US3442441A

US3442441A - Supersonic cascades

Info

Publication number: US3442441A
Application number: US654837A
Authority: US
Inventors: Wilhelm Dettmering
Original assignee: Individual
Current assignee: Individual
Priority date: 1966-07-21
Filing date: 1967-07-20
Publication date: 1969-05-06
Anticipated expiration: 1986-05-06
Also published as: GB1198515A; DE1628237C3; DE1628237B2; DE1628237A1

Description

W. DETTMERING May 6, 1969 SUPERSONIG CASCADES Sheet Filed July 20, 1967 RIGHT ANGLE THRUS T INVENTOR Wilhelm Dettmerz'r BY Y M, fl'mufl' ATTORNEYS y 1969 Q w. DETTMERING 3,442,441

SUPERSONIC CASCADES I lfiled July 20, 1967 Sheet 13 of 2 INVENTOR BY W 4 144 fiM ATTORNEYS United States Patent O Int. Cl. F04d 29736; F01d 5/06 US. Cl. 230132 6 Claims ABSTRACT OF THE DISCLOSURE The gases flowing through a turbine or a compressor are reduced from supersonic to subsonic flow by being passed through a row of shock and deflecting blades, and then through a row of gas deceleration blades which partially lap the shock and deflecting blades.

Because the stator of a turbine engine as, for example, an axial compressor is subject to supersonic gas velocities, it is possible to increase considerably the specific work turnover and thereby considerably reducing the crosssectional areas and the number of stages, and thus the axial length and material required for the obtaining of the desired increase in pressure.

While the supersonic gas flow can be obtained with relatively good efficiencies in an acceleration cascade, high losses occur in the heavy deflecting and deceleration cascades, such as disclosed in German Patents Nos. 759,892, 866,793 and 760,327, in which the gas flow is reduced to subsonic velocities by means of one or more shocks which preclude or heavily limit an economic use.

Aside from the unavoidable shock losses, there occurs a separation of the gas flow on the blade contours which considerably impairs the efliciency of a cascade. These losses which are influenced and caused by the boundary layer, by the interaction between shock and boundary layer, and the deflection and deceleration to subsonic flow are considerably lessened according to this invention. According to this invention, the cascade is composed of a row of shock blades and a row of deceleration blades which partially lap each other and the blades are positioned with respect to one another, as well as the contours of the blades so that separation of the gas flow cannot occur.

Double rows of blades have been used for the subsonic velocities, as disclosed in German Patents Nos. 390,486, 459,204, 573,799 and 760,327. However, such double rows cannot be used for supersonic gas velocities since the blade contours and the duct cross-sections do not meet the characteristics of supersonic flow.

The means by which the objects of the invention are obtained are described more fully with reference to the accompanying schematic drawings in which:

FIGURE 1 is a cross-sectional view through a cascade showing one form of the invention;

FIGURE 2 is a similar view of a modified form of the invention; and

FIGURE 3 is another similar view of a further modified form of the invention.

The very thin leading edges 1 and the trailing edges 2 of the shock blades 9 create only weak head waves with subsequent diagonal thrusts which produce no substantial deceleration and cause but little loss in static pressure. The vertical shock of the inflowing gas C is stabilized by means of throttling within the area A so that a uniform cross-section of the passage in area A is required and at most the passage is permitted to expand only slightly. The stream lines are not curved or only slightly so since the vertical shock disturbs the equilibrium of the centrifugal force and pressure gradient. The deflection is accomplished after the shock, that is with subsonic velocities substantially in the area B. The losses connected wtih this are small since the boundary layers are relatively thin and since a deceleration and subsequent detachment of the flow is avoided by means of the approximate uniform cross-section. The deflection is accomplished most appropriately as a potential vortex. After the deflection and for the purpose of building up the pressure, the cross-section in areas C and D is enlarged and therefore the flow decelerated. The deceleration blade 10 extends into the deflection area B to form areas C and D and is a condition for the exact adherence to the cross-sections which are suitable for the flow. The trailing edges 2 of the shock cascade blades are relatively thin in order to achieve a uniform annulus diagram and to minimize the low-energy wake flow.

The leading edges 3 of the deceleration cascade blades 10 are also feather-edged in order to avoid localized supersonic velocities. They are exactly tailored to the blades of the shock cascade and form a smooth blending of the cross-section, at the points where they overlap from areas B to C and D. The

contours

7 and 8 of the deceleration blades 10 form, with the contour on the suction side 6 and on the pressure side 5 of the adjacent shock blades 9, a small angle of expansion of approximately 1 to 2 in area D and up to 4 in area C. The expansion of the cross-section in area D is held small because of the danger of detachment due to the continuation of the deflection which is necessary in this type of cascade construction for geometric reasons and with due consideration for pressure equilibrium. The curvature in area B produces, by means of centrifugal force, a pressure gradient which, after the separation of the streams of gas flow in area C, effects a higher means pressure than in area D. So that no pressure diflference can prevail at the confluence point of the two partial streams passing through areas C and D which could lead to a disturbance which could cause a loss, the area D continues and ends with an inverted curvature which again produces a pressure gradient which reaches at edge 2 the mean pressure of the parallel flow in area C which is also freed from a pressure gradient so that a uniform pressure and approximate parallel flow prevails upon the surface of discontinuity. Therefore, a small pressure gradient remains only in area D which, however, causes no mentionable disturbance. In order to hold the pressure difference of the two partial flows low, it is not possible to expand the cross-sectional area C at will. The deceleration mainly takes place in the area E which is kept as short as possible. At this point, a relatively high angle of expansion can be provided since the boundary layers are thin. The accumulated boundary layers of shock blades 9 flow into the sound core of the flow and do not cause any appreciable loss. The outflowing gases C pass between the trailing edges 4.

The curvature proposed for the equilization of pressure on the surface of discontinuity in passage E can be eliminated if the geometry of the cascade is otherwise the same and where the construction for a smaller deflection in the cascade does not result in a too high pressure difference. In this case, the suction-surface contour 6 of the shock blades 9 would extend in a straight line as shown in FIGURE 2. The losses will remain low since a type of stagnation point flow is formed on the trailing edge 2 which guarantees a continuous transition.

-In the modification of FIGURE 2, the parallel direction of the two partial flows on the surface of discontinuity from edges 2 is eliminated. However, the additional losses produced by this fact are neutralized by the low inner deflection in the potential vortex area with the angle of inlet and outlet remaining the same. Thus, with identical effective deflection, the danger of detachment is reduced and the cascade shortened. The contour 7 of the deceleration blade is on one side constructed as an almost straight line while the curvature required for the thickness distribution on the contour 8 produces a further deflection.

In the modification of FIGURE 3, the shock is separated from the main deflection and the subsonic deceleration. The last two are both placed in the second subsequent deceleration cascade where the boundary layers have not yet formed any dangerous thickness. As far as the construction of the cross-sections is concerned, the above-mentioned points are valid.

The shock cascade blades 9 are very thin over their entire length since otherwise even with a small positive angle Ant of attack and/or transonic inlet flow velocity the cascade can be blocked, since the supersonic flow conditions are not maintained. Small angle [3 of the leading edges already cause a pre-compression which has little loss, whereby the efficiency of the cascade is enhanced rather than when the velocity change would be accomplished only in a vertical shock. However, because a detached head shock can occur which is connected with high losses, the angle ,8 cannot be made much more than 6. The actual vertical shock is stabilized in the vicinity of the narrowest cross-section in order to obtain minimum static pressures. A low deflection of the flow by is also accomplished in the shock cascade by designing one contuor side straight, as has already been described for the cascade in FIGURE 2. An insignificant deceleration in area B caused by the small angle 7 of expansion does not lead to any detachment.

The blades of the shock cascade 9 have in correspondence with their task only a small length so that, despite the shock, no dangerous thick boundary layers are produced. The boundary layer flows into the sound stream between the deceleration blades 10. The deceleration blades do not need to extend so far into being lapped with the shock cascade since, with the small angle 7 of the trailing edge 2 of the shock blades, no greater expansion of the cross-section occurs which must be avoided.

Sharp leading edges 3 and small leading edge angles are needed because of the desired constant transition of the cross-section without restriction. The main deflection takes place in the area E having an almost constant cross-section and with subsonic flow. The further deceleration and pressude build-up occurs in the difiusion part D of the deceleration cascade.

In the examples given, the flow has always been discharged in an axial direction. This is not entirely necessary inasmuch as constructions are possible which have a flow component in the direction of the circumference. However, in order to fit to a subsequent stage or for achieving a discharge loss which is as small as possible, an axial discharge is most appropriate in the majority of cases.

Even though the example given relate to a compressor, this invention also applies to a turbine subject to super sonic flow velocities and indeed also as the last stage for the purpose of retarding the high velocity discharge and for deflecting the flow in the axial direction so as to considerably reduce the discharge losses.

This invention also applies to stationary as well as rotating cascades for considerably diminishing flow losses and realizing a highly effective specific work turnover in which the additional losses produced by the separation of the flow, by thick boundary layers and by the interaction between shock and boundary layer are at least held to a minimum or avoided. Furthermore, it is possible to obtain larger deflections and decelerations.

Having now described the means by which the objects of the invention are obtained,

I claim:

1. A deceleration lattice for supersonic fluid inflow in the stators and rotors of axial flow compressors and turbines comprising two fixed staggered blade rows arranged in series and with one row lapped with and staggered downstream of the other, said staggered blade rows forming an upstream shock inducing cascade row and a downstream subsonic flow producing cascade row, the blades in the upstream shock cascade row having sharp leading and trailing edges and being spaced from each other to form a fluid flow passage starting with a straight entry zone portion followed by a stabilizing slight narrowing of said zone portion and further followed down- I stream by a short curved deflection zone portion having a nearly constant deflecting cross-sectional area for deflecting the fluid flow toward said subsonic cascade row and said deflection zone portion ending at the leading edge of the blades in said subsonic cascade row, the blades in said subsonic cascade row being spaced from each other and overlapping the trailing portions of the upstream shock inducing cascade to form a fluid velocity deceleration cross-sectional area larger than said deflecting cross-sectional area, and sharp leading and trailing edges on said subsonic blades for giving an almost constant crosssectional distribution of the flow passage.

2. A lattice as in claim 1, each shock cascade blade being concave on the trailing edge portion of its suction surface, and each subsonic cascade blade being convex on its pressure surface for causing a deflection of fluid flow in a direction opposite to that of the main deflection of the fluid flow through the entire lattice.

3. A lattice as in claim 2, said passage having a widened cross-sectional area in the portion lapped by the blades for obtaining an initial deceleration of the fluid flow.

4. A lattice as in claim 1, each shock cascade blade being rectilinear on its suction surface portion lapped by the subsonic cascade blade, and the subsonic cascade blade being rectilinear on its pressure surface portion lapped by the shock cascade blade.

5. A lattice as in claim 1, said shock cascade blades together with said subsonic cascade blades forming a blade set with each shock cascade blade suction surface terminating with approximately the outlet angle of the entire blade set and said shock cascade blade having a thickness for forming a different angle on its pressure surface, said ditference being compensated for the fluid flow deflection into said subsonic cascade row.

6. A lattice as in claim 1, each shock cascade blade having a rectilinear pressure surface adjacent its leading edge and forming a small angle with its pressure surface for producing an oblique shock starting at said leading edge in addition to the normal shock with only a weak deflection due to the finite thickness of the blade, and with the main deflection occurring in said subsonic cascade row.

References Cited UNITED STATES PATENTS 1,771,711 7/1930 Hahn 10397 2,839,239 6/1958 Stalker 230- 2,974,927 3/1961 Johnson 230-120 3,156,407 11/ 1964 Bourguard 230120 3,356,289 12/1967 Plotkowiak 2301 14 HENRY F RADUAZO, Primary Examiner.

U.S C1. X.R. 230-120; 253-78