turbulence and stall in plane diffusers - Computational study

The effect of free-stream turbulence level on stall conditions in plane diffusers is investigated with a finite-volume numerical method which utilizes a k-epsilon turbulence model or a differential Reynolds stress model for closure. Computed results show qualitative agreement with measured data. The ability to eliminate stall and thereby increase pressure recovery in diffusers at large included angles by utilizing high free-stream turbulence is correctly predicted.

Demuren, A. O.↗
NASA NTRS1955

Estimation of Inlet Lip Forces at Subsonic and Supersonic Speeds

The effects of lip thickness on inlet performance are estimated as functions of mass flow for subsonic and supersonic flight speeds. At subsonic speeds, pressure-recovery losses and additive drag are shown to decrease linearly with increasing lip frontal area if the maximum suction force is attained. At supersonic speeds, inlet drag increases linearly with inlet lip frontal area at full mass flow. For reduced mass flow, some reduction in additive drag is possible with lips of moderate thickness, but the magnitude of this reduction becomes negligible as flight speed increases.

Moeckel, W E↗
NASA NTRS1950

Free-flight Performance of 16-inch-diameter Supersonic Ram-jet Units II : Five Units Designed for Combustion-chamber-inlet Mach Number of 0.16 at Free-stream Mach Number of 1.60 (units B-1, B-2, B-3, B-4, and B-5) /c Wesley E. Messing and Scott H. S

Free-flight performance of five 16-inch-diameter ram-jet units was determined over range of free-stream Mach numbers of 0.50 to 1.86 and gas total-temperature ratios between 1.0 and 6.1 Time histories of performance data are presented for each unit. Correlations illustrate effect of free-stream Mach number and gas total-temperature ratio on diffuser total-pressure recovery, net-thrust coefficient, and external drag coefficient. One unit had smooth steady burning throughout the entire flight and encountered a maximum free-stream Mach number of 1.86 with a net acceleration of approximately 4.2 g's.

COMBUSTION - RAM-JET ENGINES↗
NASA NTRS1950

Investigation of Perforated Convergent-divergent Diffusers with Initial Boundary Layer

An experimental investigation was made at Mach number 1.90 of the performance of a series of perforated convergent-divergent supersonic diffusers operating with initial boundary layer, which was induced and controlled by lengths of cylindrical inlets affixed to the diffusers. Supercritical mass-flow and peak total-pressure recoveries were decreased slightly by use of the longest inlets (4 inlet diameters in length). Combinations of cylindrical inlets, perforated diffusers, and subsonic diffuser were evaluated as simulated wind tunnels having second throats. Comparisons with noncontracted configurations of similar scale indicated conservatively computed power reductions of 25 percent.

DIFFUSERS, SUPERSONIC↗
NASA NTRS1951

Investigation at Mach Number 1.88 of Half of a Conical-spike Diffuser Mounted as a Side Inlet with Boundary-layer Control / H. Fred Goelzer and Edgar M. Cortright, Jr

Experimental investigation was conducted at Mach number 1.88 to determine performance characteristics of half a 50 degree-conical-spike inlet mounted on a flat plate. Initial boundary layer was removed up-stream of inlet by a ram-type scoop of variable height. Initial boundary-layer thickness was also varied. With complete removal of initial boundary layer, total-pressure recovery of approximately 70 percent. Several alternative boundary-layer-removal systems were investigated which decreased the adverse effect of operating the ram scoop sub-critically.

BOUNDARY LAYER CONTROL-INTERNAL AERODYNAMICS↗
NASA NTRS1952

Performance of Air Inlets at Transonic and Low Supersonic Speeds

A general discussion of the air-inlet problem is presented. Recently obtained drag and pressure-recovery data for transonic-type nose, scoop, and wing-root inlets are summarized. Preliminary results concerning the performance of the sharp-edged supersonic-type inlets at transonic and subsonic speeds also are given.

Nichols, Mark R↗