US3252212A

US3252212A - Method of selectively matching a turbine wheel and turbine nozzle assembly

Info

Publication number: US3252212A
Application number: US297538A
Authority: US
Inventors: Iii Albert H Bell
Original assignee: Chrysler Corp
Current assignee: Old Carco LLC
Priority date: 1963-07-25
Filing date: 1963-07-25
Publication date: 1966-05-24
Anticipated expiration: 1983-05-24
Also published as: US3319931A; GB1057234A

Description

May 24, 1966 A. H. BELL In METHOD OF SELECTIVELY MATCHING A TURBINE WHEEL AND TURBINE NOZZLE ASSEMBLY 4 Sheets-Sheet l Filed July 25, 1963 3,252,212 EEL A. H. BELL Ill May 24, 1966 METHOD OF SELECTIVELY MATCHING A TURBINE WH AND TURBINE NOZZLE ASSEMBLY 4 Sheets-Sheet 2 Filed July 25 1963 www;

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May 24, 1966 A. H. BELL nl 3,252,212

METHOD OF' SELECTIVELY MATCHING A TURBINE WHEEL AND TURBINE NOZZLE ASSEMBLY Filed July 25, 1965 4 Sheets-Sheet I5 A R IN VENTOR.

A Mda/Ya. HZ H zzf @EL fw/m5.

May 24, 1966 H. BELL nl 3,252,212

METHOD OF' SELEC ELY MATCHING A TURBINE WHEEL AND TURBINE NOZZLE ASSEMBLY Filed July 25, 1963 4 Sheets-Sheet 4 IN VEN TOR.

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United States Patent O M 3,252,212 METHOD F SELECTIVELY MATCHING A TUR- BINE WHEEL AND TURBINE NOZZLE ASSEM- BLY Albert H. Bell III, Birmingham, Mich., assignor to Chrysler Corporation, Highland Park, Mich., a corporation of Delaware Filed July 2S, 1963, Ser. No. 297,538 9 Claims. (Cl. 29-407) This invention relates generally to turbine engines and more particularly to the fixed nozzle vanes and wheel blades employed therein.

As Well known in the art, turbine engines are comprised Of a gas generator section consisting primarily of an air compressor, burner portion and a compressor turbine wheel which obtains its energy from the gases fiowing out of the burner portion and drives the compressor. Some turbine engines, depending on their intended use, also include a power turbine which when placed in the path of the flowing gases, downstream of the compressor turbine, serves to provide additional work such as by driving a propeller in a turbojet engine or driving some other output shaft for land or water-based vehicles.

The ultimate output power developed by any turbine engine is primarily dependent on the output of the gas generator section which, in turn, is a function of the compressor speed of that particular gas generator.

In many instances it has been found that substantial variations in output horsepower exist as between any two turbine engines even though the turbine engines are of the same design and rated output power. Such variations, which in some cases have been in the magnitude of thirty percent (30%) power loss, exhibit themselves to the greatest extent in the range of compressor speeds of seventy-five percent (75%) to a hundred percent (100%) of the designed or rated maximum compressor speed. Further, these variations occur most frequently in turbine engines wherein the gas fiow through the fixed guide vanes or nozzles and compressor turbine Wheel blades closely approaches sonic velocity as the compressor speed approaches designed maximum speed.

As a consequence of such power variations, it becomes impossible to predict, with any reasonable degree of accuracy, the expected power output of any particular turbine engine. Further, since such variations have been found to exist between engines of identical design, it must be concluded that the engine delivering the lesser power is, at least to some degree, of lesser efficiency and therefore undesirable.

Now it has been discovered that if certain relationships between the elements comprising the 4turbine engine are considered as being critical and such relationships are established during the process of manufacturing the turbine engine, that the resulting engines will produce a reasonably predictable output horsepower which is consistently at a relatively higher and more eliicient value.

Accordingly, a general object of this invention is to provide, in a turbine engine, iixed guide vanes or nozzles and cooperating compressor turbine wheel blades having a relationship established therebetween which enables the turbine engine to exhibit an overall higher efficiency.

Another object of this invention is to provide in a turbine engine, fixed guide vanes or nozzles and cooperating wheel baldes having a relationship therebetween which enables the turbine engine to achieve a predicted output horsepower which is within a reasonable range of tolerances.

Still another object of this invention is to provide a method for achieving a particular critical relationship between fixed guide Vanes or nozzles and cooperating tur- 3,252,212 Patented May 24, 1966 ICC bine wheel blades of a turbine engine so as to obtain a turbine engine having a predictable horsepower output. Another object of this invention is to provide a method for achieving the above said critical relationship which method includes the selective matching of a turbine wheel and cooperating turbine Wheel nozzle assembly.

A further object of this invention is to achieve the above said critical relationship by a method of selective matching of turbine Wheel and cooperating turbine wheel nozzle assembly which method includes matching based on the effective iioW areas of said turbine wheel and said turbine wheel nozzle assembly.

The invention as hereinafter disclosed in detail, contemplates a method of selectively matching a turbine wheel to a cooperating turbine nozzle assembly which comprises the steps of creating a first iiow of a suitable fiuid through the nozzle assembly, determining from said flow of fiuid therethrough the effective flow area of said nozzle assembly, repeating the above steps with succeeding nozzle assemblies, classifying said nozzle assemblies according to their respective equivalent flow areas, creating a second iiow of a suitable fluid through said turbine Wheel, determining from said second flow of iiuid through said wheel the equivalent iiow area of said turbine wheel, repeating the steps with succeeding turbine wheels, classifying each of the turbine wheels according to their respective equivalent flow areas, and selectively matching a classified nozzle assembly to a classified turbine wheel according to a predetermined relationship of their respective equivalent flow areas.

Other objects and advantages of the invention will become apparent when reference is made to the following description and drawings wherein:

FIGURE l is a cross-sectional view of a gas turbine engine adapted to propel driving wheels of a land based vehicle;

FIGURE 2 is a graph illustrating typical compressor curves obtained by plotting the ratio of compressor discharge pressure to compressor inlet pressure against the weight-rate of air iiow through the compressor;

FIGURE 3 is a graph illustrating a typical output horsepower curve of a turbine engine;

FIGURE 4 is an enlarged fragmentary portion of the turbine engine of FIGURE 1;

FIGURE 5 is a cross-sectional View taken generally on the plane of line 5-5 of FIGURE 4, looking in the direction of the arrows;

FIGURE 6 is a cross-sectional view taken generally on the plane of line 6 6 of FIGURE 5, looking in the direction of the arrows;

FIGURE 7 is a cross-sectional view taken generally on the plane of line 7--7 of FIGURE 5;

FIGURE 8 illustrates a generally tubular conduit of converging cross-sectional area;

FIGURE 9 is a graph illustrating the relationship between the change in velocity of iiiow through the conduit of FIGURE 8 for a corresponding change in cross-sectional area; and

FIGURES l0 and ll illustrate, in cross-section, an arrangement for testing nozzle assemblies and turbine wheels, respectively, in accordance with the teachings ott this invention.

Referring now in greater det-ail t0 the drawings, FIG- URE l illustrates a turbine engine 10 adapted vlfor piropelling a land-based vehicle having driving wheels 12. The engine 10 is comprised of a housing 14 having an air intake 16 and exhaust orifice 18. A combustion chamber 20, having any suitable fuel distribution means such las a fuel distribution ring 22 therein, is located within the housing 14 between the compressor 24 and compressor turbine wheel 26. Compressor 24 and iuzrbine wheel 26 are connected to each other by means of a shaft 9 28 Iwhich may also be used for driving or in other ways operating selected engine accessories such as the fuel control schematically illustrated at 30.

Compressor 24, as illustrated, is of the radial flow type and is comprised of :a generally circular body portion 32 and cylindrical bearing portion 34 with a plurality of generally radially directed compressor vanes 36 therebetween.

Turbine wheel 26 is provided with a plurality olf circumferentially spaced blades 38 extending radially outwardly from the wheel rim 40. Upstream of turbine wheel 26, the engine 10 is provided with a plurality of radially directed guide vanes lor nozzles 42, sometimes referred to as stators, spaced oircumferentially about the body portion 44.

As compressor 24 is rotated, air is drawn in through inlet 16, compressed by vanes 36 and directed into the combustion chamber 20 where fuel, supplied in accordance with a predetermined schedule, is burned so as to heat the gases therein. The gases flow from the burner chamber 20v through nozzles 42 which accelerate and caluse the gases to impinge upon the turbine blades 38 in a manner causing rotative motion of turbine wheel 26r which, in turn, drives compressor 24.

The gases, continuing to flow towards exhaust orifice 18, pass between a second set of guide vanes or nozzles 46, similar to nozzles 42, which direct the gases against blades 48 Iof an loutput power turbine wheel 50. Turbine wheel S may be connected, a-s by any suitable power transmission means 52, 54 to the driving wheels 12.

FIGURE 2 is a graph illustrating

typical compressor curves

56, 58, 60, 62, 64, 66, 68 and 70 obtained by plotting the ratio of compressor discharge pressure, P2, to,

inlet pressure, P1, against the air flow through the compressor in pounds per second. Each of the curves can be obtained by maintaining the compressor at selected constant speeds While varying the pressure ratio P2/P1. The respective compressor speeds are indicated, for illustrative purposes, near each curve in terms lof percentages of designed maximum compressor speed.

The area above and to the left of dash line 72 represents that area in which compressor instability or surge is. encountered. Compressor instability or surge refersy to a condition sometimes referred to as hunting, that is, those portions of the compressor curves which tend to flatten out and become more horizontal indicate that slight variations in the ratio of P2/P1 result in comparatively large changes in the air flow which give rise to compressor instabil-ity. Curve '72 is determined generally by connecting together the respective points, on the individual compressor curves, at which compressor surge is first encountered.

In addition to the compressor surge area, as discussed above, a further important consideration remains. That is, the rapidreduction :of the numerical value of the ratio of P2/P1 as illustrated by the generally vertically depending portions of each of the compressor curves 56 through 70.A Since the compressor inlet pressure, P1, may be assumed to be constant, the conclusion must be that the downstream or compressor discharge pressure, P2, decreases and the rate of .P2 decrease, as illustrated yby the slope of the depending portions lorf the compressor curves, is very rapidas compared to an increment Iof change in the air flow. The reduction in comp-ressor discharge pressure, P2, is cau-sed by the choking effect of the air within the compressor. Consequently, a line 74 drawn generally through the points of each Iof the compressor curves wherein the rate of change of the ratio PZ/ P1 starts to become rapid defines .an a-rea, generally -below and to the right of line 74, which miglht be referred to as the compressor choke area.

Accordingly, with turbine engines,-precaution1s are taken to assure the operation of the compressor within the limits defined by the surge lor stall line 72 and choke line 74. In order to achieve this, it has been the practice,

generally, to aerodynamically match the compressor turbine wheel and its associated nozzle assembly, as a subassembly, to the compressor. This is sometimes referred to as gross matching. In some instances other limits similar to the choke line 74 and the stall line 72 are determined and employed for aerodynamic reasons. The aircraft industry has also employed additional means for avoiding at least portions of the compressor stal-l or surge areas by providing compressor bleed valves which at times and in accordance with selected operating parameters, vent some of the compressor air to the atmosphere.) A turbine engine having such a gross-matched compressor, lcompressor turbine wheel and nozzle assembly, so as to operate between the stall line 72 and choke line 74, should exhibit a characteristic horsepower curve 76, obtained generally Aby plotting output horsepower `and the percent of designed maximum compressor speed along a logarithmic Y-axis and an arithmetic X-axis, respectively, of the graph of FIGURE 3.

However, it ha-s been found that substantial unpredictable variations in output horsepower still exist between gas turbine engines which have been constructed in accordance with the prior art.

v For example, with reference to FIGURE 3, i-t has been found that while one turb-ine engine might deliver full rated power, another engine of the sarne design and rating might well develop substantially lessy horsepower than that predicted. That is, a particular engine might deviate from the predicted mean curve 76, as at some point 80, and after reaching a maximum output as at .point 82 (substantially lower than the rated value represented by point 7S) steadily decrease to some point 84 which may represent a developed horsepower less than that `of point 80.

Further, it V has been foundv that this phenomenon of horsepower variation, or lost horsepower, will at times occur in the same gas` turbine engine which hadoriginally developed the full rated horsepower. stated, the phenomenon of unpredictable horsepower variation occurs most frequently in gas turbine engines wherein the velocity of gas flow I'through the compressor turbine and the compressor turbine nozzleassembly is'in the transonic rangeas the compresso-r approaches its designed maximum speed.

In the past, various proposals Ihave been tried inl unsuccessful attempts to overcome theproblem of horse-- power variation. For example, the redesigning, modification and/or inspection of such areas as lubrication systems, gear trains, horsepower demands ofengine` driven accessories, alignment of turbinecomponents and engine air leakages in no Way alleviated; the problem ofy horsepower variation.

Therefore, assuming point-84 to;be the lowest horsepower ever achieved, it can be seen that turbine engines, constructed in accordance with Vthe prior art and having the compressor turbine and nozzle assembly grossly matched, as a sulbassembly, to thecompressor, can be expected lto produce an output horsepower value ranging anywhere between the limits defined by points 78l and 84. In view of this, it becomes evident that the gross aerodynamic matching of the compressor turbine wheel and its nozzle assembly, as a subassembly, to the compressor is in itself insufficient for assuring reasonable attainment of predicted horsepower and that still another influencing factor, heretofore not apparent to those skilledin the art,

remains as an important consideration in the design and` As was previouslyy sor turbine wheel 26. Preferably, the nozzle or stator vanes 42 are made integrally with a main body p-or-tion 86 which is secured to a portion 44 of the general engine housing 14. This may be accomplished as by means of a nut 88 internally threaded so as to cooperate with a threaded portion 90 in axially urging the stator body 86 against the radial surface 92 of a mounting shoulder 94. Radially outwardly, an annular shroud 96 is formed preferably integrally with the nozzle vanes 42. The shroud 96, in cooperation with the outer surface 98 of the stator body 86, defines an annular passage 97 in which the varies 42 are located yfor the directional control ofthe gases flowing therethrough.

The 'compressor turbine Wheel 26, secured to shaft 28 for rotation therewith, has its blades 38 generally Within the confines of shroud 96. The clearance between the outermost ends 100 of moving turbines blades 38 and the inner stationary surface 102 of shroud 96 is kept to a minimum. The outer surface 40 of turbine wheel 26 and the inner surface 102 of -shroud 96 also define an annular passage, generally coaxial with passage 97, for the ow of gasses therethrough.

The flow area as between any two adjacent nozzle vanes 42 Will be of the smallest cross-sectional area between such adjacent blades. For example, assuming that a plane indicated by line 6--6 of FIGURE 5 were passed through vanes 42 so as to have the cross-sectional area between the vanes a minimum area, the lflow area 103 therebetween could be determined, generally, from the dimension-s H1, W1 and B1 as illustrated -in FIGURE 6. Similarly, the ow area of the turbine blades 38 could be determined by passing a plane indicated by line 7 7 of FIGURE 5 through points resulting in a minimum area 104. The flow area between such turbine blades would then be determined, generally, by the dimensions H2, W2 and B2 as illustrated in FIGURE 7. (Because of the relatively small clearances existing between the ends 100 of turbine blades 38 and the inner surface 102 of stationary shroud 96, the height, H2, in FIGURE 7 can -be considered as the distance fro-m surface 40 of turbine Wheel 26 to the inner surface 102 of the shroud 96.)

As previously stated, the gross matching of the nozzle assembly 43 and compressor turbine wheel, as a subassembly, to the compressor, even though well known -in the art, only `assures the engine of operation within the limits defined by the stall line 72 and the choke line 74 of FIGURE 2. Such gross matchng does not, however, give any assurance that the `engine will attain the maximum horsepower as represented by poi-nt 78 of FIGURE 3, but rather that the 'horsepower attained will be some- Where |between the desired point 78 and the low limit of point 84- It has been discovered Ithat what is still further required, in addition to the gross matching with reference to the compressor, :is a critical or selective matching of the nozzle assembly 43 to the compressor turbine wheel assembly 26 based on the respective effective iiow areas of each.

' Referring to FIGURES 4, 5 and 6 it can be seen that if plane 6-6 determines the smallest cross-sectional area between two adjacent nozzle vanes 42 then the cross-sectional areas taken at local points upstream of the plane 6 6 will define progressively larger areas generally as the distance between the local area under consideration and the minimum 4area 103 increases. In other words, the progressively larger areas can be considered as defining, generally, the converging conduit 120 of FIGURE 8l wherein the throat area, AT, would be a dimensional equivalent of the minimum area 103 defined by W1, B1 and H1 of FIGURE 6 and the various local areas, Am, AL2, ALS, etc., would constitute the increasing cross-sectional areas upstream of the minimum area 103. Further, just as the throat area, AT, will determine the maximum volume rate of air fiow through conduit 120, so will the minimum area 103 between nozzle vanes 42 determine the maximum volume rate of gas iiow therethrough for any given condition. Generally, it can be said that the absolute maximum rates in both arrangement will occur when the velocity of lair ow through the most restricted areas is sonic.

Assuming that conduit is under a condition of sonic flow, that is, the velocity of air fiow through throat 122 is Mach-one, a relationship can be established -between the respective local areas and the velocity of flow through such areas. For example, an exponential curve, 124 illustrated in FIGURE 9 can be obtained by graphically plotting the ratio of the local cross-sectional areas, AL, to the throat area AT, against the velocity of flow at that particular local area.

From an inspection of curve 124, it can be seen that when the ratio of areas, AL/AT, is 1.000 then the velocity of the flow Iat the local 4area is equal to Mach-one. It should also be observed, however, that the change in local air velocity, Mach No., is large as -compared to a relatively small change in the local cross-sectional area, as refiected by AR, in that range Where the ratio of areas, AL/AT, approaches 1.000.

Even though curve 124 of FIGURE 9 has been discussed primarily with reference to the converging conduit 120 of FIGURE 8, the principles involved apply equally well to the nozzle assembly 43. That is, -as previously stated, the minimum area 103 determined by W1, B1 and H1 is the equivalent `of the throat area, AT, of throat 122, and the various local cross-Sectional areas upstream of the minimum nozzle Iarea 103v are the equivalents of the local cross-sectional are-as Am, AL2, ALs, etc.

Accordingly, it becomes evident that if the size of the minimum nozzle area 103 is calculated to transmit gas therethrough at sonic velocities and at Ia specific weightrate of flow, that very slight deviations from the calculated size `of the minimum nozzle yarea as illustrated by FIGURE 9 will result in relatively large reductions in the velocity of air or gas fiow through the minimum nozzle area 103. The immediate effect of such velocity losses is a drastic reduction in the energy available for driving the compressor turbine Wheel 26 and ultimately the power turbine 50.

It has been discovered that in some instances turbine engines, having fa nozzle assembly wherein the blade surfaces 108 and 110 (FIGURE 5) were .003 inch further away from each other than the ideal calculated dimension, experienced about a ten percent (10%) power loss. It has also been -found that such turbine engines could be made to produce their full rated horsepower if the nozzle assembly and compressor turbine wheel 26 in the engine ywere replaced by a nozzle assembly and compressor turbine wheel which were selectively matched to each other.

For example, let it be assumed that it has been empirically determined that a particular engine requires the effective fiow area of the nozzle assembly and the effective iiow :area of the compressor turbine wheel to be in-an ideal ratio of 1.000`to 1.500, respectively, as determined at a particular pressure differential across each. (The effective or equivalent flow areas are determined by total flow and velocity experienced through the entire nozzle assembly 43, that is the summation of the effective fiow areas between the individual nozzle vanes, or the entire compressor turbine wheel, as `the case may be.)

It is well known that dimensional tolerances are required in any manufacturing operation. Therefore, it becomes impossible to randomly choose any one nozzle yassembly out of many and similarly choose a compressor turbine wheel and have lany degree of assurance that the two will have their respective equivalent flow areas in precisely the assumed ideal ratio of 1.000; 1.500. Therefore, the very tolerances necessary for the production of the nozzle assembly and compressor turbine wheel cause the respective equivalent flow areas to vary to the degree sufficient to cause ow velocity losses as described with reference to FIGURES 4-9. Such variations resulting from the manufacturing tolerances also exhibit themselves in causing the respective equivalent ow areas to vary from the assumed ideal ratio of 1.000; 1.500.

However, the discovery of the cause of .the problem of variations in `and loss of engine power, does not in and of itself suggest a necessarily practical solution of that problem because dimension-al manufacturing tolerances cannot be eliminated nor can they be reduced to a degree which results in prohibitive manufacturing costs.

It has been discovered that the solution of the problem resides in the selective matching of a nozzle assembly to a cooperating compressor turbine wheel by means of matching the ow performance of each of them.

For purposes of illustration, let the following be assumed to be the ideal constants:

(l) Ratio of the total nozzle assembly ow area to thev ytotal compressor turbine wheel ow area=1.000/ 1.500;

(2) Total nozzle assembly ilow area=10.000 sq. inches;

(3) Total compressor turbine wheel flow area: 15.000 sq.

inches;

(4) Velocity of gas ow through nozzle yassembly at full rated engine power=Mach 0.94;

(5 Velocity of gas ow through compressor turbine wheel at full rated engine power=Mach one.

Further, let it be assumed that because of the small dimensional differences obtained due to reasonable manufacturing tolerances, that relatively large changes occur (as previously discussed with reference to FIGURES 8 and 9) in the equivalent flow areas of both the nozzle assembly and the compressor turbine, so as to result in the following:

(6) The effective or equivalent total nozzle assembly ow area varies, as between assemblies, from 9.700 to 10.300 sq. inches and (7) The effective or equivalent total compressor turbine ow area varies, as between turbines from 14.500 to 15.500 sq. inches.

With the above assumptions it can be seen that a nozzle assembly and compressor turbine wheel randomly selected from a large group of each could yield the following equivalent flow area combinations having the corresponding indicated equivalent flow area ratios, which would still satisfy the gross mat-ching requirements asA between the compressor and the nozzle assembly and compressor turbine as a subassembly:

Combinations I and IV result in equivalent 110W areas which are very near the assumed ideal area ratios.

It has been discovered that, in gas turbine engines wherein the velocity of motive gas ow through the nozzle .assembly or through the compressor turbine is in the range of Mach 0.9 to Mach 1.() as the compressor approaches its designed maximum speed, the numerical equivalent of the actual area ratio may Vary in the range of plus or minus 1.0% from the numerical equivalent of the ideal area ratio. Therefore, combinations I and IV would result in the engine devoloping its full rated horsepower at point 78 of the Ihorsepower curve of FIGURE 3.

However, the same nozzle assemblies combined in re- Verse order with the compressor turbine wheels, as illustrated by combinations II and III, result in equivalent flow area ratios having a substantial deviation from the assumed ideal equivalent flow area ratio. Combinations II and III, even though satisfying the gross matching requirements of the compressor, would, nevertheless, cause the engine to produce an output horsepowervmuch less than the rated output of the engine such as represented by point 84 of the curve of FIGURE 3.

Accordingly, in order to insure proper engine perform- Iance so as to consistently achieve a reasonably predictable horsepower output it becomes necessary to selectively match the nozzle .assembly and compressor turbine to each other.

A method of selectively or critically matching the nozzle assembly and compressor turbine is disclosed with reference to the apparatus of FIGURE 10 which is comprised of an outer housing 126 provided with a cover 128 at one end thereof and 'an inlet passage 130 at the other end. Conduit 132, having a valve 134 serially connected therewith, communicates at its one end with a source 135 of pressurized air while its other end serves to retain an orice plate 136 in proper position so as to maintain the orifice 138 within the path of the air flowing from conduit 132 and into the plenum chamber 140 defined generally by housing 126 and cover 128.

Conduits

142 and 144 are provided so as to communicate the pressures on opposite sides of the orifice plate 136 to a differential pressure gage 146.

Conduits

142, 144, orifice 138 and pressure gage 146 provide means for determing the mass rate of flow through the orice 138 and therefore the entire system. The mass rate of W, W, can be determined by the application of the following general compressible mass flow equation:

Where:

(1) G=flow per unit of area;

(2) W=mass rate :of flow;

(3) A=Crosssectional flow :area of orice 138;

(4) C=the discharge coefficient of orifice 138 (a calibrated function of the Reynolds number as well known in the art;)

(5) P0=P2=total pressure upstream of orice 138;

(6) M=molecular weight of the fluid flowing;

(7) Tu=total temperature upstream of orifice 133;

(8) g=acceleration due to gravity;

(9) n.=the ratio of the specific heat of the uid at constant pressure, to the specific heat of the uid at constant volume;

(10) =universal gas constant; and

(11) 11=pressure ratio of downstream static pressure,

`Ps2 to upstream total pressure, Po A testing fixture 152 having an annular passage 154 formed therein is :secured insealing engagement with cover 128. A plurality of guide vanes 156 are provided within the annular passage 154 so as to impart a swirling motion to the air passing therethrough in order to simulate the langle ofair impingement experienced by the nozzle vanes in the actual engine. The nozzle assembly 43 to be tested is mounted, as illustrated at the exit of the annular passage 154.

During testing ofthe nozzle assembly 43 valve 134 is opened to the degree necessary to establish a desired differential, AP, between the plenum chamber pressure, P3, and the barometric pressure, PB. This pressure differential and effective cross-sectional ow area of the nozzle assembly will then determine the velocity of ow through the nozzle assembly.

Since nozzle assembly 43 and orifice 138 are in series withv each other, it becomes evident that the mass-rate of fiow, W, must be the same for both the nozzle assembly 43 and the orifice 138. Accordingly, once the mass-rate of flow, W, is determined for orifice 138, the same massrate of flow can be employed in the general compressible mass ow equation as applied to the nozzle assembly. The terms of the equation, as applied to the nozzle assembly will have the following meanings:

(l) A=actual cross-sectional flow area of nozzle assembly 43;

(2) Wzmass-rate of flow through the nozzle assembly (the same as the mass-rate of flow through orifice 138);

(3) P0=P3=pressure upstream of the nozzle assembly;

(4) m=molecular weight of the fluid;

(5) T0=total temperature upstream of the nozzle assembly;

(6) C=the discharge coefficient of the nozzle assembly;

(7) g=acceleration due to gravity;

(8) n,=the ratio of the specific heat of the gas at constant pressure, Cp, to the specific heat of the gas at constant volume, Cv;

(9) R=universal gas constant; and (l0) r=pressure ratio of downstream pressure, PB, to

upstream pressure, P3.

In applying the general equation to the nozzle assembly, the equation may be rearranged into the following form:

It should also be noted that W, m, To, g, n and R are the sarne values as when the general equation was applied to the orifice 138. Further, r, is determined from the pressure differential, Pg-PB. The only values which are not known are C and A. However, the solution of the equation, as rearranged above, yields a value which is the product of C, the discharge coefficient, and A, the actual cross-sectional flow area or, in -other words, the effective or equivalent flow area of the nozzle assembly 43 which is precisely the value desired.

Similarly, the compressor turbine wheel 26 can be secured at the outlet end of a test fixture 158 (FIGURE 1l) which also has an annulus 154 and guide vanes 156. A shroud 162 located about turbine wheel 26 and secured atop fixture 158 is provided in order to duplicate the confining effect that the shroud 96 will have within the engine. The remaining portion of the apparatus of FIG- URE ll can be made so as to be identical with that disclosed in FIGURE 10.

The general procedure employed in testing the nozzle assembly is followed in the testing of the turbine wheel 26. That is, the valve 134 is opened to the degree necessary to again establish the differential, AP, between P3 and PB.

The mass-rate of flow, W, through orice 138 is then determined by the application of the general compressible mass ow equation. Once, W, is determined it then becomes possible to solve the same equation, as applied to the compressor turbine wheel, for the combined term, CA, (as discussed with reference to the nozzle assembly of FIGURE 10) of the turbine wheel 26.

By employing the above method it then becomes possible to test a plurality of nozzle assemblies, one at a time, and determine and record thereon the equivalent or eective flow area exhibited by that particular nozzle assembly. Likewise, the compressor turbine wheels can be tested and the equivalent flow area, so determined, recorded on the individual compressor turbine wheels. It then becomes .a matter of simply selectively matching a nozzle assembly and compressor turbine wheel, based on their respective equivalent ow areas, which will result in an equivalent flow area ratio within the critical limits defined on both sides of the ideal equivalentqow area ratio. Such a selectively matched set, when placed within the turbine engine will insure the engine of not only operating within the limits of the choke and stall lines of FIGURE 2 as achieved by gross aerodynamic matching to the compressor, but will further insure the engine of producing an output horsepower at least very closely approaching the designed maximum output horsepower as represented by point 78 of FIGURE 3.

Further, it has been discovered that as between any two nozzle assemblies `or compressor turbine wheels, slight differences may exist between the respective curves defining the mass-rate of gas flow therethrough as a function of the pressure differential across such nozzle assemblies or compressor turbine wheels. This is believed due to the slightly varying aerodynamic characteristics of the particular nozzle .assembly or compressor turbine wheel under consideration. Accordingly, the method for determining and critically selecting a particular nozzle assembly to a cooperating compressor turbine wheel can be modified so as to even avoid such slight discrepancies arising out of such varying aerodynamic characteristics.

That is, preferably, when the ideal equivalent flow areas are empirically determined, they should be determined by employing a pressure differential substantially equivalent to that pressure differential which the nozzle assembly or the compressor turbine wheel, as the case may be, will experience within the engine as during full rated power operation or maximum designed compressor speed.

Subsequently, the pressure differential employed during calibration of nozzle assemblies would vary from that pressure differential employed for Calibrating compressor turbine wheels. For example, with nozzle assemblies, a sufficient pressure differential may be indicated when pressure gage indicates a value which is 0.75 times the then existing barometric pressure, whereas, in the case of the compressor turbine Wheel a reading of 1.17 times the then existing barometric pressure may indicate a proper pressure differential.

In each event, however, the calibration is conducted employing a pressure differential substantially equivalent to the pressure differential actually experienced by the test piece under conditions of actual engine operation thereby avoiding any inuencing factors which might otherwise exhibit themselves due to the aerodynamic characteristics of the turbine wheel or nozzle assembly.

The drawings and the foregoing specification constitute a description of the invention in such terms as to enable any persons skilled in the art to practice the invention, the scope of which is indicated by the yappended claims.

I claim:

1. A method of selectively matching a gas turbine engine compressor turbine wheel to the compressor turbine wheel nozzle assembly, comprising the steps of creating a first pressure differential across said nozzle assembly substantially equivalent to a predetermined pressure differential which said nozzle assembly will experience within said engine during selected periods of normal engine operation, employing said first pressure differential to flow a suitable gasvthrough said nozzle assembly, determining from said pressure differential and the Volume rate of flow of said gas therethrough the effective flow area of said nozzle assembly, repeating the above steps with succeeding nozzle assemblies, classifying each of the nozzle assemblies according to their respective effective flow areas, creating a second pressure differential across said turbine wheel substantially equivalent to a second predetermined pressure differential which said turbine wheel will experience within said engine during selected periods of normal engine operation, employing said second pressure differential to ow a suitable gas through said turbine wheel, determining from said second pressure differential and the volume rate of ow of said gas through said turbine wheel the effective flow area of said turbine wheel, repeating the steps with succeeding turbine wheels, classifying each of the turbine ll l Wheels according to their respective effective flow areas, and selectively matching a nozzle assembly to turbine Wheel so as to have the numerical value of the ratio of their respective effective flow areas within at least one percent of the numerical value of the ratio of their respective ideal flow areas.

2. A method of selectively matching a gas turbine engine lcompressor turbine wheel to the compressor turbine wheel nozzle assembly, comprising the steps o f creating a first pressure differential across said nozzle assembly substantially equivalent to a predetermined pressure differential which said nozzle assembly will experience Within said engine during selected periods of normal engine operation, employing said first pressure differential to flow a suitable gas through said nozzle assembly, determining from said pressure differential and the rate of fiow of said gas therethrough the effective flow area of said nozzle assembly, repeating the above steps with succeeding nozzle assemblies, classifying said nozzle assemblies according to their respective equivalent flow areas, creating a second pressure differential across said turbine wheel substantially equivalent to a second predetermined pressure differential which said turbine wheel will experience Within said engine during selected periods of normal engine operation, employing said second pressure differential to flow a suitable gas through said turbine Wheel, determining from said second pressure differential and the rate of flow of said gas through said turbine wheel the equivalent flow area of said turbine Wheel, repeating the steps with succeeding turbine Wheels, classifying each of the turbine Wheels according to their respective equivalent flow areas, and selectively matching a classified n-ozzle assembly to a classified turbine wheel so -as to have the numerical value of the ratio of their respective equivalent flow areas Within a predetermined range of the numeral value of the ideal ratio of their respective ideal flow areas.

3. A method of matching a gas turbine engine turbine wheel to a cooperating turbine Wheel nozzle assembly each of which is comprised of a plurality of circum- -ferentially spaced radially directed vanes defining respective gas flow areas, comprising the steps offsecuring said nozzle assembly to an outlet of a relatively large plenum chamber, directing a fiow of suitable pressurized gas to an inlet of said plenum chamber, throttling the flow of said suitable pressurized gas to the degree necessary to establish ya first pressure differential between the interior of said plenum chamber and the ambient pres,-`

sure which is substantially equivalent to a predetermined pressure differential which said nozzle assmbly will experienec Within said engine vduring selected periods of normal engine operation, measuring the volume rate of gas flow supplied to said plenuml chamber, determining from said first pressure differential and .said volume rate of gas flow the equivalent ow area of the actual nozzle assembly gas flow area, terminating the flow of pressurized gas to said plenum chamber, removing said nozzle assembly from said outlet, classifying said nozzle as-` sembly according to its equivalent flow area, repea ting the `above steps with succeeding nozzle assemblies, secur-v ing a turbine wheel to an outlet of a relatively large plenum chamber, directing a fiow of suitable pressurized gas to an inlet of said last mentioned plenum chamber, throttling the flow of the last mentioned pressurized gas to the degree necessary to establish a second pressure differential between the interior of said last mentioned plenum chamber and the ambient pressure which is substantially equivalent to `a predetermined pressure differ-v ential which said turbine Wheel Will experience Within said engine during selected periods of normal engine operation, measuring the volume rate of flow of said last mentioned pressurized gas supplied to said last mentioned plenum chamber, determining from said second pressure` differential and said last mentioned volumey rate of gasA flow the equivalent area of the actual turbine Wheel gas flow area, terminating the flow of said last mentioned pressurized gas to said last mentioned plenum chamber, removing said turbine Wheel from said last mentioned outlet, classifying said turbine wheel according to its equivalent flow area, repeating the steps with succeeding turbine Wheels, and selectively matching a classified noz- Zle assembly to a classified turbine wheel according to a predetermined relationship of their respective equivalent flow area classifications.

4. A method of selectively matching a gas turbine engine compressor turbine Wheel to the compressor turbine wheel` nozzle assembly, comprising the steps of creating a first pressure differential across said nozzle assembly substantially equivalent to a predetermined pressure differential which said nozzle assembly will eX- perience Within said engine during selected periods of normal engine operation, employing said first pressure differential to flow a suitable gas through said nozzle assembly, determining from said` pressure differential and the rate of flow of said gas therethrough the effective fiow area of said nozzle assembly, repeating the above steps With succeeding nozzle assemblies, classifying said nozzle assemblies according to their respective equivalent fiow areas, creating a second pressure differential across said turbine wheel substantially equivalent to a second predetermined pressure differential which said turbine Wheel will experience Within said engine during selected peri-ods of normal engine operation, employing said second pressure differential to flow a suitable gas through said turbine wheel, determining from said second pressure differential and the rate of flow of said gas through said turbine Wheel-the equivalent flow area of said turbine WheeLrepeating the steps With succeedingturbine Wheels, classifying each of the turbine Wheels according to their respective equivalent flow areas, and selectively matching a classified nozzle assembly to a classified turbine wheelraccordingto a predetermined relationship of their respective equivalent flow areas. t

V5. A method lof fselectively matching a gas turbine engine compressor turbine Wheelto the compressor turbine wheel nozzle assembly, comprising lthe steps of creating a first pressure differential across said nozzle assembly substantially equivalent to a predetermined pressure differential which said nozzle assembly will experience Within said engine during designed maximum engine speed operation, employing said first pressure differential to flow a suitablegas through said nozzle assembly, determiningfrorn said pressure differential and the rate of flow of said gas therethrough the effective flow area kof said nozzle assembly,repeating the above steps with succeeding nozzle assemblies, classifying said nozzle assemblies according to their respective equivalent iiow areas, creating a second pressure differential across said turbine Wheel substantially equivalent to a second predetermined pressure, differential which said turbine Wheel will experience within said engine during designed maximum engine speed operation, employing said second pressure differential to flow a suitable gas through said turbine Wheel, determining fromV said second pressure differential and the rate of flow of said gas through said turbine Wheel ythe equivalent iiow area of said turbine wheel, repeating they steps with succeeding turbine wheels, classifying each of the turbine wheels according to their ,respective equivalent flow areas, and selectively matching a; classifiednozzle Aassembly to a classified turbine wheel according to a predetermined relationship of their respective equivalent flow areas.

o 6. A method of selectively-matching a gas turbine engine compressor turbinel Wheel to the compressor turbine Wheel nozzle assembly, comprising the steps of creating a first pressure differential across said nozzleassembly, employing said first pressure differential to flow a suitable gas through said nozzle assembly, determining from said first pressure differential .and the rate of flow of said gas therethrough the effective flow area of said nozzle assembly, repeating thev above steps with succeeding nozzle assemblies, classifying said nozzle assemblies according to their respective equivalent flow areas, creating a second pressure differential across said turbine wheel substantially equivalent to said first pressure differential, employing said second pressure differential to flow a suitable gas through said turbine wheel, determining from said second pressure differential and the rate of flow of said gas through said turbine Wheel the equivalent flow area of said turbine Wheel, repeating the steps With succeeding turbine wheels, classifying each of the turbine Wheels according to their respective equivalent flow areas, and selectively matching a classified nozzle assembly to a classified turbine wheel according to a -predetermined relationship of their respective equivalent flow areas.

7. A method of selectively matching a gas turbine engine compressor turbine Wheel to the compressor turbine wheel nozzle assembly, comprising the steps of creating a first pressure differential across said nozzle assembly, employing said first pressure differential to flow a suitable fluid through said nozzle assembly, determining from said first pressure differential and the rate of flow of said fluid therethrough the effective flow area of said nozzle assembly, repeating the above steps with succeeding nozzle assembles, classifying said nozzle assemblies according to their respective equivalent flow areas, creating a second pressure differential across said turbine wheel substantially equivalent to said first pressure differential, employing said second pressure differential to flow a suitable fluid through said turbine wheel, determining from said second pressure differential and the rate of flow of said fluid through said turbine wheel the equivalent flow area of said turbine wheel, repeating the steps with succeeding turbine wheels, classifying each of the turbine Wheels according to their respective equivalent flow areas, and selectively matching a classified nozzle assembly to a classified turbine wheel according to a predetermined relationship of their respective equivalent flow areas.

8. A method of selectively matching a gas turbine engine compressor turbine wheel to the compressor turbine Wheel nozzle assembly, comprising the steps of creating a first pressure differential across said nozzle assembly substantially equivalent to a predetermined pressure differenttial which said nozzle assembly will experience Within said engine during selected periods of no-rmal engine operation, employing said first pressure differential to flow a suitable fluid through said nozzle assembly, determining from said pressure differential and the rate of flow of said fluid therethrough the effective flow area of said nozzle assembly, repeating the above steps with succeeding nozzle assemblies, classifying said nozzle assemblies according to their respective equivalent flow areas, creating a second pressure differential across said turbine Wheel substantially equivalent to a second predetermined pressure differential which said turbine wheel will experience Within said engine during selected periods of normal engine operation, employing said second pressure differential to flow a suitable fluid through said turbine wheel, determining from said second pressure differential and the rate of flow of said fluid through said turbine Wheel the equivalent flow area lof said turbine wheel, repeating the steps with succeeding turbine wheels, classifying each of the turbine wheels according to their respective equivalent flow areas, and selectively matching a classified nozzle assembly to a classified turbine wheel according to a :predetermined Arelationship of their respective equivalent flow areas.

9. A method of selectively matching a turbine Wheel to a cooperating turbine nozzle assembly, compris-ing the steps of creating a first flow of a suitable fluid through said nozzle assembly, determining from said flow of fluid therethrough the effective flow area of said nozzle assembly, repeating the above steps with succeeding nozzle assemblies, classifying said nozzle assemblies according to their respective equivalent flow areas, creating a second flow of a suitable fluid through said turbine wheel, determining from said second flow of fluid through said Wheel the equivalent flow area of said turbine wheel, repeating the steps with su-cceeding turbine Wheels, classifying each of the turbine wheels according to ltheir respective equivalent flow areas, and selectively matching a classified nozzle assembly to a classified turbine wheel according to a predetermined relationship of their respective equivalent flow areas.

References Cited by the Examiner UNITED STATES PATENTS 2,796,658 6/1957 Aller 29-407 X 3,034,343 5/1962 Henry et al. 73-116 3,037,348 6/1962 Gassmann S0-39.16 3,038,331 6/1962 Lindberg 73--116 3,044,262 7/ 1962 Chadwick et al. 60-39.l6 3,077,030 2/ 1963 Carlson 29-407 3,120,053 2/1964 Lewis 29-407 WHITMORE A. WILTZ, Primary Examiner.

ROBERT M. WALKER, THOMAS H. EAGER,

Examiners.

UNITED STATES PATENT OFFICE CERTIFICATE 0F CORRECTION Patent No 3 ,252 ,212 May Z4, 1966 Albert H. Bell III It is hereby certified that error appears in the above numbered patent requiring correction and that the said Letters Patent should read as corrected below.

Column 4 line 6 for The" read [The Column 5, line 17, for "turbines" read turbine line 47, for "matchng" read matching column 6, line 3 for "arrangement" read arrangements column 8 lines 38 to 4l and column 9 lines 27 tonSO in the equations for "r each occurrence read r Column 9 lines 27 to 30 the left-hand portion of the equation should appear as shown below instead of as in the patent: I

P0 '\|TO column l1 1lines 49 and 50 for "experienec" read experience column 13 line 44 for "differenttial" read differential Signed and sealed this 7th day of November 1967.

(SEAL) Attest:

EDWARD M.FLETCHER,JR. EDWARD J. BRENNER Attesting Officer Commissioner of Patents

Claims

1. A METHOD OF SELECTIVELY MATCHING A GAS TURBINE ENGINE COMPRESSOR TURBINE WHEEL TO THE COMPRESSOR TURBINE WHEEL NOZZLE ASSEMBLY, COMPRISING THE STEPS OF CREATING A FIRST PRESSURE DIFFERENTIAL ACROSS SAID NOZZLE ASSEMBLY SUBSTANTIALLY EQUIVALENT TO A PREDETERMINED PRESSURE DIFFERENTIAL WHICH SAID NOZZLE ASSEMBLY WILL EXPERIENCE WITHIN SAID ENGINE DURING SELECTED PERIODS OF NORMAL ENGINE OPERATION, EMPLOYING SAID FIRST PRESSURE DIFFERENTIAL TO FLOW A SUITABLE GAS THROUGH SAID NOZZLE ASSEMBLY, DETERMINING FROM SAID PRESSURE DIFFERENTIAL AND THE VOLUME RATE OF FLOW OF SAID GAS THERETHROUGH THE EFFECTIVE FLOW AREA OF SAID NOZZLE ASSEMBLY, REPEATING THE ABOVE STEPS WITH SUCCEEDING NOZZLE ASSEMBLIES, CLASSIFYING EACH OF THE NOZZLE ASSEMBLIES ACCORDING TO THEIR RESPECTIVE EFFECTIVE FLOW AREAS, CREATING A SECOND PRESSURE DIFFERENTIAL ACROSS SAID TURBINE WHEEL SUBSTANTIALLY EQUIVALENT TO A SECOND PREDETERMINED PRESSURE DIFFERENTIAL WHICH SAID TURBINE WHEEL WILL EXPERIENCE WITHIN SAID ENGINE DURING SELECTED PERIODS OF NORMAL ENGINE OPERATION, EMPLOYING SAID SECOND PRESSURE DIFFERENTIAL TO FLOW A SUITABLE GAS-