US2830754A - Compressors - Google Patents

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US2830754A
US2830754A US39571553A US2830754A US 2830754 A US2830754 A US 2830754A US 39571553 A US39571553 A US 39571553A US 2830754 A US2830754 A US 2830754A
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rotor
blades
blade
stage
downstream
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Edward A Stalker
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Edward A Stalker
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE DISPLACEMENT PUMPS
    • F04D27/00Control, e.g. regulation, of pumps, pumping installations or systems
    • F04D27/02Surge control
    • F04D27/0207Surge control by bleeding, bypassing or recycling fluids
    • F04D27/0215Arrangements therefor, e.g. bleed or by-pass valves
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE DISPLACEMENT PUMPS
    • F04D19/00Axial-flow pumps
    • F04D19/02Multi-stage pumps
    • F04D19/028Layout of fluid flow through the stages
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE DISPLACEMENT PUMPS
    • F04D27/00Control, e.g. regulation, of pumps, pumping installations or systems
    • F04D27/02Surge control
    • F04D27/0207Surge control by bleeding, bypassing or recycling fluids
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE DISPLACEMENT PUMPS
    • F04D29/00Details, component parts, or accessories
    • F04D29/66Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F04D29/682Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps by fluid extraction
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE DISPLACEMENT PUMPS
    • F04D29/00Details, component parts, or accessories
    • F04D29/66Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F04D29/684Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps by fluid injection

Description

April 15, 1958 E. A. STALKER COMPRESSORS 3 Sheets-Sheet l Original Filed Dec. 26, 1947 E64 INVENTOR. z /4m APril 1,5, 1958 E. A. STALKER 2,830,754

COMPRESSORS Original Filed Dec. 26, 1947 3 Sheets-Sheet 2 $6 5637' asf \39 @dm/M CGMPRESSORS Edward A. Stalker, Bay City, Mich.

riginal application December 26, 1947, Serial No. 794,018, now Patent No. 2,749,027, dated .lune S, 1956. Divided and this application December 2, 1953, Serial No. 395,715

My invention relates to compressors.

This application is a division of my application Serial No. 794,018 tiled December 26, 1947 now Patent No. 2,749,027. It distinguishes from the earlier application in that it is directed to the relative positioning of axial ow blades in a stage of either rotor or stator blades and to f the relation of a stator and rotor stage. l

An object of the invention is to provide an axial flow compressor which maintains its pressure and eiciency over a wider range of volume flow per revolution.

Another object is to provide a compressor having proper relations between stator and rotor blades to reduce the range of angles of approach which a bladerequires.

Another object is to provide a multistage compressor having a combination of axial flow stages of one type with an axial ow stage of another type to ameliorate the great loss in eiciency at oli-design conditions when the compressor has been designed for a high compression ratio.

Other objects will appear from the drawings, specification, and claims. v

The above objects are accomplished by the means illustrated in the accompanying drawings in which:

Figure l shows a vector `diagram for the air approaching a rotor blade; i

Figure 2 showsa vector diagram for the air approaching a rotor blade at a dat angle;

Figure 3 is a chordwise section along the line 3--3 in Fig. 4;

Figure 4 is an axial section through an axial flow compressor according to this invention;

Figure 5 is a fragmentary diagrammatic development of some of the stages of the compressor of Fig. 1 with the blades shown solidralthough they are hollow inthe machine; y

Figure 6 is a section along line 6-6 in Fig. 4;

Figure 7 is a fragmentary development of the last stage of the compressor to show the vector relations;

Figure 8 is a fragmentary development of the last -rotor showing blades in section as they are in the machine;

Figure 8a is a fragmentary development as an alternate to that of Figure 8 showing blades of parallel sides;

Figure 9 is an alternate rotorconstruction to, that` ot' Fig. 8; v

Figure l0 is an `enlarged fragmentary axial section through a part of the last rotor and part of the case of the compressor of Fig. 4; p

Figure ll `shows an isolated group of blades and a connecting duct, one blade having `an Vinduction slot and the other having a` discharge slot; u

Figure 12 `is a fragmentary axial section through another compressor in which the rotor and stator blades have discharge slots and the other walls 'have slots for facilitating the entrance of-'a supersonic ow into the passages between blades; andf- Figure 13 isasection along line 13 -13 in Fig.`12. `v When a multi-stage axial ttor.' compr ser is operating yUnited a States Patent O stages.

2,830,754 Patented Apr. l5, 1958 riice.

at a mass flow per revolution less than the optimum or design value with a back pressure that is relatively low the axial velocity in the downstream stages may be as much as three times the velocity which would prevail at the optimum or design condition. This is so because the upstream stages do a certain amount of compressing at olf design conditions and the lack of back pressure permits the ow compressed by the up stream stages to stream at greatly increased velocity through the later This leads to a great change in the direction of the Huid approaching a later rotor or stator with respect to the direction for the optimum operating condition, reducing the angle of attack of the blades and their compressing ability.

For instance Figure 1 shows the vector diagram for a conventional axial flow compressor for a downstream stage when fthe compressor is operating under optimum condition, that is at about best efciency and corresponding pressure ratio. In this instance the axial velocity for optimum operation is CM equal to a fraction of u the peripheral velocity. Under this condition the direction of the lluid leaving the stator blade 1 and approaching therotor blade 2 is the vector 4. Now if the axial velocity is increased to 3 times CM the new direction is the vector 6 and the change in the angle of approach is Aal which is equal to about 30. This is a greater range of angles of attack than a blade can accommodate.

Now consider a case as in Fig. 2 where the leaving velocity vector from blade 10 is C directed at the positive angle B toward the rotor blade 12. The resultant vector is 14. if the axial component of C' is increased from CM as ffor Fig. l to 3CM the new resultant velocity vector is 20, whose peripheral component is much larger than that of Vector 14. The peripheral component is not magnied as greatly as the axial since the new triangle of which 20 is the longer side is not symmetrical with respect to the triangle lllcu. The change in angle of approach to blade 12 is now A112 equal to about 7. This is not only within the range of angles which the blade can accommodate but is also well within the range of angles of attack for best el'liciency of the blade itself.

It is thus shown that deiiecting the air toward the oncoming rotor blades reduces the range of approach angles or angles of attack which the blade must accommodate when the compressor is operating at a pressure and speed substantially below optimum conditions provided the deliection through the angle B is accompanied by a rise in velocity.

The angle B for the vector representing the entering vector for a rotor (or stator) is positive when the vector has a peripheral component directed toward the concave face of the blade of the succeeding stage. Thus in Fig. 2 B is positive. Also in Fig, 3, B is positive since the vector 32 approaching the stator attacks the concave side of the blade 86.

The range of approach angles which can be accommodated by the downstream stages can also be extended to a considerable extent by making the noses of the blades successively thicker in successive stages in the downstream direction. Thus in Fig. 3 the nose 30 is substantially semicircular so that the relative flow will he able to flow about the nose without burbling when the approach vectors vary from vector 32 to vector 34 disposed angularly with respect to each other by the angle (delta).

To further encourage the ow the nose is provided with the slots 36 and 38 (Figs. 3, 4, and 8) through which a ow may be inducted to control the boundary layer.

Since the fluid is compressed in successive stages the temperature rises along the compressor axis. Consequently the velocity of sound in the lluid increases in magnitude so that the velocity of the Huid relative to the blades can be increased without precipitating a compressibility shock. In other words the local velocity on the blade surfaces can be higher on the downstream blades without reaching the critical Mach number of one.

Thickening the nose of the blades makes possible a wider range of angles (see Fig. 3) but increases the local velocity on the nose. However by taking advantage of the rise of temperature from stage to stage, the noses of the blades of successive stages may be thickened without the local Mach number exceeding the critical value.

Figs. 4 to 8 show a compressor incorporating the foregoing features. Figs. 5 and 8 illustrate the preferred range of positions of one blade 94 relative to another. The maximum thickness of each blade is preferably at about mid-chord and nearer thereto than to the normal projection of the leading edge of the following blade on the adjacent leading blade.

In Fig. 4 the compressor is indicated generally by 40 comprised of the case 4-2 the rotor stages 41-46 and the stators S1-56. (See also Fig. 5). Fluid enters the inlet 60 and is pumped through the anular or main ow passage 62 to the exit passage 64.

At the upstream end (Fig. 4) the stator 51 dellects the incoming air by means of the stator blades 66 in the direction of rotation 67 of rotor stage 41 composed of blades 68. The next stator 52 also deects the fluid in the direction of rotation of rotor stage 42, but to a less extent, by blades 70. At the third stage the stator 53 deects the fluid substantially axially toward the rotor 43. This stage is comprised of blades 74 and 76.

In the succeeding stages of Fig. 4 the stator blades deflect the flow with increasing peripheral velocity components against the direction of motion of the rotor blades.

The blades of the fourth stage are 78 and 80 and the blades of the fifth stage are 82 and 84. It is to be noted that in each of these stator stages (see Fig. 5) and in the sixth stator stage the stator blades are curved to give the flow a progressively greater peripheral component in successive downstream stages.

The stator blades 86 for instance in the sixth stage have tail portions directed substantially in the peripheral direction.

In Fig. 7 the velocity vector 90 leaving the blades 86 when combined with the peripheral velocity vector n of the rotor gives the velocity vector 92 acting relative to the rotor 46. The vector 90 makes the positive angle B with the axial direction and hence even for a great increase in axial velocity through the compressor the direction of 92 relative to the blades 94 of rotor 46 will change only a small amount in direction.

The rotor blades 94, Fig. 8 may be hollow and as shown in Figs. 4 and ll each has its interior in communication by means of individual ducts 96 with the hollow blades 76 of the third stage. Since the fluid pressure is greater in the sixth stage than in the third stage fluid will enter the blades 94 through slot 95 and be discharged through the discharge slots 98 in blades 76. Thus the flow is induced to follow the curved portion of blade 94 making it possible to discharge the flow from the stage with a velocity direction closely perpendicular to the plane of rotation. The slots may also be omitted as shown in Fig. 5 with some small loss in effectiveness.

The last set of stators 100 (Fig. 4) takes out the peripheral component of velocity relative to the case 42 and directs the discharge of fluid axially along the passage 64.

As an alternate form the rotor may be formed as in Fig. 9. Here the blade is made in two parts, the fore part 102 and the aft part 104 spaced from the fore part to provide the slot l06. The ow through the slot provides a jet to control the boundary layer on the convex portion of the blade and induces the flow in the passage 108 between blades 102 to follow the blade surface.

The stators as shown in Fig. 4 are also interconnected by ducts such as 109 to provide for flows of uid through the blade slots. This construction is similar to that shown in my U. S. Patent No. 2,344,835 issued March 2l, 1944.

By providing the stator which gives a large positive angle B, the variation (Fig. 3) is kept small and consequently the blades 94 (Figs. 4 and 5) may be thin at the nose and particularly efficient for high velocities of flow.

As shown in Figs. 4 and l0, particularly the latter, the case 42 diverges from the wall 110 of the rotor 48 so that each passage 112 between blades 94 is expanding in cross sectional area until the locality of the blade curvature is reached where the passage area is preferably made to contract slightly so that the ilow about the curve is in a favorable pressure gradient. This facilitates an eiiicient iiow about the curve but is not essential. Even though the peripheral width of the ilow passage becomes less in the downstream direction, the cross sectional areas of the passages increase by reason of the divergence of the walls of the case, as described above.

There is also another advantage in the divergence of the hub and case walls. The increase in the cross sectional areas of the rotor passages in the downstream direction slows down the lvelocity of ow before the flow is turned by the blade. Hence the appearance of compressibility shock waves is delayed. That is, the peripheral tip speed of the blades can be higher before the shock wave appears in the passages between blades. This means that substantially greater pressure ratios can be obtained from a rotor.

The rst shock waves appear at the leading edge of a blade but the critical shock wave which limits the mass flow through the rotor occurs in the passage downstream from the nose of the blade.

If thel passages between blades begin to diverge radially opposite the blade noses, the radial expansion can compensate for the peripheral contraction due to the blade thickness. Hence there need not be a throat along the passages between blades or at least the throat may be placed far downstream from the inlet of each rotor passage. Thus as shown in Figs. 5 and 8 the maximum thickness may be in the neighborhood of mid chord or even further rearward. In this connection the blades may have substantially parallel sides as shown by blades 102 in Fig. 9.

The opposite sides of the blade sections such as blades 86 (Fig. 5) are substantially parallel along a substantial length between the nose portion and the aft portion.

Blades 94, Figs. 5 and 8, can also be modified to have parallel sides as shown by 94 in Fig. 8a providing for fabricating them from sheet metal. Since the thickness can be constant along the blade span the nose radius can be cut by a simple tool passing along the nose of the blade. The trailing edge can be made relatively sharp also by a simple tool operation. Thus the blade can be given its proper shape including faired nose and tail portions at low cost.

By making the last stage with thin blades and relatively sharp noses it can operate with very high fluid velocities without generating shock waves at the nose or in the passage. However in some applications the velocity may become supersonic in the last stage if the back pressure is reduced suciently when the rate of rotation of the rotor is near the optimum speed for the compressor as a whole. For this reason the type of rotor shown in the last stage is very advantageous since it can operate even at a supersonic velocity as has been disclosed in my application Serial No. 624,013 filed October 23, 1945 entitled Compressors, now Patent No. 2,648,493. Furthermore for a high performance compressor the last stage is preferably made to have a supersonic velocity of approach of the air at the optimum condition of operation. For such a compressor it is important that the angular range of the approach vector should be small to obtain the proper shock waves at the nose of the blades and within the rotor Aor stator passages.- These are provided by this invention. i v

In an axial flow compressor if the pressure rise is great between inlet and exit for the design condition, then the machine will be much less efficient at a lower delivery, that is at a lower value of the mass of iluid delivered per revolution. The greater the pressure rise, the greater the drop in efficiency at an oit-design delivery.

The compressor of this invention using the type of rotor ed is provided to assuagc this undesirable condition and places the axial ow compressor on a more favorable footing with respect to other compressors, such as for instance the centrifugal compressor, than heretofore existed,

When the fluid approaches the blades at supersonic values shock waves iirst appear at the leading edges of the blades and if the back pressure is substantial the shock waves may occur ahead of the leading edges and the iiov/ may refuse to enter the passages between the blades at high supersonic velocities. This difficulty can be overcome by discharge slots properly located with respect to the leading edges of the blades.

Figure `l2 shows an alternate structure for the last rotor and the stator ahead of it. The balance of the compressor ahead of this stator would have a structure similar to that of Figs. 4 and 5.

In Fig. l2 the slots 141) and 142 are located in peripheral walls, that is the shroud ring 143 and the hub wall respectively. That is the rotor blades 144 of the last rotor are encircled by the shroud ring and its leading edge forms the slot 140 with the case Wall 42.

Air for the case or outer wall slot 146 is bled from the passage 64 via the annular duct 150 formed in the case. The air is at a higher pressure in 64 than in the passage at the leading edge of blade 144 and hence can ow Aat a higher velocity from the slot 140 than the velocity of the local main flow.

Air is also supplied from duct 150 to the slot 152 positioned in the rotor passage 62 a substantial distance inward from the leading edge of blade 144.

Air is also supplied to the slot 142 and slot 156 from passage 6d via the annular ducts 160 and 162. Air also enters the hollow interior of blade 144 via 162 to serve the slot 154.

The discharge slots 171i and 172 of stator blade 174 is also served with air from duct 150. As shown in Fig. 13 this blade has a well rounded nose 176 and the discharge slots located near the ends of the nose contour.

The passages 112 in the rotor between the blades in Fig. l2 are similar to those in Fig. 8 and are bounded v by walls on four sides. VThe walls 11d of the hub of the rotor and the shroud ring 143 bound the passage on radially apposite sides while the adjacent blades bound the other two opposite sides. All of the Walls may have slots therein but preferably only hub and case walls and one blade have slots. The slots in opposite walls within the passages are preferably not directly opposite each other.

The blades discussed herein are to be considered thin blades if their maximum thickness is less than l percent of the blade section chord length.

In the preferred forms of the blades the chord-wise length of the blade, that is the dimension along the direction of flow is preferably vsmaller than the spanwise dimension or length or at least the chord is not more than twice the span. The blades also have free leading and trailing edges extending in the same general radialdirection. j

Axial ilow compressors have blade structures whose main flow passages extend in the general axial direction from an inlet at the front to an exit at the rear to discharge fluid in the general axial direction.

It is to be noted that the blades of the downstream rotor have blade sections of fair or streamline contours. That is there are no sharp changes of direction in them,

the fore portions of the blades fairing into the aft portions thereof as shown in Figs. 5, 8 and 9 particularly. Thus these sections will operate eiiciently at subsonic speeds.

As shown in Figs. 5, 8 and 9 the blades are placed close together so that they overlap in axial view over a substantial portion of their span. Each passage between the blades may be considered as bounded by a leading blade and a following blade (for succinctness in the claims). The normal projection of the leading edge of a following blade on a leading blade should fall at a point forward of its mid-chord point and substantially ahead of the maximum thickness of the blade section to assure that a rotor passage throat, if there is one, is a substantial distance downstream from the passage inlet, as remarked earlier. More briefly the leading edge of the following blade may be said to be preferably opposite the mid-chord point of the leading blade or ahead of this point. Since the maximum thickness is preferably at about mid-chord as remarked earlier, the maximum thickness is preferably nearer the mid-chord point than to said normal projection. With this construction the cross sectional area lof each flow passage measured along the shortest line from the leading edge of one blade to the upper surface of the next adjacent blade is less than the succeeding cross sectional areas to exclude a throat in each such passage. These proportions insure efficient operation of the blades in converting the high velocities in the neighborhood of sonic velocities to static pressure.

It will be clear that this invention provides the means of increasing the range of operation of axial ow compressors by providing stators which direct their ow lagainst the adjacent downstream rotor with very high velocity at a value about equal to sonic velocity in the uid approaching the rotor or at somewhat higher than sonic velocity. The flow leaving the stator, since it is directed against the direction of rotation of the rotor, is defined as making a positive angle with the axis of rotation.Y

The rotor passages between blades preferably diverge radially but this is not essential for utilizing the positive angle of ow from they stators to a substantial extent to augment the range of mass flow through the compressor. The range of mass flow and the increases in pressure are in part accomplished by having the positive direction of the ilow from the'stators, by the sonic or high speed nature of the ow, and by the relatively at angles which the rotor blades make with the plane of rotation. Also the blades are adapted by their fair or streamline contours and their close peripheral spacing to handle elastic uid at` velocities in the neighborhood of sonic values.

Thus the blades have relatively thin blade sections and they overlap in axial view.

The fore portions of the blades are set on the rotors at small pitch angles, that is angles between the blades and the plane of rotation. See Fig; 7 for instance where the fore portions of blades 94 are set parallel to vector 92 and are accordingly positioned more along a plane of rotation than along the axis of rotation, that is more nearly parallel to the rotation plane than to the rotation axis.

While I have illustrated a speciiic form of this invention it is to be understood that I do not intend to limit myself to thisl exact form but intend to claim my invention breadly as indicated by the appended claims.

I claim:

l. In combination in a compressor lfor pumping elastic liuid therethrough, a case, a plurality of rotor stages of blades positioned in said case for rotation about an axis with rotor flow passages between said blades, a plurality of stages of`stator blades alternated with said rotor stages `including a downstream stator stage positioned upstream adjacent to al downstream said rotor stage, said downstream stator stage receiving compressed uid for tlow f 7 therethrough `from said upstream stages during operating of said compressor, the blades of said downstream stator stage having a substantial chordwise curvature for directing said fluid flow with a substantial positive angle relative to the direction of said axis so that the direction of said fluid leaving said downstream stator stage has a limited range of angles relative to said blades of said rotor stage downstream adjacent to said downstream stator stage for varying rates of mass flow, said downstream stator stage directing said fluid toward said downstream rotor stage blades against their direction of rotation with a velocity relative to .any element of said downstream rotor stage blades equal to or greater than the velocity of sound in said fluid, each said downstream rotor stage passage increasing in cross sectional area rearward with the exit cross sectional area greater than the inlet cross sectional area, the cross sectional area of each said llow passage measured along the shortest line from the leading edge of one blade to the upper surface of the next adjacent blade being less than the succeeding cross sectional areas to exclude a throat in each such passage, said downstream rotor stage blades being constructed each with a relatively thin faired leading edge for high speed operation, each following blade of said downstream rotor stage having the normal projection of its leading edge on the convex surface of the adjacent leading blade closely adjacent the mid-chord point thereof, said downstream rotor stage blades substantially overlapping in axial View and each having blade sections each with its maximum thickness at about its mid-chord point and nearer thereto than to the point of said normal projection, each said downstream rotor stage blade having a substantially convex upper contour extending from substantially said leading edge over said maximum thickness ordinate, providing for effective operation with fluid at about or greater than said sonic velocity in said fluid approaching said downstream rotor stage, and means to rotate said downstream rotor stage at blade tip speeds relative to said fluid of sonic or greater than sonic speed in said fluid adjacent to said blades during the normal condition of operation of said compressor corresponding to fluid pressures at substantially maximum values thereof and at substantially the maximum normal rate of rotation of said downstream rotor stage.

2. In combination in an axial flow elastic fluid compressor for impelling a flow of elastic fluid therethrough with a pressure rise, a wall means defining a case, a plurality of rotor and stator stages positioned in said case in alternating relation and including an upstream rotor stage and a downstream rotor stage having a hub wall and axial flow blades mounted thereon peripherally spaced thereabout with rotor flow passages between said blades, said stator stages being supported in said case and including a downstream stator stage positioned upstream adjacent to said downstream rotor stage, said rotor stages supplying said fluid in a compressed state to said downstream rotor stage for flow therethrough during operation of said compressor, each said following blade of said downstream rotor stage having the normal projection of its leading edge on the convex surface of the adjacent leading blade at about the mid-chord point ofthe adjacent leading blade with the chords thereof angularly set more along a plane of rotor stage rotation than along said axis, said downstream rotor stage blades and the portions of the walls of said case and hub opposite said blades cooperating to provide cross sectional areas of said downstream rotor stage flow passages substantially increasing in the downstream direction from substantially the leading edges of said blades to substantially the trailing edges thereof to provide for efficient compression of said fluid at fluid speeds relative to said blades of the order of the speed of sound in said fluid at the inlets to said passages of said downstream rotor stage, and means to rotate said downstream rotor stage at blade tip speeds relative to said fluid of sonic or greater than sonic speed in saiduid adjacent to said blades during the normal condition of operation of said compressor corresponding to fluid pressures at substantially maximum values thereof and at substantially the maximum normal rate of rotation of said downstream rotor stage.

3, In combination in a multistage axial flow compressor for raising the pressure of an elastic fluid flowing therethrough, a case, a plurality of upstream stator stages positioned in said case having peripherally spaced stator blades with passages therebetween, a plurality of axial flow rotor stages positioned in said case for rotation about an axis, said rotor stages being alternated with said stator stages, said plurality of upstream rotor stages having blades with fore portions positioned at substantial angles with respect to said axis, said blades being peripherally spaced bounding rotor passages therebetween, said plurality of rotor stages also including a downstream axial flow rotor stage, said plurality of stator stages including a downstream axial flow stator stage positioned upstream adjacent to said downstream rotor stage, at least one of said upstream stator stages being formed to deflect said flow in the direction of rotation of a said upstream roter stage in operation, said downstream stator stage being formed to direct said flow less in the direction of rotation providing a fluid flow velocity relative to said downstream rotor stage of a magnitude of about or more than sonic value in said fluid between said downstream stator stage and said downstream rotor stage, each said downstream. rotor stage passage increasing in cross sectional area rearward with the exit cross sectional area greater than the inlet cross sectional area, said downstream rotor stage blades being designed each with a relatively thin faired leading edge for high speed operation, each following blade of said downstream rotor stage having the normal projection of its leading edge on the convex surface of the adjacent leading blade substantially at the mid-chord point thereof, said downstream rotor stage blades overlapping in axial view over a substantial portion of their span and each having blade sections each with its maximum thickness at about its mid-chord point and nearer thereto than to said normal projection, cach said blade of said downstream rotor stage having a substantially convex upper contour extending from said leading edge over said maximum thickness ordinate, said downstream rotor blades having thin blade sections of chordwise fair contours and having their aft ends directed more nearly parallel to said axis than to said fore portion of said downstream rotor blades, to be effective in converting said fluid velocity to static pressure, all said blades of said rotor having similar blade sections, each said blade section presenting said concave contours toward the direction of rotation for effecting compressive action, and means to rotate said downstream rotor stage at blade tip speeds relative to said fluid of sonic or greater than sonic speed in said fluid adjacent to said blades during the normal condition of operation of said compressor corresponding to fluid pressures at substantially maximum values thereofvand at substantially the maximum normal rate of rotation of said downstream rotor stage.

4. In combination in a multistage axial ow compressor for raising the pressure of an elastic fluid flowing therethrough, a case, a plurality of upstream stator stages positioned in said case having peripherally spaced stator blades with passages therebetween, a plurality of axial flow rotor stages positioned in said case for rotation about an axis, said rotor stages being alternated with said stator stages, said rotor stages having blades with fore portions positioned at substantial angles with respect to said axis, the last said blades being peripherally spaced with rotor passages therebetween, said plurality of rotor stages including a downstream axial flow rotor stage, said plurality of stator stages including a downstream axial flow stator stage positioned upstream adjacent to said downstream rotor stage, at least one of said upstream stator stages being formed to deflect said flow in the direction of rotation of a said upstream rotor stage in operation of said compressor, said downstream stator stage being formed to direct said ow less in the direction of rotation than the last said stator stage providing a fluid How velocity relative to said downstream rotor stage of a magnitude of about or more than sonic value in said fluid between said downstream stator stage and said downstream rotor stage, each said blade of said downstream rotor stage having a fore portion angularly positioned at a greater angle relative to said axis than said fore portions of said blades of an upstream said rotor stage, each said passage of said downstream rotor stage being bounded by a leading said blade and a following said blade, said downstream rotor stage blades overlapping in axial View by a substantial amount and each having its maximum thickness at about mid-chord point and nearer thereto than to the normal projection on the convex surface of the adjacent leading blade of the leading edge of said following blade to be effective in converting said fluid velocity to static pressure, and means to rotate said downstream rotor stage at blade tip speeds lrelative to said uid of sonic or greater than sonic speed` in said uid adjacent to said blades during the normal condition of operation of said compressor corresponding to fluid pressures at substantially maximum values thereof and atsubstantially the maximum normal rate of rotation yofsaid downstream rotor. stage.

5. In combination in an axial flow elastic fluid compressor for substantially increasing the` pressure and density of an elastic fluid, a case, va rotor positioned in said case for rotation about an axis to compress and impel said uid through said case, said rotor comprising ahub and a plurality of thin axial How blades carried thereon with rotor flow passages between said blades for ow of said uid therethrough, each said rotor passage increasing' in cross sectional area rearward tl'lerealong4 with the exit'cross sectional area greater than the inlet cross sectional area thereof, said blades being designed each with a relatively sharp trailing edge and a relatively thin faired leading edge foroperation at said high speeds, each following blade having thev normal projection of its leading edge on the convex surface of the adjacent leading blade at about the mid-chord point thereof to be effective in converting' velocity pressure to static pressureat tip speeds equal to sonic or greater therethan' in adjacentsaid uid, said blades having blade sections each with the maximum thickness ordinate thereof nearer to said mid-chord point than to said projection, the cross sectional 'area of each said tlow passage measured along the shortest line from the leading edge of one blade to the upper surface of the next adjacent blade being less than the succeeding cross sectional areas to exclude a throat lin each such passage, each said blade having asubstantially convexupper contour extending fro1n-said-leadir i g edge oversaidjmaximum thickness ordinate tobejefectiye ,throughout a subsonic range of speeds into a supersonic range at the tips of said blades, all said blades of` said rotor having similar blade sections, each said blade section presenting .said concave contoursitowardnthe directionfof rotation for effecting compressing action, and means to rotate said rotor at blade tip speeds relative to said fluid at sonic or greater than sonic speed in said uid adjacent said` blades during the normal optimum conditions of operation of said compressor corresponding to efciencies or pressures at substantially maximum values thereof and at substantially the maximum normal rate of rotation of said rotor.

6. in combination in an axial ow elastic uid compressor for substantially increasing the pressure and density of an elastic fluid, a case, a rotor positioned in 1G said case for rotation about an axis to compress and impel said fluid through said case, said rotor comprising a hub and a plurality of thin axial flow blades carried thereon with rotor flow passages between said blades for ilow of said fluid therethrough, each said rotor passage increasing in cross sectional area rearward therealong with the exit cross sectional area greater than the inlet cross sectional area thereof, said blades being designed each with a relatively sharp trailing edge and a relatively thin faired leading edge for operation at high speeds, each following blade having the normal projection of its leading edge on the convex surface of the adjacent leading blade at about the mid-chord point to be elective in converting velocity to static pressure at blade tip speeds equal to sonic or greater there than in adjacent said fluid, said blades substantially overlapping in axial view over a substantial portion of their span and each having blade sections each with the maximum thickness ordinate substantially at said mid-chord point, the cross sectional area of each said ow passage measured along the shortest line from the leading edge of one blade to the upper surface of the next adjacent blade being less than the succeeding cross sectional areas to exclude a throat in each such passage, each said blade having a substantially convex upper contour extending from said leading edge over said maximum thickness ordinate, to be effective throughout a subsonic range of speed into a supersonic range at the tips of said blades, all said blades of said rotor having similar blade sections, each said blade section presenting said concave contours toward the direction of rotation for effecting compressing action, and means to rotate said rotor at blade tip speeds relative to said lluid of sonic or greater than sonic speed in said fluid adjacent to said blades during the normal condition of operation of said compressor corresponding to fluid pressures at substantially maximum values thereof and at substantially the maximum normal rate of ,rotation of said rotor.

vtherethrough fromy said upstream stages during operation of said compressor, said downstream stator stage having stator blades of substantial chordwise curvature for directing said fluid llow with a substantial positive `angle relative to said axis so that the direction of said `fluid leaving said downstream stator stage has a limited range of angles relative to the blades of a said rotor stage downstream adjacent to said downstream stator stage, said downstream stator directing said ilow toward said downstream rotor stage blades with a velocity relativeto any element thereof equal to or greater than the velocity ofsound in said luid, each said blade of said l'downstream rotortstag'e having smaller maximum thickness atvthe root end thereof than the maximum thicknesses at the root ends of the blades of an upstream said rotor stage vto be effective at higher relative ilow speeds, said thicknesses being expressed as fractions of the blade chord lengths, said downstream rotor stage blades overlapping in axial view over a substantial portion of their span for eective operation with fluid at about or greater than said sonic velocity in said lluid approaching said rotor, and means to rotate said downstream rotor stage at blade tip speeds relative to said fluid at about sonic or greater than sonic speed in said fluid adjacent to said blades during the normal optimum condition of operation of said compressor corresponding to efficiency or fluid pressure at substantially maximum values thereof and at substantially the maximum normal rate of rotation of said downstream rotor stage.

8. ln combination in an axial llow elastic lluid compressor for substantially increasing the pressure and density of an elastic iluid, a case, a rotor positioned in said case for rotation about an axis to compress and impel said fluid through said case, said rotor comprising a hub and a plurality of thin axial flow blades carried thereon with rotor llow passages between said blades for flow of said lluid therethrough, each said rotor passage increasingy in cross sectional area rearward therealong with the exit cross sectional area greater than the inlet cross sectional area thereof, the cross sectional area of each said ilow passage measured along the shortest line from the leading edge of one blade to the upper surface of the next adjacent blade being less than the succeeding cross sectional areas to exclude a throat in each such passage, said blades being designed each with a relatively sharp trailing edge and a relatively thin faired leading edge for operation at high speeds, said blades each having blade sections each with the maximum thick# ness ordinate substantially at said mid-chord point, each said blade having a substantially convex upper contour extending from said leading edge over said maximum thickness ordinate and a concave lower contour, to be etlective throughout a subsonic range of speed into a supersonic range at the tips of said blades, all said blades of said rotor having similar blade sections, each said blade section presenting said concave contours toward the direction of rotation for effecting compressing action, and means to rotate said rotor at blade tip speeds relative to said iluid of sonic or greater than sonic speed in said fluid adjacent to said blades during the normal condition of operation of said compressor corresponding to fluid pressures at substantially maximum values thereof and at substantially the maximum normal rate of rotation of said rotor.

9. In combination in an axial llow elastic fluid compressor for substantially increasing the pressure and density of an elastic lluid, a case, a rotor mounted in said case for rotation about an axis to compress and impel said .tluid through said case, said rotor comprising a hub and a plurality of thin axial ilow blades carried thereon with rotor ilow passages between said blades for flow of said lluid therethrough, each said rotor passage increasing in cross sectional area rearward therealong with the exit cross sectional area greater than the inlet cross sectional area thereof, said blades being designed each with a relatively sharp trailing edge and a relatively thin faired leading edge for operation at said high speeds, each following blade having the normal projection of its leading L edge on the convex surface of the adjacent leading blade at about the mid-chord point thereof to be elective in converting velocity pressure to static pressure at tip speeds equal to sonic or greater than sonic in the adjacent fluid, said blades each having blade sections each with a subsonic range of speeds into a supersonic range at the tips of said blades, all said blades of said rotor having similar blade sections, each said blade section presenting said concave contours toward the direction of rotation for ellecting compressing action, and means to rotate said rotor at blade tip speeds relative to said iluid of sonic or greater than sonic speed in said lluid adjacent to said blades during the normal condition of operation of said compressor corresponding to fluid pressures at substantially maximum values thereof and at substantially the maximum normal rate of rotation of said rotor.

10. ln combination in an axial flow elastic fluid compressor for impelling a flow of elastic fluid therethrough with a pressure rise, a wall means dening a case, a plurality of rotor and stator stages positioned in said case in alternating relation and including an upstream rotor stage and a downstream rotor stage having a hub wall and axial flow blades mounted thereon peripherally spaced thereabout with rotor ilow passages between said blades, said stator stages being supported in said case and including a downstream stator stage positioned upstream adjacent to said downstream rotor stage, said rotor stages supplying said lluid in a compressed state to said downstream rotor stage for flow therethrough during operation of said compressor, said blades of said downstream rotor stage having their maximum thicknesses substantially at the mid-chord points thereof and having their leading edges transversely across the adjacent said rotor passages from about the mid-chord points of the adjacent leading blades with the chords thereof angularly set more along a plane of rotor stage rotation than along said axis, said downstream rotor stage blades and the portions of the walls of said case and hub opposite said blades cooperating to provide cross sectional areas of said downstream rotor stage llow passages substantially increasing in the downstream direction from substantially the leading edges of said blades to substantially the trailing edges thereof to provide for ellicient compression of said fluid at lluid speeds relative to said blades of the order of the speed of sound in said lluid at the inlets to said passages of said downstream rotor stage, the cross sectional area of each said llow passage measured along the shortest line from the leading edge of one blade to the upper surface of the next adjacent blade being less than the succeeding cross sectional areas to exclude a throat in each such passage, and means to rotate said downstream rotor stage at blade tip speeds relative to said lluid of sonic or greater than sonic speed in said lluid adjacent to said blades'during the normal condition of operation of said compressor corresponding to fluid pressures at substantially maximum values thereof and at substantially the maximum normal rate of rotation of said downstream rotor stage.

References Cited in the file of this patent UNITED STATES PATENTS 2,326,072 Seippel Aug. 3, 1943 2,378,372 Whittle June 12, 1945 2,406,126 Zweifer Aug. 20, 1946 2,527,971 Stalker Oct. 31, 1950 FOREIGN PATENTS 386,039 Great Britain Ian. 12, 1933

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US2958456A (en) * 1954-10-06 1960-11-01 Power Jets Res & Dev Ltd Multi-stage aerofoil-bladed compressors
US2990106A (en) * 1956-10-12 1961-06-27 English Electric Co Ltd Axial flow multi-stage compressors
US3009630A (en) * 1957-05-10 1961-11-21 Konink Maschinenfabriek Gebr S Axial flow fans
US3032313A (en) * 1956-04-09 1962-05-01 Bertin & Cie Turbo-machines
US3751909A (en) * 1970-08-27 1973-08-14 Motoren Turbinen Union Turbojet aero engines having means for engine component cooling and compressor control
US6260349B1 (en) 2000-03-17 2001-07-17 Kenneth F. Griffiths Multi-stage turbo-machines with specific blade dimension ratios
US6378287B2 (en) 2000-03-17 2002-04-30 Kenneth F. Griffiths Multi-stage turbomachine and design method
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US20070295011A1 (en) * 2004-12-01 2007-12-27 United Technologies Corporation Regenerative Turbine Blade and Vane Cooling for a Tip Turbine Engine
US20080014078A1 (en) * 2004-12-01 2008-01-17 Suciu Gabriel L Ejector Cooling of Outer Case for Tip Turbine Engine
US20080093174A1 (en) * 2004-12-01 2008-04-24 Suciu Gabriel L Tip Turbine Engine with a Heat Exchanger
US20090071162A1 (en) * 2004-12-01 2009-03-19 Suciu Gabriel L Peripheral combustor for tip turbine engine
US20090142184A1 (en) * 2004-12-01 2009-06-04 Roberge Gary D Vectoring transition duct for turbine engine
US20090145136A1 (en) * 2004-12-01 2009-06-11 Norris James W Tip turbine engine with multiple fan and turbine stages
US20090155079A1 (en) * 2004-12-01 2009-06-18 Suciu Gabriel L Stacked annular components for turbine engines
US20090232650A1 (en) * 2004-12-01 2009-09-17 Gabriel Suciu Tip turbine engine and corresponding operating method
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US2958456A (en) * 1954-10-06 1960-11-01 Power Jets Res & Dev Ltd Multi-stage aerofoil-bladed compressors
US3032313A (en) * 1956-04-09 1962-05-01 Bertin & Cie Turbo-machines
US2990106A (en) * 1956-10-12 1961-06-27 English Electric Co Ltd Axial flow multi-stage compressors
US3009630A (en) * 1957-05-10 1961-11-21 Konink Maschinenfabriek Gebr S Axial flow fans
US3751909A (en) * 1970-08-27 1973-08-14 Motoren Turbinen Union Turbojet aero engines having means for engine component cooling and compressor control
US6260349B1 (en) 2000-03-17 2001-07-17 Kenneth F. Griffiths Multi-stage turbo-machines with specific blade dimension ratios
US6378287B2 (en) 2000-03-17 2002-04-30 Kenneth F. Griffiths Multi-stage turbomachine and design method
US20070295011A1 (en) * 2004-12-01 2007-12-27 United Technologies Corporation Regenerative Turbine Blade and Vane Cooling for a Tip Turbine Engine
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US20080014078A1 (en) * 2004-12-01 2008-01-17 Suciu Gabriel L Ejector Cooling of Outer Case for Tip Turbine Engine
US20080093174A1 (en) * 2004-12-01 2008-04-24 Suciu Gabriel L Tip Turbine Engine with a Heat Exchanger
US20090071162A1 (en) * 2004-12-01 2009-03-19 Suciu Gabriel L Peripheral combustor for tip turbine engine
US20090142184A1 (en) * 2004-12-01 2009-06-04 Roberge Gary D Vectoring transition duct for turbine engine
US20090148273A1 (en) * 2004-12-01 2009-06-11 Suciu Gabriel L Compressor inlet guide vane for tip turbine engine and corresponding control method
US20090145136A1 (en) * 2004-12-01 2009-06-11 Norris James W Tip turbine engine with multiple fan and turbine stages
US20090155079A1 (en) * 2004-12-01 2009-06-18 Suciu Gabriel L Stacked annular components for turbine engines
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