Aeronautics and astronautics – Aircraft sustentation – Sustaining airfoils
Reexamination Certificate
2000-11-24
2002-12-03
Poon, Peter M. (Department: 3643)
Aeronautics and astronautics
Aircraft sustentation
Sustaining airfoils
C244S207000, C244S208000
Reexamination Certificate
active
06488238
ABSTRACT:
The present invention relates in general to aerodynamic and thermal wall boundary layer control of aerodynamic airfoils, in particular to application of such to aircraft, turbo engines and automotive components, especially but not exclusively, to isolated airfoils or fuselage parts, wings, flaps and blades, casing, end walls and liners of turbo engines in order to increase the efficiency and to improve off design performance.
As well known, a boundary layer develops on a wall of a body as soon as it interacts with a flow stream. The aerodynamic performance (lift or loading and drag) and the thermal performance (heat exchange efficiency and working temperature of the component) on nominal and off design conditions depend on the properties and the structure of the boundary layer.
The purpose of boundary layer control, also known as BLC (Boundary Layer Control), is to affect the flow by influencing the structure of the boundary layer, in order to increase the efficiency, the loading and the stage pressure ratio of turbo engines and off design performance of isolated airfoils and bodies.
The main advantages of boundary layer control are:
to delay transition from a laminar to a turbulent boundary layer and thus reduce skin friction and heat transfer;
to prevent or delay boundary layer separation and thereby increase the allowable blade or airfoil loading and range of angles of attack;
to cancel or attenuate flow disturbances in transonic flows.
Such effects can be achieved either by suction of the boundary layer in the regions of interest, or by injection of a working fluid which in turn can be either the same as the main flow or a different one (binary boundary layer).
Boundary layer control by means of a suction method appears, in general, to be the more efficient, but the injection method would be necessary for its use in conjunction with air cooling of turbine blades.
Concerning aircraft or automotive applications, the suction area may be applied either at wing leading edges or at the forward edge of the trailing edge flaps.
Concerning turbo engines, as in the case of isolated airfoils, it is possible to increase the blade loading considerably by the use of boundary layer control. This increase is achieved by prevention or delay of flow separation and the resultant stalling of the blades, and the greatest gain can be obtained if the blades are specially designed for the use of the boundary layer control so as to maintain high loading over most of the blade section. By use of boundary layer control a higher blade loading can be obtained for a given inlet Mach number without exceeding a given Mach number on the blade. In fact the blade can be designed to have a uniformly high Mach number over a larger portion of the upper surface of the blade without flow separation. For the same reason, higher inlet Mach numbers can be used with a given loading without exceeding a given Mach number on the blade. The gain resulting from maintaining a high velocity over a larger portion of the upper surface may be materially reduced, however, by increased velocity over the lower surface as a result of the practical requirement of thicker blades for boundary layer control.
With a conventional blading, the maximum pressure ratio per stage in a multistage compressor is obtained by increasing the axial velocity component and maintaining an essentially symmetrical velocity diagram throughout in order to assure the maximum allowable Mach number on all blade elements. This increase in the axial velocity component can only be obtained by using a sufficiently large taper for the annular passage to more than compensate for the reduction in axial velocity due to the increase in density. The use of taper large enough to maintain constant relative Mach numbers, however, leads to very small passages in the later stages of high pressure-ratio compressors and to high exit velocities. The resultant narrow annular passage tends to produce low efficiencies in the later stages and the high exit velocities either produce large exit losses or require long diffusers. For these reasons, most commercial compressors use much less taper than required to produce constant Mach number and consequently obtain relatively low Mach numbers and pressure ratios in the later stages.
The drop-off in pressure ratio in the later stages due to this decrease in relative Mach number in the wider passage could be prevented, however, if the blade loading could be increased in these stages. Because of the lower Mach numbers, it should be possible to use considerably higher blade loading without obtaining eccessive local Mach numbers on the blades. With conventional blading, however the blade loading is limited by the early stalling of the blades.
The situation can be somewhat improved by the use of blades of high camber, but the gain is limited and the useful range of angles of attack may be reduced. Some further slight increase in pressure ratio can he obtained by the use of solidifies higher than conventional limits of about 1.2, but the gain is generally obtained with some drop in efficiency.
A definitely greater increase in loading and stage pressure ratio should be possible through the use of boundary-layer control on the rotor, stator blades and hub and tip endwalls. The results from isolated airfoils indicate that there should be no difficulty in doubling the loading obtainable without boundary-layer control with a corresponding increase in pressure ratio. The effect of boundary-layer control on the stage efficiency is less easy to evaluate than its effect on stage pressure ratio.
The profile drag-lift ratio should be decreased because of the large increase in lift coefficient possible without boundary-layer separation and because of the decreased profile drag resulting from reduced boundary-layer thickness behind the control slot or porous structure. In addition, some improvement in efficiency might be expected from the fact that, for a given drag-lift ratio, the velocity diagram theoretically most favorable to high profile efficiency (symmetrical diagram with axial velocity equal to one-half rotor-blade velocity), can be approached more closely throughout a multistage compressor when a high stage pressure ratio is obtained by use of boundary-layer control on both rotor and stator blades. The effect is generally obtained with conventional blades by varying the axial velocity to give a constant Mach number entering all blade rows.
The boundary-layer control when applied at critical points of the blades and the casing, allows for positively action on the secondary-flow losses leading to a consistent reduction of the overall losses of the stages.
In addition to the effect on internal aerodynamic efficiency, the power required for supplying the boundary-layer-control air must be considered in evaluating the over-all efficiency of the machine. Because of this pumping loss, it appears desirable to limit the boundary-layer control to the later stages where the largest gains are possible.
In relation to the turbine blade performance, the boundary layer control appears to have useful applications in conjunction with turbine blade cooling.
One of the most effective methods of cooling the trailing edge region of turbine blades is by ejecting cooling air at or near the trailing edge of the blade. This ejected air can be used at the same time as an effective method of boundary layer control for increasing the blade loading and thereby reducing the total blade area for a given power output. The reduction in blade surface may, in turn, be expected to reduce the amount of cooling required to maintain a given blade temperature.
A further application of the BLC has been made for turbines requiring very high specific work with limited rotor blade speed, where negative reaction may be required. The flow through the turbine rotor is then similar to that through a typical compressor rotor and, if the blade loading is to be high, boundary layer control may be required to prevent stalling of the blade. In this case, as in the more usual case, boundary layer
Collins Timothy D.
Dykema Gossett PLLC
Poon Peter M.
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