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Publication numberUS3128939 A
Publication typeGrant
Publication dateApr 14, 1964
Filing dateJul 19, 1960
Publication numberUS 3128939 A, US 3128939A, US-A-3128939, US3128939 A, US3128939A
InventorsJoseph Szydlowski
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Szydlowski
US 3128939 A
Abstract  available in
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Claims  available in
Description  (OCR text may contain errors)

April 14, 1964' J. SZYDLOWSKI 3,128,939

AXIAL-FLOW COMPRESSORS Filed July 19, 1960 United States Patent O F 3,128,939 AXIAL-FLOW COMPRESSORS Joseph Szydlowslri, Usines Turbomeca, Bordes, France Filed July 19, 1960, Ser. No. 43,750 Claims priority, application France Feb. 3, 1960 8 Claims. (Q1. 230-134) The present invention relates to improvements in the wheels of axial-flow compressors, and more particularly of such compressors which operate at ultrasonic and supersonic speeds and comprise a wheel or wheels equipped with blades the camber-line of which is curved in opposite directions at the tip and at the root thereof, such as described in the applicants US. Patent No. 2,915,238. Such axial-flow compressors operate with acceptable efficiencies when the Mach number of their flow speed at the tip of the blades is high.

The present invention has for main object to provide means for increasing such efficiencies for equal Mach numbers with respect to those of said axial-flow compressors while permitting the obtention of acceptable efliciencies for higher Mach numbers.

For this purpose, according to the invention, at least one cross-section of the blade and more particularly the cross-section at the tip of the blade has a profile with two successive opposed curvatures while the cross-section at the root has a profile with a single curvature. It is thus possible to obtain an increasing in efficiency rang ing from 3% to 4% with respect to the compressors described in the aforementioned US. patent.

A further object of the invention is to provide jet engines, gas turbine engines and other similar machines, more particularly aviation machines, with compressors having rotor wheels of the type specified above.

Conveniently, an axial-flow wheel such as specified above is used as the single wheel of a compressor or as the first stage wheel of a multi-stage compressor, or as the single or first stage Wheel of the low-pressure compressor of a jet engine, a double-flow jet engine or other similar machines having more than one compressor.

The invention will be better understood and other features thereof will appear by means of the following description of a practical embodiment of the invention, reference being made to the accompanying drawings, in which:

FIG. 1 is a perspective view of an axial-flow compressnr wheel constructed according to the invention.

FIG. 2 is a radial view of a blade.

FIG. 3 is a partial developed view of a cylindrical section of the wheel adjacent the blade roots along the line IIIIII shown in FIG. 1.

FIG. 4 is a partial developed view of a cylindrical section of the wheel adjacent the blade tips along the line IVIV shown in FIG. 1.

FIG. 5 is a view similar to that of FIG. 2 for a modification of the blade shown in said FIG. 2.

Referring to FIG. 1, a compressor Wheel 2 is provided with a plurality of blades 1. The centers c3 and c4 of the cross-section 3 and 4 of each blade (FIGS. 3 and 4) through planes parallel to the plane tangent to the rotor at the blade root and passing through the lines III-III and IVIV of FIG. 1 are substantially aligned from the blade root to the blade tip along the radius passing through the center c3 of the blade foot.

Each blade 1 has a trailing edge 6 and a leading edge 5 which curve in opposite directions, the trailing edge 6 having a spanwise curvature from the radial inner end of the blade to its radial outer end in a counter rotating direction with respect to the direction of rotation F of the wheel while the leading edge 5 has a spanwise curvature in said direction of rotation. Such curvatures are so determined that the blade is flared at its tip.

3,128,939 Patented Apr. 14, 1964 As illustrated in FIGS. 3 and 4, the chordal pitches p3 and p4 of the cross-sections 3 and 4 numerically decreases from a positive value p3 from the root to a positive value p4 at the tip of the blade, whereby the chord 78 of the cross-section 3 at the blade root is slightly inclined to the axis of rotation A of the wheel 2 in the direction F of the wheel rotation. The chords of the successive cross-section from the blade root to the blade tip are more and more inclined to said axis so that the chord 910 of the cross-section 4 at the blade tip forms a substantially greater angle with said axis A than does the cross-section at the root.

At the blade root the cross-section has a single curvature and a concave-convex shape, the concave curve C being on the pressure side of the blade and the convex curve on the back side thereof. Near the blade tip the cross-section has a configuration with two successive opposed curvatures. The curve C, on the pressure side of the blade comprises, close to the leading edge 5, a convex curve portion C and close to the trailing edge 6 a concave curve portion C The curve on the back side of the blade near the blade tip has a concave portion close to the leading edge and a convex portion close to the trailing edge. The center of curvature 0 of the concave curve C is located with respect to the blade sections in the direction F of the wheel rotation as well as the center of curvature 0 of the concave curve portion C close to the trailing edge of the blade. On the contrary the center of curvature 0 of the convex curve portion C close to the leading edge of the blade is directed in the direction opposite to said direction F. The intermediate cross-sections of the blade change progressively from the single curvature cross-section C to the twin-curvature cross-section C The tip of the blade is located in a plane substantially parallel to the plane tangent to the body at the blade root and has an outline similar to that of said twin-curvature cross-section C The Wheel of an axial-flow compressor provided with blades of the character described has an increasing in efficiency ranging from 3% to 4%, with respect to that of a compressor according to the aforesaid US. patent, other things being the same.

Conveniently, the compressor wheel of the invention is more particularly used with advantage in double flow jet engines and in jet engines of the type described in the following US. patents in the name of applicant, No. 2,397,060, No. 2,795,372 and No. 2,922,278, in the US. patent application in the name of applicant Ser. No. 653,148 filed April 16, 1957, now abandoned and corresponding to the French Patent No. 1,012,233 patented April 9, 1952 or in the following French patents in the name of applicant: No. 907,666 patented July 9, 1945, No. 908,121 patented August 13, 1945, No. 915,552 patented July 22, 1946, No. 1,046,175 patented July 8, 1953 and No. 1,107,179 patented August 3, 1955.

Of course modifications may be made to the described embodiment without departing from the scope of the invention as defined in the appended claims. Thus, for example and as illustrated in FIG. 5, the curve portions C and C forming the curve C, with a twin-curvature instead of being joined together tangentially may be united by a rectilinear portion 11 which is tangent to both curve portions C and C What I claim is:

1. A wheel for an axial flow compressor adapted for operating at ultrasonic and supersonic speeds, said wheel comprising: a body adapted for rotation about an axis, and a plurality of blades supported on said body, said blades having a root at the location where said blades contact said body and a remote tip, said blades also having leading and trailing edges, each blade in crosssections in planes parallel to a plane tangent to the body at the blade root having centers which are substantially aligned from the blade root to the blade tip along a radial line, said blades each having a chordal pitch which uniformly decreases from a relatively large positive value at the root of the blade to a relatively small positive value at the tip of the blade, said cross-sections of each blade progressively varying in outline as viewed from a pressure side of the blade from a concave curve at the root of the blade to a curve having successive convex and concave portions at the blade tip, said convex and concave portions being respectively disposed adjacent the leading and trailing edges.

2. A wheel as claimed in claim 1 wherein the trailing edge of each blade has a spanwise curvature from the radial inner end to the radial outer end of the blade in a counter rotating direction with respect to the rotary direction of the wheel while the leading edge of the blade has a spanwise curvature in said rotary direction, said curvatures being sutficient to cause the blade to be flared at the tip thereof.

3. A wheel as claimed in claim 1 wherein the crosssections of each blade at the root is concave on the pressure side of the blade and convex on the other side of the blade and said cross-sections vary to a shape at the tip respectively having on pressure and back sides convex-concave and concave-convex outlines.

4. A wheel for an axial-flow compressor adapted for operating at ultrasonic and supersonic speeds, said wheel comprising: a body, said body being adapted for rotation about an axis, and a plurality of blades supported at the roots thereof to said body and having a remote tip, said blades further having leading and trailing edges, the trailing edge of each blade having a spanwise curvature from the root to the tip in a counter-rotating direction relative to the rotary direction of the wheel, while the leading edge has a spanwise curvature in said rotary direction, said curvatures being suflicient to cause the blades to be flared at the tips thereof, the centers of the cross-sections of each blade in planes parallel to a plane tangent to the rotor at the blade root being substantially aligned from the blade root to the blade tip along a radial line passing through the center of the blade root, said blades having a chordal pitch numerically decreasing from a relatively large positive value at the root to a relatively small positive value at the tip of the blade while the outline of said cross-sections progressively vary on the pressure side from a concave curve at the blade root to a curve having successive convex and concave portions at the blade tip, said convex and concave portions being respectively disposed close to the leading and trailing edges.

5. An axial compressor having at least one stage with at least one wheel and intended for operation at ultrasonic and supersonic speeds, at least one of the wheels of at least one stage of the compressor including blades supported at the roots thereof on said one Wheel, said blades having leading and trailing edges, the centers of the cross-sections of each blade in planes parallel to a plane tangent to the wheel at the blade root being substantially aligned from the blade root to the blades tip along a radial line passing through the center of the blade root, said blades each having a chordal pitch of said crosssections which numerically decrease from a relatively large positive value at the root to a relatively small positive value at the tip of the blade while the outline of said cross-sections progressively vary as viewed on the pressure side of the blade from a concave curve at the blade root to a curve having successive convex and concave portions at the blade tip, said convex and concave portions being respectively disposed close to the leading and trailing edges.

6. A compressor comprising at least one stage operating at ultrasonic and superosnic speeds, said compressor constituting the inlet compressor of a jet engine, said compressor having at least one wheel in at least one stage of said compressor, said wheel having a plurality of blades including roots and tips and leading and trailing edges, the centers of the cross-sections of each blade through planes parallel to the plane tangent to the wheel at the blade root being substantially aligned from the blade root to the blade tip along a radial line passing through the center of the blade root, said blades having a chordal pitch at said cross-sections which numerically decreases from a relatively large positive value at the root to a relatively small positive value at the tip of the blade while the outline of said cross-sections progressively vary as viewed from the pressure side of the blade from a concave curve at the blade root to a curve having successive convex and concave portions of the blade tip, said convex and concave portions being respectively disposed close to the leading and trailing edges.

7. A compressor comprising at least one stage with at least one wheel operating at ultrasonic and supersonic speeds and acting as low pressure compressor for a double flow jet engine, at least one of said wheels or at least one stage of said compressor having blades including roots and tips and leading and trailing edges, the centers of the cross-esctions of each blade through planes parallel to the plane tangent to the wheel at the blade root being substantially aligned from the blade root to the blade tip along a radial line passing through the cen ter of the blade root, said blade having a chordal pitch at said cross-sections which numerically decreases from a relatively large positive value at the root to a relatively small positive value at the tip of the blade while the outline of said cross-sections progressively vary as viewed from the pressure side of the blade from a concave curve at the blade root to a curve having successive convex and concave portions at the blade tip, said convex and concave portions being respectively disposed close to the leading and trailing edges.

8. A compressor having at least one stage operating at ultrasonic and supersonic speeds and forming the inlet compressor of a gas turbine, said compressor comprising at least one wheel forming at least one stage of said compressor and having blades including roots and tips and leading and trailing edges, the centers of the crosssections of each blade through planes parallel to the plane tangent to the wheel at the blade root being substantially aligned from the blade root to the blade tip along a radial line passing through the center of the blade root, said blades having a chordal pitch at said crosssections which numerically decreases from a relatively large positive value at the root to a relatively small positive value at the tip of the blade while the outline of said cross-sections progressively vary as viewed from the pressure side of the blade from a concave curve at the blade root to a curve having successive convex and concave portions at the blade tip, said convex and concave portions being respectively disposed close to the leading and trailing edges.

References Cited in the file of this patent UNITED STATES PATENTS 2,269,287 Roberts Jan. 6, 1942 2,628,768 Kantrowitz Feb. 17, 1953 2,667,936 Clark Feb. 21, 1954 2,784,551 Karlby et al Mar. 12, 1957 2,915,238 Szydlowski Dec. 1, 1959 2,931,563 Eggleton Apr. 5, 1960 2,935,246 Roy May 3, 1960 2,955,746 Stalker Oct. 11, 1960

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2269287 *Nov 29, 1939Jan 6, 1942Roberts Wilmer SFan
US2628768 *Mar 27, 1946Feb 17, 1953Arthur KantrowitzAxial-flow compressor
US2667936 *Sep 16, 1950Feb 2, 1954William F ClarkBoat propeller
US2784551 *Jun 1, 1951Mar 12, 1957George Zilliac EVortical flow gas turbine with centrifugal fuel injection
US2915238 *Oct 18, 1954Dec 1, 1959Joseph SzydlowskiAxial flow compressors
US2931563 *Sep 10, 1956Apr 5, 1960Frederick EggletonConstruction of axial flow compressors
US2935246 *Jun 1, 1950May 3, 1960Onera (Off Nat Aerospatiale)Shock wave compressors, especially for use in connection with continuous flow engines for aircraft
US2955746 *May 24, 1954Oct 11, 1960Stalker Edward ABladed fluid machine for increasing the pressure of a fluid
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US3724968 *Mar 23, 1971Apr 3, 1973Cit AlcatelAxial supersonic compressor
US3867062 *Sep 19, 1973Feb 18, 1975Theodor H TrollerHigh energy axial flow transfer stage
US4012165 *Dec 8, 1975Mar 15, 1977United Technologies CorporationFan structure
US4373861 *Apr 16, 1980Feb 15, 1983Papst-Motoren KgAxial-flow fan
US4564335 *Apr 2, 1985Jan 14, 1986Papst-Motoren Gmbh & Co. KgAxial flow fan
US4682935 *Dec 12, 1983Jul 28, 1987General Electric CompanyBowed turbine blade
US5044885 *Mar 1, 1990Sep 3, 1991Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A."Mobile blade for gas turbine engines providing compensation for bending moments
US5137417 *Jun 12, 1991Aug 11, 1992Lund Arnold MWind energy conversion system
US6164919 *Dec 14, 1998Dec 26, 2000Vanmoor; ArthurPropeller and impeller blade configuration
US6168384 *Dec 14, 1998Jan 2, 2001Arthur VanmoorPropeller blade configuration
US6910861 *Nov 15, 2001Jun 28, 2005Leybold Vakuum GmbhTurbomolecular vacuum pump with the rotor and stator vanes
US6981839Mar 9, 2004Jan 3, 2006Leon FanWind powered turbine in a tunnel
US7195456Dec 21, 2004Mar 27, 2007United Technologies CorporationTurbine engine guide vane and arrays thereof
US8337164 *Jul 9, 2009Dec 25, 2012Osaka Vacuum, Ltd.Turbomolecular pump
US20110008176 *Jul 9, 2009Jan 13, 2011Tetsuro OhbayashiTurbomolecular pump
Classifications
U.S. Classification416/242
International ClassificationF04D29/32, F04D21/00
Cooperative ClassificationF04D21/00, F04D29/324
European ClassificationF04D21/00, F04D29/32B3