Engineering topics

Boldman, D. R.

Publications and source records attributed to Boldman, D. R..

At least 19 records

NASA NTRS1994

Flow analysis for the nacelle of an advanced ducted propeller at high angle-of-attack and at cruise with boundary layer control

An axisymmetric panel code and a three dimensional Navier-Stokes code (used as an inviscid Euler code) were verified for low speed, high angle of attack flow conditions. A three dimensional Navier-Stokes code (used as an inviscid code), and an axisymmetric Navier-Stokes code (used as both viscous and inviscid code) were also assessed for high Mach number cruise conditions. The boundary layer calculations were made by using the results from the panel code or Euler calculation. The panel method can predict the internal surface pressure distributions very well if no shock exists. However, only Euler and Navier-Stokes calculations can provide a good prediction of the surface static pressure distribution including the pressure rise across the shock. Because of the high CPU time required for a three dimensional Navier-Stokes calculation, only the axisymmetric Navier-Stokes calculation was considered at cruise conditions. The use of suction and tangential blowing boundary layer control to eliminate the flow separation on the internal surface was demonstrated for low free stream Mach number and high angle of attack cases. The calculation also shows that transition from laminar flow to turbulent flow on the external cowl surface can be delayed by using suction boundary layer control at cruise flow conditions. The results were compared with experimental data where possible.

Hwang, D. P.↗

NASA NTRS1993

Effect of a rotating propeller on the separation angle of attack and distortion in ducted propeller inlets

The present study represents an extension of an earlier wind tunnel experiment performed with the P&W 17-in. Advanced Ducted Propeller (ADP) Simulator operating at Mach 0.2. In order to study the effects of a rotating propeller on the inlet flow, data were obtained in the UTRC 10- by 15-Foot Large Subsonic Wind Tunnel with the same hardware and instrumentation, but with the propeller removed. These new tests were performed over a range of flow rates which duplicated flow rates in the powered simulator program. The flow through the inlet was provided by a remotely located vacuum source. A comparison of the results of this flow-through study with the previous data from the powered simulator indicated that in the conventional inlet the propeller produced an increase in the separation angle of attack between 4.0 deg at a specific flow of 22.4 lb/sec-sq ft to 2.7 deg at a higher specific flow of 33.8 lb/sec-sq ft. A similar effect on separation angle of attack was obtained by using stationary blockage rather than a propeller.

Boldman, D. R.↗

NASA NTRS1993

Effect of a rotating propeller on the separation angle of attack

The present study represents an extension of an earlier wind tunnel experiment performed with the P&W 17-in. Advanced Ducted Propeller (ADP) Simulator operating at Mach 0.2. In order to study the effects of a rotating propeller on the inlet flow, data were obtained in the UTRC 10- by 15-Foot Large Subsonic Wind Tunnel with the same hardware and instrumentation, but with the propellar removed. These new tests were performed over a range of flow rates which duplicated flow rates in the powered simulator program. The flow through the inlet was provided by a remotely located vacuum source. A comparison of the results of this flow-through study with the previous data from the powered simulator indicated that in the conventional inlet the propeller produced an increase in the separation angle of attack between 4.0 deg at a specific flow of 22.4 lb/sec-sq ft to 2.7 deg at a higher specific flow of 33.8 lb/sec-sq ft. A similar effect on separation angle of attack was obtained by using stationary blockage rather than a propeller.

Boldman, D. R.↗

NASA NTRS1991

Evaluation of panel code predictions with experimental results of inlet performance for a 17-inch ducted prop/fan simulator operating at Mach 0.2

An axisymmetric panel code was used to evaluate a series of ducted propeller inlets. The inlets were tested in the Lewis 9 by 15 Foot Low Speed Wind Tunnel. Three basic inlets having ratios of shroud length to propeller diameter of 0.2, 0.4, and 0.5 were tested with the Pratt and Whitney ducted prop/fan simulator. A fourth hybrid inlet consisting of the shroud from the shortest basic inlet coupled with the spinner from the largest basic inlet was also tested. This later configuration represented the shortest overall inlet. The simulator duct diameter at the propeller face was 17.25 inches. The short and long spinners provided hub-to-tip retios of 0.44 at the propeller face. The four inlets were tested at a nominal free stream Mach number of 0.2 and at angles of attack from 0 degrees to 35 degrees. The panel code method incorporated a simple two-part separation model which yielded conservative estimates of inlet separation.

Boldman, D. R.↗

NASA NTRS1991

Evaluation of panel code predictions with experimental results of inlet performance for a 17-inch ducted prop/fab simulator operating at Mach 0.2

An axisymmetric panel code was used to evaluate a series of ducted propeller inlets. The inlets were tested in the Lewis 9 by 15 Foot Low Speed Wind Tunnel. Three basic inlets having ratios of shroud length to propeller diameter of 0.2, 0.4, and 0.5 were tested with the Pratt and Whitney ducted prop/fan simulator. A fourth hybrid inlet consisting of the shroud from the shortest basic inlet coupled with the spinner from the largest basic inlet was also tested. This later configuration represented the shortest overall inlet. The simulator duct diameter at the propeller face was 17.25 inches. The short and long spinners provided hub-to-tip ratios of 0.44 at the propeller face. The four inlets were tested at a nominal free stream Mach number of 0.2 and at angles of attack from 0 degrees to 35 degrees. The panel code method incorporated a simple two-part separation model which yielded conservative estimates of inlet separation.

Boldman, D. R.↗

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 NTRS1987

Wall shear stress measurement in blade end-wall corner region

The magnitude and the direction of wall shear stress and surface pressure in the blade end-wall corner region were investigated. The measurements were obtained on a specially designed Preston tube, the tip of which could be concentrically rotated about its axis of rotation at the measurement location. The magnitude of wall shear stress in the vicinity of the corner was observed to increase significantly (170 percent) compared to its far-upstream value; the increase was consistently higher on the blade surface compared to the value on the plate surface of the blade end-wall corner. On both surfaces in the blade end-wall corner, the variation of the wall shear stress direction was found to be more predominant in the vicinity of the blade leading-edge location. The trend of the measured wall shear stress direction showed good agreement with the limiting streamline directions obtained from the flow visualization studies.

Bhargava, R.↗

NASA NTRS1986

Three-dimensional fluid flow phenomena in the blade end wall corner region

Flow visualization, static and total pressure measurements, and mean velocity profile measurements with a single-sensor inclined hot wire probe, are used in a study of three-dimensional flow at a turbine blade end wall corner region for six critical axial stations along the blade chord. Three vortices are identified: (1) a horseshoe vortex near the leading edge; (2) a corner eddy between the horseshoe vortex and the corner; and (3) a vortex at the rear portion of the corner due to the corner eddy's secondary flow. Attention is given to the relative size and rate-of-spread of the vortices in the streamwise direction.

Hazarika, B. K.↗

NASA NTRS1986

Experimental evaluation of two turning vane designs for high-speed corner of 0.1-scale model of NASA Lewis Research Center's proposed altitude wind tunnel

Two turning vane designs were experimentally evaluated for corner 1 (downstream of the test section) of a 0.1-scale model of the NASA Lewis Research Center's proposed Altitude Wind Tunnel (AWT). Vane A was a controlled-diffusion airfoil shape; vane B was a circular-arc airfoil shape. The vane designs were tested over corner inlet Mach numbers from 0.16 to 0.465. Several modifications in vane setting angle and vane spacing were also evaluated for vane A. The overall performance obtained from total pressure rakes indicated that vane B had a slightly lower loss coefficient than vane A. At Mach 0.35 (the design Mach number without the engine exhaust removal scoop), the loss coefficients were 0.150 and 0.178 for vanes B and A, respectively. Resetting the vane A angle by -5 deg. (vane A10) to turn the flow toward the outside corner reduced the loss coefficient to 0.119. The best configuration (vane A10) was also tested with a simulated engine exhaust removal scoop. The loss coefficient for that configuration was 0.164 at Mach 0.41 (the approximate design Mach number with the scoop).

Moore, R. D.↗

NASA NTRS1986

Comparison of analytical and experimental performance of a wind-tunnel diffuser section

Wind tunnel diffuser performance is evaluated by comparing experimental data with analytical results predicted by an one-dimensional integration procedure with skin friction coefficient, a two-dimensional interactive boundary layer procedure for analyzing conical diffusers, and a two-dimensional, integral, compressible laminar and turbulent boundary layer code. Pressure, temperature, and velocity data for a 3.25 deg equivalent cone half-angle diffuser (37.3 in., 94.742 cm outlet diameter) was obtained from the one-tenth scale Altitude Wind Tunnel modeling program at the NASA Lewis Research Center. The comparison is performed at Mach numbers of 0.162 (Re = 3.097x19(6)), 0.326 (Re = 6.2737x19(6)), and 0.363 (Re = 7.0129x10(6)). The Reynolds numbers are all based on an inlet diffuser diameter of 32.4 in., 82.296 cm, and reasonable quantitative agreement was obtained between the experimental data and computational codes.

Shyne, R. J.↗

NASA NTRS1986

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfols were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a clasical usnteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Shaw, L. M.↗

NASA NTRS1985

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfroils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Shaw, L. M.↗

NASA NTRS1983

Turbulence spectra in the noise source regions of the flow around complex surfaces

The complex turbulent flow around three complex surfaces was measured in detail with a hot wire. The measured data include extensive spatial surveys of the mean velocity and turbulence intensity and measurements of the turbulence spectra and scale length at many locations. The publication of the turbulence data is completed by reporting a summary of the turbulence spectra that were measured within the noise source locations of the flow. The results suggest some useful simplifications in modeling the very complex turbulent flow around complex surfaces for aeroacoustic predictive models. The turbulence spectra also show that noise data from scale models of moderate size can be accurately scaled up to full size.

Olsen, W. A.↗

NASA NTRS1983

Experimental evaluation of shockless supercritical airfoils in cascade

Surface Mach number distributions, total pressure loss coefficients, and schlieren images of the flow are presented over a range of inlet Mach numbers and air angles. Several different trailing edge geometries were tested. At design conditions a leading edge separation bubble was observed resulting in higher losses than anticipated. The minimum losses were obtained at a negative incidence condition in which the flow was accelerating over most of the supercritical region. Relatively minor differences in losses were measured with the different trailing edge geometries studied.

Boldman, D. R.↗

NASA NTRS1983

Experimental evaluation of shockless supercritical airfoils in cascade

Boldman, D. R.↗

NASA NTRS1982

Three-dimensional shock structure in a transonic flutter cascade

Rapid double-pulse holography was employed to obtain detailed, two-dimensional images of the shock forming during simulated flutter in a transonic flowfield. The experiment comprised a linear cascade of airfoils externally oscillated in torsion and viewed tangentially at the shock surface. Three biconvex airfoils were subjected to harmonic pitching motion about the midchord axis at a frequency of 0.53 while immersed in a Mach 0.81 flow. Failure to produce observable shocks led to use of choked flow with a Mach number near one, of which 50 holograms were taken. The images revealed a narrow shock surface with a spanwise variation in the shock properties. The method is concluded to be useful for examining transonic flowfield shocks in the presence of airfoil flutter.

Boldman, D. R.↗

NASA NTRS1981

Stall flutter experiment in a transonic oscillating linear cascade

Two dimensional biconvex airfoils were oscillated at reduced frequencies up to 0.5 based on semi-chord and a free stream Mach number of 0.80 to simulate transonic stall flutter in rotors. Steady-state periodicity was confirmed through end-wall pressure measurements, exit flow traverses, and flow visualization. The initial flow visualization results from flutter tests indicated that the oscillating shock on the airfoils lagged the airfoil motion by as much as 80 deg. These initial data exhibited an appreciable amount of scatter; however, a linear fit of the results indicated that the greatest shock phase lag occurred at a positive interblade phase angle. Photographs of the steady-state and unsteady flow fields reveal some of the features of the lambda shock wave on the suction surface of the airfoils.

Boldman, D. R.↗

NASA NTRS1978

Wind tunnel tests of a blade subjected to midchord torsional oscillation at high subsonic stall flutter conditions

A mechanical drive system for oscillating blades in a wind tunnel at frequencies up to 767 hertz and amplitudes of + or - 1.2 deg is described. High-speed motion pictures of schlieren images of the flow over a double-circular arc blade oscillating in harmonic motion about the midchord revealed extensive shock patterns at a nominal free stream Mach number of 0.7, a mean angle of attack of 4 deg, and reduced frequency of about 0.7. A phase lag resulting from the slow response of the flow to the motion of the blade increased with increasing reduced frequency. This phase lag, based on the difference between the time the blade attained its maximum angle of attack and the time required for the normal shock to reach its extreme downstream position, was nominally 100 deg at the above conditions.

Boldman, D. R.↗