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Modified single-shaft turbopump incorporates inducer and main pump, each separately driven at different speeds through coaxial-shaft arrangement. Inducer operates at low speed for low net positive suction pressure, main pump operates at high speed to generate high pressure. This arrangement requires no external control for the inducer.
Influence of suction on shock wave turbulent boundary layer interactions for two dimensional and axially symmetric flows
Leading edge suction analogy for predicting low speed lift and drag-due-to-lift characteristics of sharp edge delta and related wing planforms
Conical shock wave-turbulent boundary layer interaction data obtained for adiabatic wall conditions at supersonic free stream Mach numbers, including suction effects
Design, characteristics, and performance tests of short two dimensional curved wall diffusers using suction slots
Performance tests and efficiency measurement of hub-slit-suction stator in single stage compressor
Leading edge suction analogy for predicting low speed lift and drag-due-to-lift characteristics of sharp edge delta and related wing planforms
The leading-edge-suction analogy of Polhamus has been used to develop the longitudinal load distribution of the vortex lift for delta wings. This distribution is shown to be similar in shape to that of the potential-flow longitudinal loading for delta wings having aspect ratios of 2 or less. The totals of the two theoretical distributions for delta wings with an aspect ratio near 1 are in good agreement with the experimentally determined loadings over the angle-of-attack range from 0 to 30 deg. The corresponding predicted pitching moments show slightly more stability than those measured, because of loss of lift near the wing tips.
The modelling of turbulent transport processes with an eddy viscosity representation of the inner, wall region for cases with suction or injection is considered in detail. It is shown how the van Driest solid wall model can be logically extended for such cases. Further, a new inductive model is developed. Finite difference calculations using the new inductive model are shown to be in good agreement with experimental data over a wide range of conditions. Lastly, a new method for calculating the wall shear within a finite difference solution is developed. This procedure greatly relaxes the requirement for fine grid spacing in the vertical direction.
The effect on aerodynamic performance of a single row of spanwise-spaced coolant holes located at four positions chordwise along the suction surface was investigated. In addition, multiple-row data were obtained. The data are presented in terms of primary efficiency, coolant-to-primary-flow percentage, and coolant pressure ratio. Primary efficiencies of multirow blade configurations compare satisfactorily with that calculated from the single-row efficiency increments. At any given coolant flow percentage, the efficiency was about the same for all row locations. For a given coolant pressure, the efficiency varied by as much as 1 percent, depending on the row location.
An asymmetric annular diffuser equipped with wall bleed (suction) capability was tested for controllability of exit velocity profile. The diffuser area ratio was 3.2, and the length to inlet height ratio was 1.6. Results show that the diffuser radial exit velocity profile could be controlled from a hub peaked to a tip peaked form by selective use of bleed on the outer wall or on both diffuser walls. Based on these results, application of the diffuser bleed technique to gas turbine combustors may be possible. Diffuser bleed could be used to tailor the airflow distribution for optimizing combustor performance at a variety of operating conditions.
The performance of a short highly asymmetric annular diffuser equipped with wall bleed (suction) capability was evaluated at nominal inlet Mach numbers of 0.188, 0.264, and 0.324 with the inlet pressure and temperature at near ambient values. The diffuser had an area ratio of 2.75 and a length- to inlet-height ratio of 1.6. Results show that the radial profiles of diffuser exit velocity could be controlled from a severely hub peaked to a slightly tip biased form by selective use of bleed. At the same time, other performance parameters were also improved. These results indicate the possible application of the diffuser bleed technique to control flow profiles to gas turbine combustors.
Film-cooling-air ejection from the suction surface of a turbine vane was investigated. This investigation was conducted in a four-vane cascade on a J75 size turbine vane which had a row of holes near the leading edge. The experimental data are presented. It was found that a small amount of film-cooling air has a detrimental effect on the downstream vane wall cooling effectiveness. It was also shown that the presence of the film-cooling holes, without blowing, also cause an increase in vane wall temperatures. These results came from an increase in the gas-side heat transfer coefficient that was apparently caused by a laminar or transitional boundary layer becoming transitional or turbulent.
Results are presented of a coordinated experimental and theoretical study of a sound shield concept which aims to provide a means of noise reduction in the test section of supersonic wind tunnels at high Reynolds numbers. The model used consists of a planar array of circular rods aligned with the flow, with adjustable gaps between them for boundary layer removal by suction, i.e., laminar flow control. One of the basic requirements of the present sound shield concept is to achieve sonic cross flow through the gaps in order to prevent lee-side flow disturbances from penetrating back into the shielded region. Tests were conducted at Mach 6 over a local unit Reynolds number range from about 1.2 x 10 to the 6th power to 13.5 x 10 to the 6th power per foot. Measurements of heat transfer, static pressure, and sound levels were made to establish the transition characteristics of the boundary layer on the rod array and the sound shielding effectiveness.
Eleven semi-span wing models were tested in the 1/8-scale model of the Langley V/STOL tunnel to qualitatively study vortex bursting. Flow visualization was achieved by using helium filled soap bubbles introduced upstream of the model. The angle of attack range was from 0 deg to 45 deg. The results show that the vortex is unstable, that is, the bursting point location is not fixed at a given angle of attack but moves within certain bounds. Upstream of the trailing edge, the bursting point location has a range of two inches; downstream, the range is about six inches. Anhedral and dihedral appear to have an insignificant effect on the vortex and its bursting point location. Altering the section suction distribution by improving the triangularity generally increases the angle of attack at which vortex bursting occurs at the trailing edge.
Most fuselage geometries cover a portion of the wing leading edge near the plane of symmetry, and it seems reasonable to expect that a large fraction of the leading edge suction which would be developed by the covered wing at high angles of attack is not developed on the fuselage. This is one of the reasons that the Oswald span efficiency factor for the wing body combination fails to approach the value predicted by lifting line theory for the isolated wing. Some traditional and recent literature on wing-body interference is discussed and high Reynolds number data on wing-body-nacelle drag are reviewed. An exposed central leading edge geometry has been developed for a sailplane configuration. Low Reynolds number tests have not validated the design concept.
Tabulated data from a series of experimental studies of the interaction of a shock wave with a turbulent boundary layer in axisymmetric flow configurations is presented. The studies were conducted at the walls of circular wind tunnels and on the cylindrical centerbody of an annular flow channel. Detailed pitot pressure profiles and wall static pressure profiles upstream of, within and downstream of the interaction region are given. Results are presented for flows at nominal freestream Mach Numbers of 2, 3 and 4. For studies at the tunnel sidewalls, the shock waves were produced by conical shock generators mounted on the centerline of the wind tunnel at zero angle of attack. The annular ring generator was used to produce the shock wave at the centerbody of the annular flow channel. The effects of boundary layer bleed were examined in the investigation. Both bleed rate and bleed location were studied. Most of the bleed studies were conducted with bleed holes drilled normal to the wall surface but the effects of slot suction were also examined. A summary of the principal results and conclusions is given.
In this paper we present a numerical method for solving the compressible laminar boundary-layer equations with suction on swept and tapered wings. The method employs an efficient two-point finite-difference method to solve the governing equations, and a very convenient similarity transformation which removes the wall normal velocity as a boundary condition and places it into the governing equations as a parameter. In this way the awkward nonlinear boundary condition which couples all the variables is avoided. To test and demonstrate the method, we present a sample calculation for a typical laminar-flow-control (LFC) wing.