Engineering topics

Neumann, H. E.

Publications and source records attributed to Neumann, H. E..

At least 19 records

NASA NTRS1990

Qualitative investigation of cryogenic fluid injection into a supersonic flow field

The behavior of liquid nitrogen injected into a supersonic nitrogen flow field was investigated using an experimental apparatus in which a Mach 2.7 2D gas nitrogen tunnel is coupled with a high-pressure cryogenic source. Observations were monitored and recorded via a video camera and a motion picture camera. It was found that the penetration of a supersonic flow field by injection of liquid nitrogen is strongly dependent on the flow Mach number, the cryogen injection pressure (Pi/Pc), the injector configuration, and the cryogen temperature. For a 2D gaseous N2, Mach 2.7 tunnel, with cryogen injection Pi/Pc approaching 2, the injected fluid penetration for the 1/8-in. injection port approached one half of the tunnel width at 90-deg injection, and one fourth of the tunnel width at 20-deg injection.

Hendricks, R. C.↗

NASA NTRS1985

Performance and surge limits of a TF30-P-3 turbofan engine/axisymmetric mixed-compression inlet propulsion system at Mach 2.5

Steady-state performance and inlet-engine compatibility were investigated with a low-bleed inlet. The inlet had minimum internal contraction, consistent with high total pressure recovery and low cowl drag. The inlet-engine combination displayed good performance with only about 2% of inlet performance bleed. The inlet-engine combination had 5.58 deg angle-of-attack capability with 6% bleed.

Wasserbauer, J. F.↗

NASA NTRS1983

Dynamic distortion in a short s-shaped subsonic diffuser with flow separation

An experimental investigation of the time varying distortion at the diffuser exit of a subsonic HiMAT forebody and inlet was conducted at Mach 0.9 in the Lewis 8 by 6 foot Supersonic Wind Tunnel. A transitory separation was detected within the subsonic diffuser. Vortex generators were installed to eliminate the flow separation. Results from a study of the instantaneous pressure variations at the diffuser exit are presented. The time unsteady total pressures at the diffuser exit are computer interpolated and presented in the form of a movie showing the transitory separation. Limited data showing the instantaneous distortion levels are also presented. Previously announced in STAR as N83-26838

Stumpf, R.↗

NASA NTRS1983

An analytical and experimental comparison of the flow field of an advanced swept turboprop

An argon ion laser velocimeter with four beams was used to measure the detailed flow-field of an advanced eight blade propeller with 45% of tip sweep in an 8x6 foot supersonic wind tunnel. Data were obtained at a free stream Mach number of 0.8, the design advance ratio of 3.06 and a power coefficient of 1.8. Data are presented for inlet flow, exit flow, flow within the blades and flow slightly outside the blade tips. The data are compared to a lifting line theory. In general, the results of the comparison are considered favorable.

Neumann, H. E.↗

NASA NTRS1983

New test techniques and analytical procedures for understanding the behavior of advanced propellers

Analytical procedures and experimental techniques were developed to improve the capability to design advanced high speed propellers. Some results from the propeller lifting line and lifting surface aerodynamic analysis codes are compared with propeller force data, probe data and laser velocimeter data. In general, the code comparisons with data indicate good qualitative agreement. A rotating propeller force balance demonstrated good accuracy and reduced test time by 50 percent. Results from three propeller flow visualization techniques are shown which illustrate some of the physical phenomena occurring on these propellers.

Stefko, G. L.↗

NASA NTRS1983

Dynamic Distortion in a Short S-Shaped Subsonic Diffuser with Flow Separation

An experimental investigation of the time varying distortion at the diffuser exit of a subscale HiMAT forebody and inlet was conducted at Mach 0.9 in the Lewis 8 by 6 foot Supersonic Wind Tunnel. A transitory separation was detected within the subsonic diffuser. Vortex generators were installed to eliminate the flow separation. Results from a study of the instantaneous pressure variations at the diffuser exit are presented. The time unsteady total pressures at the diffuser exit are computer interpolated and presented in the form of a movie showing the transitory separation. Limited data showing the instantaneous distortion levels is also presented.

Stumpf, R.↗

NASA NTRS1983

An analytical and experimental comparison of the flow field of an advanced swept turboprop

(Previously announced in STAR as N83-14061)

Neumann, H. E.↗

NASA NTRS1982

LV measurements with an advanced turboprop

Nonintrusive measurements of velocity about a spinner-propeller-nacelle configuration were made at a Mach number of 0.8. A laser velocimeter (LV) specifically developed for these measurements was used to determine the flow field of the advanced swept SR-3 propeller. The data will be used to study the flow and to verify computer prediction codes. The usefulness of the LV data in detecting flow anomalies and to substantiate the data quality was demonstrated. Some typical results are given. Mach number profiles at the entrance of the propeller are compared with theoretical predictions. The LV data is in excellent agreement with the axisymmetric, compressible, inviscid theory (without blades) ahead of the propeller except near the hub. The data indicate blade blockage near the spinner. Blade to blade variations in axial velocity for four radial positions at the propeller exist are also given. The large apparent wake near the hub is associated with the hub choking. The blade to blade variation of axial velocity ahead of a shock within the blade passage is given.

Neumann, H. E.↗

NASA NTRS1981

Laser-velocimeter flow-field measurements of an advanced turboprop

Non-intrusive measurements of velocity about a spinner-propeller-nacelle configuration at a Mach number of 0.8 have been performed. A laser velocimeter, specifically developed for these measurements in the NASA Lewis 8- by 6-foot Supersonic Wind Tunnel, was used to measure the flow-field of the advanced swept SR-3 propeller. The laser velocimeter uses an argon ion laser and a 2-color optics system to allow simultaneous measurements of 2-components of velocity. The axisymmetric nature of the propeller-nacelle flow-field permits two separate 2-dimensional measurements to be combined into 3-dimensional velocity data. Presented are data ahead of and behind the propeller blades and also a limited set in between the blades. Aspects of the observed flow-field such as the tip vortex are discussed.

Serafini, J. S.↗

NASA NTRS1981

Laser-velocimeter flow-field measurements of an advanced turboprop

Non-intrusive measurements of velocity about a spinner-propeller-nacelle configuration at a Mach number of 0.8 were performed. A laser velocimeter, specifically developed for these measurements in the NASA Lewis 8-foot by 6-foot Supersonic Wind Tunnel, was used to measure the flow-field of the advanced swept SR-3 turboprop. The laser velocimeter uses an argon ion laser and a 2-color optics system to allow simultaneous measurements of 2-components of velocity. The axisymmetric nature of the propeller-nacelle flow-field permits two separate 2 dimensonal measurements to be combined into 3 dimensional velocity data. Presented are data ahead of and behind the prop blades and also a limited set in between the blades. Aspects of the observed flow-field such as the tip vortex are discussed.

Serafini, J. S.↗

NASA NTRS1980

An analytical and experimental study of a short s-shaped subsonic diffuser of a supersonic inlet

A subscale HiMAT forebody and inlet was investigated over a range of Mach numbers to 1.4. The inlet exhibited a transitory separation within the diffuser but steady state data indicated reattachment at the diffuser exit. A finite difference procedure for turbulent compressible flow in axisymmetric ducts was used to successfully model the HiMAT duct flow. The analysis technique was further used to estimate the initiation of separation and delineate the steady and unsteady flow regimes in similar S-shaped ducts.

Neumann, H. E.↗

NASA NTRS1980

An analytical and experimental study of a short S-shaped subsonic diffuser of a supersonic inlet

An experimental investigation of a subscale HiMat forebody and inlet was conducted over a range of Mach numbers to 1.4. The inlet exhibited a transitory separation within the diffuser but steady state data indicated reattachment at the diffuser exit. A finite difference procedure for turbulent compressible flow in axisymmetric ducts was used to successfully model the HiMAT duct flow. The analysis technique was further used to estimate the initiation of separation and delineate the steady and unsteady flow regimes in similar S-shaped ducts.

Neumann, H. E.↗

NASA NTRS1979

Reynolds number, scale and frequency content effects on F-15 inlet instantaneous distortion

An inlet instantaneous distortion study program sponsored by NASA was recently completed using an F-15 fighter aircraft. Peak distortion data from subscale inlet model wind tunnel tests are shown to be representative of full-scale flight test peak distortion. The effects on peak distortion are investigated for engine presence, Reynolds number, scale and frequency content. Data are presented which show that: (1) the effect of engine presence on total pressure recovery, peak distortion, and turbulence is small but favorable, (2) increasing the Reynolds number increases total pressure recovery, decreases peak distortion, and decreases turbulence, and (3) increasing the filter cutoff frequency increases the calculated values of both peak distortion and turbulence.

Stevens, C. H.↗

NASA NTRS1976

A method of distortion pattern synthesis for high response data screening

An empirically derived method of estimating maximum levels of instantaneous inlet total pressure distortion from steady state measurements is presented. The estimation procedure uses steady state total pressure and root mean square turbulent pressure data measurements to synthesize instantaneous distortion patterns. The synthesis method is tailored to intensify distortion patterns in a manner consistent with the characteristics of the particular index to be used in evaluating inlet/engine compatability. The maximum expected value of distortion, a statistically determined function of the length of time of inlet operation, is used to constrain the amount of pattern intensification. Comparison with an independent set of data is presented for verification of the synthesis method.

King, R. W.↗

NASA NTRS1976

Boundary layer bleed system study for a full-scale, mixed-compression inlet with 45 percent internal contraction

The results of an experimental bleed development study for a full-scale, Mach 2.5, axisymmetric, mixed-compression inlet were presented. The inlet was designed to satisfy the airflow requirements of the TF30-P-3 turbofan engine. Capabilities for porous bleed on the cowl surface and ram-scoop/flush-slot bleed on the centerbody were provided. A configuration with no bleed on the cowl achieved a minimum stable, diffuser exit, total pressure recovery of 0.894 with a centerbody-bleed mass flow ratio of 0.02. Configurations with cowl bleed had minimum stable recoveries as high as 0.900 but suffered range decrement penalties from the increased bleed mass flow removal. Limited inlet stability and unstart angle-of-attack data are presented.

Shaw, R. J.↗

NASA NTRS1975

Performance of vortex generators in a Mach 2.5 low-bleed full scale 45-percent-internal-contraction axisymmetric inlet

Steady-state and dynamic flow characteristics associated with two sets of vortex generators having different mixing criteria were determined. The inlet performance with and without these vortex generators is presented. The vortex generators were successful in eliminating separation, increasing area-weighted total pressure recovery, and decreasing distortion. Transmission times obtained from cross-correlations of the wall static pressures and the diffuser exit total pressure showed no effect of the upstream flow characteristics on the diffuser exit pressures when generators were used. Without generators, separation occurred and the upstream pressure characteristics had immediate effects on the diffuser exit pressure characteristics.

Neumann, H. E.↗

NASA NTRS1975

Design of a very-low-bleed Mach 2.5 mixed-compression inlet with 45 percent internal contraction

A full-scale, mixed-compression inlet was designed for operation with the TF30-P-3 turbofan engine and tested at Mach numbers of 2.5 and 2.0. The two-cone axisymmetric inlet had minimum internal contraction consistent with high total pressure recovery and low cowl drag. At Mach 2.5, inlet recovery was 0.906 with only 0.021 centerbody bleed mass-flow ratio and no cowl bleed. Increased centerbody bleed gave a maximum inlet unstart angle of attack of 6.85 deg. At Mach 2.0, inlet recovery was 0.94 with only 0.014 centerbody bleed mass-flow ratio and no cowl bleed. Inlet performance and angle-of-attack tolerance is presented for operation at Mach numbers of 2.5 and 2.0.

Wasserbauer, J. F.↗

NASA NTRS1974

Distortion in a full-scale bicone inlet with internal focused compression and 45 percent internal contraction

The distortion characteristics were investigated at the subsonic diffuser exit of a full-scale, Mach 2.5, axisymmetric, mixed compression inlet. Performance and steady-state distortion characteristics were obtained at zero and maximum angle of attack and during an inlet unstart-restart sequence. For the configuration with no cowl bleed, steady-state distortion P(max)P(min)P(bar) ranged from 0.10 for critical inlet operation at 0 deg angle-of-attack to 0.306 for supercritical inlet operation at 6.84 deg angle-of-attack. Vortex generators provided a 50 percent reduction in steady-state distortion for critical operation. Bleed has a smaller effect on steady-stated distortion.

Wasserbauer, J. F.↗