Garth V. Hobson

Prediction of Cascade Performance Using an Incompressible Navier–Stokes Technique

A fully elliptic, control volume solution of the two-dimensional incompressible Navier-Stokes equations for the prediction of cascade performance over a wide incidence range is presented. The numerical technique is based on a new pressure substitution method. A Poisson equation is derived from the pressure weighted substitution of the full momentum equations into the continuity equation. The analysis of a double circular arc compressor cascade is presented, and the results are compared with the available experimental data at various incidence angles. Good agreement is obtained for the blade pressure distribution, boundary layer and wake profiles, skin friction coefficient, losses and outlet angles. Turbulence effects are simulated by the Low-Reynolds-Number version of the k-epsilon turbulence model.

Effects of Reynolds Number and Surface Roughness Magnitude and Location on Compressor Cascade Performance

An experimental investigation has been conducted to characterize the influence of Reynolds number and surface roughness magnitude and location on compressor cascade performance. Flow field surveys have been conducted in a low-speed, linear compressor cascade. Pressure, velocity, and loss have been measured via a five-hole probe, pitot probe, and pressure taps on the blades. Four different roughness magnitudes, Ra values of 0.38 lm (polished), 1.70 lm (baseline), 2.03 lm (rough 1), and 2.89 lm (rough 2), have been tested. Furthermore, various roughness locations have been examined. In addition to the as manufactured (baseline) and entirely rough blade cases, blades with roughness covering the leading edge, pressure side, and 5%, 20%, 35%, 50%, and 100% of suction side from the leading edge have been studied. All of the tests have been carried out for Reynolds numbers ranging from 300,000 to 640,000. For Reynolds numbers under 500,000, the tested roughnesses do not significantly degrade compressor blade loading or loss. However, loss and blade loading become sensitive to roughness at Reynolds numbers above 550,000. Cascade performance is more sensitive to roughness on the suction side than pressure side. Furthermore, roughness on the aft 2=3 of suction side surface has a greater influence on loss. For a given roughness location, there exists a Reynolds number at which loss begins to significantly increase. Finally, increasing the roughness area on the suction surface from the leading edge reduces the Reynolds number at which the loss begins to increase. [DOI: 10.1115/1.4003821]

Effect of Reynolds Number on Separation Bubbles on Compressor Blades in Cascade

A detailed experimental investigation of second-generation, controlled-diffusion, compressor stator blades at an off-design inlet-e ow angle was performed in a low-speed-cascade wind tunnel primarily using laser Doppler velocimetry (LDV). The object was to characterize the e owe eld in the Reynolds number range of 2 :1‐6:4 £ 10 5 and to obtain LDV measurements of the suction surface boundary-layer separation that occurred near midchord. Surfacee owvisualizationshowedthatatthelowReynoldsnumberthemidchordseparationbubblestartedlaminar and reattachedturbulentwithin 20%chordonthesuctionsideoftheblade.Theextentofthebubblecomparedvery well with the measured blade surface pressure distribution, which showed a classical plateau and then diffusion in the turbulent region. LDV measurements of the e ow reversal in the bubble were performed. At the intermediate Reynolds number, the boundary layer was transitional upstream of the separation bubble that had decreased signie cantly in size (down to 10% chord ). At thehighestReynoldsnumber, the e owwas turbulentfrom close to the leading edge, and three-dimensional e ow reversal as a result of endwall effects appeared at approximately 80% chord. These data, particularly the low Reynolds number data, are an excellent test case for either large-eddy simulation or direct numerical simulation of cascade e owe elds.

Turbulence Amplification with Incidence at the Leading Edge of a Compressor Cascade

Detailed measurements, with a two-component laser-Doppler velocimeter and a thermal anemometer were made near the suction surface leading edge of controlled-diffusion airfoils in cascade. The Reynolds number was near 700,000, Mach number equal to 0.25, and freestream turbulence was at 1.5% ahead of the cascade.

Effect of Surface Roughness Location and Reynolds Number on Compressor Cascade Performance

An experimental work has been conducted in a linear compressor cascade to find out the effect of surface roughness and Reynolds number. Surveys were conducted with different roughness size and Reynolds number. The k s /c value of each roughness is 0.0006, 0.0090, 0.00150, 0.00213, and 0.00425. The range of Reynolds number is 300,000~600,000 and conducted with roughened blade, which roughness Ra is 2.89 microns. Flow pressure, velocity, and angle have been found out via 5 hole probe. Pressure loss and deviation increased with increasing roughness. In the low Reynolds number under 500,000, tested roughness does not affect to the performance of compressor cascade. However, roughness is very sensitive to pressure loss in high Reynolds number over 550,000.

Pre-Stall Instability Distribution Over a Transonic Compressor Rotor

The present study was part of the compressor research program sponsored by the Propulsion and Power Department of the Naval Air Warfare Centre, Patuxent River, MD with Ravi Ravindranath as the technical monitor.

Three-Component LDV Measurements in the Wake of a Compressor Cascade With Flow Separation

Measurements were taken of the vortex system and turbulent flow that resulted from the interaction between a stator blade and the approaching endwall boundary layer in a linear cascade of compressor blades. Data were taken at a Reynolds number based on blade chord of 640,000. Five-hole pressure measurements were conducted upstream and downstream of the blade row. The approaching boundary layer was also characterized with the laser-Doppler velocimeter. Downstream three-component laser-Doppler velocimetry surveys were conducted at three streamwise stations to map the location and velocity characteristics of the wake and vortex system. Results clearly showed the extent of the vortex emanating from the separation of the boundary layer on the suction side of the blade. Finally, all components of mean flow velocity and turbulence are documented for the last survey station. These data will form a challenging test case for numerical code validation.

Impact of Non-Uniform Leading Edge Coatings on the Aerodynamic Performance of Compressor Airfoils

Abstract : A computational fluid dynamic (CFD) investigation is presented that provides predictions of the aerodynamic impact of uniform and non-uniform coatings applied to the leading edge of a compressor airfoil in a cascade. Using a NACA 65(12)10 airfoil, coating profiles of varying leading edge non-uniformity were added. This non-uniformity is typical of that expected due to fluid being drawn away from the leading edge during the coating process. The CFD code, RVCQ3D, is a steady, quasi-three-dimensional Reynolds Averaged Navier-Stokes (RANS) solver. A k-omega turbulence model was used for the Reynolds Stress closure. The code predicted that these changes in leading edge shape can lead to alternating pressure gradients in the first few percent of chord that create small separation bubbles and possibly early transition to turbulence. The change in total pressure loss and trailing edge deviation are presented as a function of the coating non-uniformity parameter. Results are presented for six leading edge profiles over a range of incidences and inlet Mach numbers from 0.6 to 0.8. Reynolds number was 600,000 and free-stream turbulence was 6%. A two-dimensional map is provided that shows the allowable degree of coating non-uniformity as a function of incidence and inlet Mach number.

Laser-Doppler-Velocimetry Measurements in a Cascade of Compressor Blades at Stall

Compressor stall was simulated in the Low-Speed Cascade Wind Tunnel at the Turbopropulsion Laboratory of the Naval Postgraduate School. The test blades were of controlled-diffusion design with a solidity of 1.67, and stalling occurred at 10 deg of incidence above the design inlet air angle. All measurements were taken at a flow Reynolds number, based on chord length, of 700,000. Laser-sheet flow visualization techniques showed that the stalling process was unsteady and occurred over the whole cascade. Detailed laser-Doppler-velocimetry measurements over the suction side of the blades showed regions of continuous and intermittent reverse flow. The measurements of the continuous reverse flow region at the leading edge were the first data of their kind in the leading edge separation bubble. The regions of intermittent reverse flow, measured with laser-Doppler velocimeter, corresponded to the flow visualization studies. Blade surface pressure measurements showed a decrease in normal force on the blade, as would be expected at stall. Data are presented in a form that characterizes the unsteady positive and negative velocities about their mean, for both the continuous reverse flow regions and the intermittent reverse flow regions.

Prediction of Cascade Performance Using an Incompressible Navier-Stokes Technique

A fully elliptic, control volume solution of the two-dimensional incompressible Navier-Stokes equations for the prediction of cascade performance over a wide incidence range is presented in this paper. The numerical technique is based on a new pressure substitution method. A Poisson equation is derived from the pressure weighted substitution of the full momentum equations into the continuity equation. The analysis of a double circular arc compressor cascade is presented, and the results are compared with the available experimental data at various incidence angles. Good agreement is obtained for the blade pressure distribution, boundary layer and wake profiles, skin friction coefficient, losses and outlet angles. Turbulence effects are simulated by the Low-Reynolds-Number version of the k-e turbulence model.Copyright