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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

Effect of Propeller on Engine Cooling System Drag and Performance

The pressure recovery of incoming cooling air and the drag associated with engine cooling of a typical general aviation twin-engine aircraft was Investigated experimentally. The semispan model was mounted vertically in the 40 x 80-Foot Wind Tunnel at Ames Research Center. The propeller was driven by an electric motor to provide thrust with low vibration levels for the cold-now configuration. It was found that the propeller slip-stream reduces the frontal air spillage around the blunt nacelle shape. Consequently, this slip-stream effect promotes flow reattachment at the rear section of the engine nacelle and improves inlet pressure recovery. These effects are most pronounced at high angles of attack; that is, climb condition. For the cruise condition those improvements were more moderate.

Katz, Joseph↗
NASA NTRS2000

Two-Dimensional Bifurcated Inlet Variable Cowl Lip Test Completed in 10- by 10-Foot Supersonic Wind Tunnel

Researchers at the NASA Glenn Research Center at Lewis Field successfully tested a variable cowl lip inlet at simulated takeoff conditions in Glenn s 10- by 10-Foot Supersonic Wind Tunnel (10x10 SWT) as part of the High-Speed Research Program. The test was a follow-on to the Two-Dimensional Bifurcated (2DB) Inlet/Engine test. At the takeoff condition for a High-Speed Civil Transport aircraft, the inlet must provide adequate airflow to the engine with an acceptable distortion level and high-pressure recovery. The test was conducted to study the effectiveness of installing two rotating lips on the 2DB Inlet cowls to increase mass flow rate and eliminate or reduce boundary layer flow separation near the lips. Hardware was mounted vertically in the test section so that it extended through the tunnel ceiling and that the 2DB Inlet was exposed to the atmosphere above the test section. The tunnel was configured in the aerodynamic mode, and exhausters were used to pump down the tunnel to vacuum levels and to provide a maximum flow rate of approximately 58 lb/sec. The test determined the (1) maximum flow in the 2DB Inlet for each variable cowl lip, (2) distortion level and pressure recovery for each lip configuration, (3) boundary layer conditions near variable lips inside the 2DB Inlet, (4) effects of a wing structure adjacent to the 2DB Inlet, and (5) effects of different 2DB Inlet exit configurations. It also employed flow visualization to generate enough qualitative data on variable lips to optimize the variable lip concept. This test was a collaborative effort between the Boeing Company and Glenn. Extensive inhouse support at Glenn contributed significantly to the progress and accomplishment of this test.

Hoffman, T. R.↗
NASA NTRS2006

Optimal Design of Passive Flow Control for a Boundary-Layer-Ingesting Offset Inlet Using Design-of-Experiments

This research will investigate the use of Design-of-Experiments (DOE) in the development of an optimal passive flow control vane design for a boundary-layer-ingesting (BLI) offset inlet in transonic flow. This inlet flow control is designed to minimize the engine fan face distortion levels and first five Fourier harmonic half amplitudes while maximizing the inlet pressure recovery. Numerical simulations of the BLI inlet are computed using the Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, developed at NASA. These simulations are used to generate the numerical experiments for the DOE response surface model. In this investigation, two DOE optimizations were performed using a D-Optimal Response Surface model. The first DOE optimization was performed using four design factors which were vane height and angles-of-attack for two groups of vanes. One group of vanes was placed at the bottom of the inlet and a second group symmetrically on the sides. The DOE design was performed for a BLI inlet with a free-stream Mach number of 0.85 and a Reynolds number of 2 million, based on the length of the fan face diameter, matching an experimental wind tunnel BLI inlet test. The first DOE optimization required a fifth order model having 173 numerical simulation experiments and was able to reduce the DC60 baseline distortion from 64% down to 4.4%, while holding the pressure recovery constant. A second DOE optimization was performed holding the vanes heights at a constant value from the first DOE optimization with the two vane angles-of-attack as design factors. This DOE only required a second order model fit with 15 numerical simulation experiments and reduced DC60 to 3.5% with small decreases in the fourth and fifth harmonic amplitudes. The second optimal vane design was tested at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel in a BLI inlet experiment. The experimental results showed a 80% reduction of DPCPavg, the circumferential distortion level at the engine fan face.

Allan, Brian G.↗
NASA NTRS2006

Optimal Design of Passive Flow Control for a Boundary-Layer-Ingesting Offset Inlet Using Design-of-Experiments

This research will investigate the use of Design-of-Experiments (DOE) in the development of an optimal passive flow control vane design for a boundary-layer-ingesting (BLI) offset inlet in transonic flow. This inlet flow control is designed to minimize the engine fan-face distortion levels and first five Fourier harmonic half amplitudes while maximizing the inlet pressure recovery. Numerical simulations of the BLI inlet are computed using the Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, developed at NASA. These simulations are used to generate the numerical experiments for the DOE response surface model. In this investigation, two DOE optimizations were performed using a D-Optimal Response Surface model. The first DOE optimization was performed using four design factors which were vane height and angles-of-attack for two groups of vanes. One group of vanes was placed at the bottom of the inlet and a second group symmetrically on the sides. The DOE design was performed for a BLI inlet with a free-stream Mach number of 0.85 and a Reynolds number of 2 million, based on the length of the fan-face diameter, matching an experimental wind tunnel BLI inlet test. The first DOE optimization required a fifth order model having 173 numerical simulation experiments and was able to reduce the DC60 baseline distortion from 64% down to 4.4%, while holding the pressure recovery constant. A second DOE optimization was performed holding the vanes heights at a constant value from the first DOE optimization with the two vane angles-of-attack as design factors. This DOE only required a second order model fit with 15 numerical simulation experiments and reduced DC60 to 3.5% with small decreases in the fourth and fifth harmonic amplitudes. The second optimal vane design was tested at the NASA Langley 0.3- Meter Transonic Cryogenic Tunnel in a BLI inlet experiment. The experimental results showed a 80% reduction of DPCP(sub avg), the circumferential distortion level at the engine fan-face.

Allan, Brian G.↗
NASA NTRS1988

Wind tunnel pressurization and recovery system

The high density, low toxicity characteristics of refrigerant-12 (dichlorofluoromethane) make it an ideal gas for wind tunnel testing. Present limitations on R-12 emissions, set to slow the rate of ozone deterioration, pose a difficult problem in recovery and handling of large quantities of R-12. This preliminary design is a possible solution to the problem of R-12 handling in wind tunnel testing. The design incorporates cold temperature condensation with secondary purification of the R-12/air mixture by adsorption. Also discussed is the use of Freon-22 as a suitable refrigerant for the 12 foot wind tunnel.

Pejack, Edwin R.↗
NASA NTRS1952

A Low-speed Investigation of an Annular Transonic Air Inlet

Low-speed wind-tunnel tests were conducted as preliminary steps in the study of fuselage-air-inlet arrangements believed suitable for use at transonic speeds. The forward part of the model consisted of an NACA 1-85-050 cowling located at the base of the long protruding fuselage nose designed to maintain substream surface velocities everywhere ahead of the entrance and thereby to avoid or minimize adverse boundary-layer-shock interaction effects up to low supersonic speeds. Pressure-recovery and surface-pressure measurements are presented for the model with three fuselage nose shapes for ranges of angle of attack and inlet-velocity ratio useful for high-speed flight

AIR INLETS, SCOOPS↗
NASA NTRS1952

A Low-speed Investigation of a Fuselage-side Air Inlet for Use at Transonic Flight Speeds

Low-speed pressure-recovery and surface-pressure measurements for a twin fuselage-side air-inlet arrangement intended for use at transonic speeds are presented. The model was developed by adding a canopy and nose-wheel fairing to basic transonic annular inlet studied in TN 2685. The results indicate shape requirements of fuselage nose, canopy, and nose-wheel fairing for maintaining substream surface velocities everywhere ahead of entrance.

AIR INLETS, SIDE↗
NASA NTRS2015

Enhanced Performance of Streamline-Traced External-Compression Supersonic Inlets

A computational design study was conducted to enhance the aerodynamic performance of streamline-traced, external-compression inlets for Mach 1.6. The current study explored a new parent flowfield for the streamline tracing and several variations of inlet design factors, including the axial displacement and angle of the subsonic cowl lip, the vertical placement of the engine axis, and the use of porous bleed in the subsonic diffuser. The performance was enhanced over that of an earlier streamline-traced inlet such as to increase the total pressure recovery and reduce total pressure distortion

performance↗
NASA NTRS1975

Hypersonic research engine project. Phase 2: Some combustor test results of NASA aerothermodynamic integration model

Combustor test results of the NASA Aerothermodynamic Integration Model are presented of a ramjet engine developed for operation between Mach 3 and 8. Ground-based and flight experiments which provide the data required to advance the technology of hypersonic air-breathing propulsion systems as well as to evaluate facility and testing techniques are described. The engine was tested with synthetic air at Mach 5, 6, and 7. The hydrogen fuel was heated to 1500 R prior to injection to simulate a regeneratively cooled system. Combustor efficiencies up to 95 percent at Mach 6 were achieved. Combustor process in terms of effectiveness, pressure integral factor, total pressure recovery and Crocco's pressure-area relationship are presented and discussed. Interactions between inlet-combustor, combustor stages, combustor-nozzle, and the effects of altitude, combustor step, and struts are observed and analyzed.

Sun, Y. H.↗
NASA NTRS1990

Parametric evaluation of the aerodynamic performance of an annular combustor-diffuser system

This study experimentally determined the effects of dominant geometric and flow parameters on the aerodynamic performance of an annular combustor-diffuser system. For each parametric configuration, data were obtained at 13 different flow rates through the inner annulus, outer annulus, and combustor dome. The parametric effects investigated included the dump gap dimension, flow splits, prediffuser inlet velocity profile, and combustor channel height. During these measurements, the average prediffuser inlet Mach number was 0.305. The measured static pressure recovery coefficient and the total pressure loss coefficient in the prediffuser as well as the overall system are presented in this paper. The study found that the performance of the combustor-diffuser system improves with increasing dome flow rate or dump gap dimension.

Srinivasan, R.↗
NASA NTRSDate not supplied

A Separation Control CFD Validation Test Case: Baseline and Steady Suction - Part 1

Low speed flow separation over a wall-mounted hump, and its control using steady suction, were studied experimentally in order to generate a data set for a workshop aimed at validating CFD turbulence models. The baseline and controlled data sets comprised static and dynamic surface pressure measurements, flow field measurements using Particle Image Velocimetry (PIV) and wall shear stress obtained via oil-film interferometry. In addition to the specific test cases studied, surface pressures for a wide variety of conditions were reported for different Reynolds numbers and suction rates. Stereoscopic PIV and oil-film flow visualization indicated that the baseline separated flow field was mainly two- dimensional. With the application of control, some three-dimensionality was evident in the spanwise variation of pressure recovery, reattachment location and spanwise pressure fluctuations. Part 2 of this paper, under preparation for the AIAA Meeting in Reno 2005, considers separation control by means of zero-efflux oscillatory blowing.

David Greenblatt↗
NASA NTRS2004

A Separation Control CFD Validation Test Case: Baseline and Steady Suction - Part 1

Low speed flow separation over a wall-mounted hump, and its control using steady suction, were studied experimentally in order to generate a data set for a workshop aimed at validating CFD turbulence models. The baseline and controlled data sets comprised static and dynamic surface pressure measurements, flow field measurements using Particle Image Velocimetry (PIV) and wall shear stress obtained via oil-film interferometry. In addition to the specific test cases studied, surface pressures for a wide variety of conditions were reported for different Reynolds numbers and suction rates. Stereoscopic PIV and oil-film flow visualization indicated that the baseline separated flow field was mainly two-dimensional. With the application of control, some three-dimensionality was evident in the spanwise variation of pressure recovery, reattachment location and spanwise pressure fluctuations. Part 2 of this paper, under preparation for the AIAA Meeting in Reno 2005, considers separation control by means of zero-efflux oscillatory blowing.

Greenblatt, David↗
NASA NTRS1974

The 727 airplane side inlet low-speed performance confirmation model test for refanned JT8D engines

The results of a low-speed wind tunnel test of a 0.3 scale model 727 airplane side inlet for JT8D-100 engines are presented. The objectives of the test were to develop lines for a full-scale flightworthy inlet, to evaluate inlet total pressure recovery and steady-state total pressure distortion, and to obtain model-scale distortion data which can be used in the assessment of the compatibility of the inlet with the JT8D-100 series engines. A secondary objective was to obtain internal/external cowl static pressures for the determination of nacelle loads. Two basic inlet models were tested at static, forward speed, angle-of-attack (inflow angle), and cross-wind conditions. One model was with and one without an acoustic ring. Two modifications to the models were also tested, one with the ring closer to the inlet throat and one with a larger lip. Test measurements consisted of inlet surface static pressure, engine face total pressure, inlet airflow, tunnel total pressure, tunnel total temperature and tunnel velocity. Total pressure traverses were taken directly behind the ring and strut. No dynamic measurements were taken.

Schuehle, A. L.↗
NASA NTRS1977

Aerodynamic and directional acoustic performance of a scoop inlet

Aerodynamic and directional acoustic performances of a scoop inlet were studied. The scoop inlet is designed with a portion of the lower cowling extended forward to direct upward any noise that is propagating out the front of the engine toward the ground. The tests were conducted in an anechoic wind tunnel facility at free stream velocities of 0, 18, 41, and 61 m/sec and angles of attack from -10 deg to 120 deg. Inlet throat Mach number was varied from 0.30 to 0.75. Aerodynamically, at a free stream velocity of 41 m/sec, the design throat Mach number (0.63), and an angle of attack of 50 deg, the scoop inlet total pressure recovery was 0.989 and the total pressure distortion was 0.15. The angles of attack where flow separation occurred with the scoop inlet were higher than those for a conventional symmetric inlet. Acoustically, the scoop inlet provided a maximum noise reduction of 12 to 15 db below the inlet over the entire range of throat Mach number and angle of attack at a free-stream velocity of 41 m/sec.

Abbott, J. M.↗
NASA NTRS2006

Boundary-Layer-Ingesting Inlet Flow Control

This paper gives an overview of a research study conducted in support of the small-scale demonstration of an active flow control system for a boundary-layer-ingesting (BLI) inlet. The effectiveness of active flow control in reducing engine inlet circumferential distortion was assessed using a 2.5% scale model of a 35% boundary-layer-ingesting flush-mounted, offset, diffusing inlet. This experiment was conducted in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel at flight Mach numbers with a model inlet specifically designed for this type of testing. High mass flow actuators controlled the flow through distributed control jets providing the active flow control. A vortex generator point design configuration was also tested for comparison purposes and to provide a means to examine a hybrid vortex generator and control jets configuration. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion and pressure recovery were determined by 40 total pressure measurements on 8 rake arms each separated by 45 degrees and were located at the aerodynamic interface plane. The test matrix was limited to a maximum free-stream Mach number of 0.85 with scaled mass flows through the inlet for that condition. The data show that the flow control jets alone can reduce circumferential distortion (DPCPavg) from 0.055 to about 0.015 using about 2.5% of inlet mass flow. The vortex generators also reduced the circumferential distortion from 0.055 to 0.010 near the inlet mass flow design point. Lower inlet mass flow settings with the vortex generator configuration produced higher distortion levels that were reduced to acceptable levels using a hybrid vortex generator/control jets configuration that required less than 1% of the inlet mass flow.

Owens, Lewis R.↗
NASA NTRS2006

Boundary-Layer-Ingesting Inlet Flow Control

This paper gives an overview of a research study conducted in support of the small-scale demonstration of an active flow control system for a boundary-layer-ingesting (BLI) inlet. The effectiveness of active flow control in reducing engine inlet circumferential distortion was assessed using a 2.5% scale model of a 35% boundary-layer-ingesting flush-mounted, offset, diffusing inlet. This experiment was conducted in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel at flight Mach numbers with a model inlet specifically designed for this type of testing. High mass flow actuators controlled the flow through distributed control jets providing the active flow control. A vortex generator point design configuration was also tested for comparison purposes and to provide a means to examine a hybrid vortex generator and control jets configuration. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion and pressure recovery were determined by 40 total pressure measurements on 8 rake arms each separated by 45 degrees and were located at the aerodynamic interface plane. The test matrix was limited to a maximum free-stream Mach number of 0.85 with scaled mass flows through the inlet for that condition. The data show that the flow control jets alone can reduce circumferential distortion (DPCP(sub avg)) from 0.055 to about 0.015 using about 2.5% of inlet mass flow. The vortex generators also reduced the circumferential distortion from 0.055 to 0.010 near the inlet mass flow design point. Lower inlet mass flow settings with the vortex generator configuration produced higher distortion levels that were reduced to acceptable levels using a hybrid vortex generator/control jets configuration that required less than 1% of the inlet mass flow.

Owens, Lewis R.↗