i

NASA CR-72562

P_VA-3546

//

DESIGN REPORT

SINGLE-STAGE EVALUATION OF HIGHLY-LOADED

HIGH-MACH-NUMBER COMPRESSOR STAGES

by

N. r. Monsarrat, M. J, Keenan and P. C. Tramm

I " I

1969021491

NOTICE

This _'eport was prepared as an account of Governmentsponsoredwo,k. Neitiler the United States, nor +.'_-National AeronauticsandSpace Administration(NASA), nor any person acting on behalf ofNASA:

A.) Mokesany warranty or representation,expressedor implied,with respectto the accuracy,completeness,or usefulnessof theinformationcontainedin this report,or that the useof any in-re:mutton,apparatus,method,or processdisclosedin this re-portmaynotinfflngeprivately ownedrights;or

B.) Assumestanyliabilitieswithrespectto the useof or for dama_lesresultingfromthe useof any information,apparatus,method

_ or professdisclosedin thisreport.

As_usedabove, "person actingon behalf of NASA" includesany em-ployeeoFcontractorof NASA, o_ emp_eyeeof suchcontractor,to the

_ extentthat suchemployeeor contractorof NASA, or employeeo_suchcontractorprepares,disseminates,or providesaccessto, ",nyinformation

, pursu_,n_tohis employmentor contractwith NASA, or his employmentwithsu-'_contractor.

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NASA CR-72562:' PWA-3546

DESIGN RE PORT

SINGLE-STAGE EVALUATION OF HIGHLY-LOADED

HIGH-MACH-NUMBER COMPRESSOR STAGES

" by": N.T. Monsarrat, M. J. Keenan and P. C. Tramm

f

:: Pratt & Whitney Aircraft DivisionUnited Aircraft Corporation

_. East Hartford, Connecticut 06108

-_.._.

Prepared for:_ NATIONAL AERONAUTICS AND S2ACE ADMINISTRATION

- )

_. July 16, 1969 -_

:z, ;_

Contract NAS3-10482f

NASA Lewis Research Center

._, Cleveland, Ohio

_: L. Reid, Project Manager

_i Fluid System Components Division

1969021491-003

|

t FOR EWOR D

This report was prepared by the Pratt & Whitney Aircraft Division of United Aircraft:: Corporation, East Hartford, Connecticut, to describe the aerodynamic and mechanical

; - design work conducted under Contract NAS3-10482, Single-Stage Evaluation of Highly-Loaded High-Mach-Number Compressor Stages. Mr. L. Reid, NASA-Lewis Research

I _ Center, Fluid System Components Division, was Project Manager.

_-

.?:?•. 2,

ii

1969021491-004

: TABLE OF CONTENTS

Page

FOR E WOR D iiTABLE OF CONTENTS iii

LIST OF FIGURES ivLIST OF TABLES vii

I. SUMMARY 1

II. INTRODUCTION 3III. AERODYNAMIC DESIGN 4IV. AIRFOIL DESIGN 15

A. Rotor 15(= B. Stator 28

C. Stator Slit Design 35V. STRUCTURAL AND VIBRATION ANALYSIS 37

A. Blade and Disc Vibration 37, B. Bladv Stress 37

_ C. Torsional Blade Flutter 41

9:; D. Disc and Attachment Stresses 43_, E. Critical Speeds 44

_ APPENDIX 1 - Aerodynamic Calculation Procedure 457',_ APPENDIX 2 - Loss System 47

APPENDIX 3 - Incidence Selection 53

: APPENDIX 4 - Stream-Tube Analysis in Channels Between Blades 60., APPENDIX 5 - Deviation System 642

_ APPENDIX 6 - Aerodynamic Summary 68_: APPENDIX 7 - Airfoil Sections on Conical Surfaces 73"?

: (Multiple-Circular-Arc Definitions) 73_ APPENDIX 8 - Rotor Slot Design 76{" APPENDIX 9 - Tandem Rotor Design 90, APPENDIX 10 - Airfoil Coordinates for Manufacturing Sections 97

APPENDIX 11 - Nomenclature 105

REFERENCES ii0

iI .,

1969021491-005

LIST OF FIGURES

Number Title Pa_ag&

1 Flowpath 6

2 Rotor and Stator Portion of Flowpath 7

3 Rotor and Stator Inlet and Exit Meridional Velocity Profiles 8

4 Rotor and Stator Inlet. Mach Numbers 9

5 Rotor and Stator Diffusion Factor Profile 11

" 6 Rotor Adiabatic Efficiency 12

7 Statc_: Spanwise Loss Profile 13

_ 8 Rotor Inlet and Exit Relative and Stator Inlet Absolute

Meridional Air Angle Profiles 14i

_' 16_, 9 Multiple-C ircular-Arc Airfoil Definitions

;" 10 Four Views of Rotor 17

._. 11 Axial Projection of Rotor 18

e' 12 Rotor Incidence 19

13 Slotted Rotor Blade Section 20

.' 14 Tandem Rotor Blade Section 21

S

,: 15 Rotor Entrance Region Relative Air Angle Showing Effectsof Annulus Convergence and Front Camber 23

!16 Rotor Entrance Region Relative Mach Number Showing

Effects of Annulus Convergence and Front Camber 24 i

17 Rotor Spanwise Variation of a/a* 25

_",. 18 Rotor Deviation 26

_ 19 Rotor Inlet and Exit Mean Camber Line Metal Angles on

• IV

1969021491-006

' LIST OF FIGURES (Cont'd)

•._ Number Title Page

20 Four Views of Stator 29

21 Axial Projection of Stator 30

22 Stator Leading-Edge Suction Surface and Mean CamberLine Incidence 31

23 Stator Spanwise Variation of a/a* 32

: 24 Stator Deviation 33

i 25 Stator Inlet and Exit Mean Camber Line Metal Angles on:, Conical Surface 34

: 26 Slit Flow to Prevent Corner Separation at Intersection?

• of Stator Suction Surface and Hub Wall 36

27 Blade and Disc Resonance Diagram 38

_!i 28 Maximum Stress Locations 397

29 Effect of Tangential Tilt on Rotor Gas Bendiug Stress 40

. 30 Rotor Goodman Diagram 42

_: 31 Compressor Rig Spring Location and Spring Rates 44

_; 32 Location of Assumed Normal Shock for Loss Model 49

_ 33 Rotor Profile Loss Parameter vs D Factor 5].Y

34 Stator Profile Loss Parameter vs D Factor 52

35 Multiple-Circular-Arc Stator B Loss Data for 50 Percent Spanfrom Contract NAS3-7614 55

36 Multiple-Circular-Arc Stator B Loss Data for 5 Percent Span'_ from Root, Contract NAS3-7614 56

_. 37 Multiple-Circular-Arc Stator B Loss Data for 95 Percent Span' from Root, Contract NAS3-7614 57

I •" V

1969021491-007

P

LIST OF FIGURES (Cont'd)

i Number Title Page

i 38 Wave PatternforSuper-CriticalOperation

of Curved-Blade EntranceRegions 58

'. 39 ChannelArea 62t

s

k 40 Loss DistributionthroughChannel 63b[

[ 41 Modification to Carter's Rule for Rotor 66

42 Modification to Carter's Rule for Stator 67Lp

[43 Recovery Increase vs Slot Angle 78-80

44 Axial Projection of Slotted Rotor 81u

45 Rotor Efficiency vs Rotor Exit Diameter 88

" 46 Rotor Blade Resonance Diagram 89

47 Axial Projection of Tandem Rotor 92

48 Airfoil Coordinates for Manufacturing Sections 98

I"

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ivi |

1969021491-008

LIST OF TABLES

N_lmber Titl.___e Page

1 Aerodynamic Design Point 4

2 Design Parameters 4

I 3 Rotor and Stator Stresses 41

I 4 Disc and Attar-" ment Stresses 43

i 5 Compressors Analyzed for Profile Loss Correlation 50i

6 Aerodynamic Summary - Rotor Inlet 69

_.. 7 Aerodynamic Summary - P_otor Exit 70

_ 8 Aerodynamic Summary - Stator Inlet 71

%_ 9 Aerodynamic Summary - Stator Exit 72

_i_ 10 Airfoil Sec*ions on C, [ Surfaces - Rotor 74

...'. 11 Airfoil Sections on Cor_cal Surfaces - Smtor 75

_, 12 Rotor Coordinates - Sections A0 B, C and D 99

_. 13 Rotor Coordinates - Sections E, F, G and H 100

_:.- 14 Rotor Coordinates - Sections J, K, L and M 101

15 Stator Coordina_s - Sections A, B, C and D 102

: 16 Stator Coordinates - Sections E, F, G and H 103

17 Stator Coordinates - Sections J, K and L 104

_. vii

1969021491-009

ABSTRACT

= DESIGN R E PORTSINGLE-STAGE EVALUATION OF HIGHLY-LOADED

HIGH-MACH-NUMBER COMI_ELSOR STAGES

A compressor stage of 1600 ft/sec (487.7 m/sec) tip speed,a D factor at 10 percent span from the tip of 0.5, and a hub/tip ratio of 0.5 was designed to deliver a pressure ratio of1. 936 with an efficiency of 0. 842. The design flow rate is187.1 lb/sec (84. b69 kg/sec). Three rotors were designed,

_ the first with multiple-circular-arc airfoils, the second with

_ slotted airfoils, and the third with tandem airfoil sections.The same stator will be used with all three rotors.

t,

1969021491-010

I. SUMMARY

A high-tip-speedsingle-stagecompressor has been de_dgnedunder ContractNAS3-

10482. This reportpresentsthedetailedaerodynamic and mechanicaldesign. Thedesignisintendedtodocument theperformance ofa stagewitha high rotortipspeed

and withhighaerodynamic loadingson boththerotorand stator. The feasibilityofusinghighrotter-inletrelativeMach numbers and highbladeloadingstoobtaina highpressureratioatan acceptablelevelofefficiencywillbe evaluated. The importantfeaturesof thisdesignare thatthereisno inletguidevane, the rotor_ipspeed is

1600 ft/sec(487.7m/sec), overallstagepressureratiois 1.936,rotorleading-edgetipdiameter is33 inches(0.8382m),rotorinlethub/tipratio0.5, rotor-tipsolidity1.3, and state)r-hubsolidity2.0.

The aerodynamic designwas based on theapproximatedesignparameters specifiedinthecontract,typicalperformance requirementsfor advanced fans,and theus_ ofex-istingtestfacilities.The designflowper unitof annulusarea at therotorleading

edge is42 Ib/sec/ft2(205.07 kg/sec/m2). The rotorwas designedfor a constantspanwisepressureratioof2.0. Blade lossesforboth rotorand statorwere estim-atedusinga lossmodel inwhich shock lossand profilelossare consideredseparate-

lywiththeirsum equaltothetotalbladeloss. Performance calibrationsusingthismodel predictedan adiabaticrotorefficiencyof 88.7 percent,and an overallstageav-erage adiabaticefficiencyof 84.2percent.

: 7hree rotorconfigurationswere designed,one with multiple-circular-arcairfoilsec-tions,thesecondwith slots,and thethirdwith tandem airfoils.The multiple-circul-

ar-arcairfoilis thebasicbladedesign,theslottedand tandem designsbeingmodific-_ ationswhich use essentiallythesame solidity,and the same chordwisedistributions• of camber and thickness.

_: The multiple-circular-arcrotorairfoilsectionswere designedtoobtainlow lossby ap-propriatecamber distribution.Slotswere designedforthesemultiple-circular-arcYsectionstodeflectsupersonicflowwithslotexitflow,causingan obliqueshock toform

" upstream of thenormal shock. Combined shock lossesof theprecompression and nor-realshockswere calculatedtobe lessthannormal shock losseswithoutprecompres-sion.

Tandem rotorairfoilswere designedby dividingthemultiple-circular-arcsections

jm_tto therear oftheassumed shock impingement pointon thesuctionsurface.Shockboundary-layerinteractioneffectsare isolatedfrom subsonicdiffusionby a stream ofhigh-energyairflowingbetween frontand rear tandem sections.Sincetheinitialcon-

dition of the boundary layer on the rear airfoil should not be affected by the shock, ef--_ ficient subsonic diffusion should be obtained._c

The rotor-inlet hub-tip ratio is 0.50 and was set primarily by mechanical considera-tions. The rotor and stator aspect ratios are 1. 663 and 1.721 respectively, using root

1

1969021491-011

chord and average length. The rotor chord is constant spanwise at 4.4 inches(0. 11176m) which, with 30 blades, yields a rotor tip solidity of 1.297. The low as-pect ratio is not typical of engine designs, but was considered preferable tc use of apart-span shroud which would have given unreliable blade-element data in the regionof the part-span shroud. The stator; chord is also constant spanwise at 3.0 inches,(0.07620 m) and, with 44 blades, yields a root solidity of 2. 047.

Mechanical design included a structural and vibration analysis. Centrifugar stress, gasbending stress, and stress produced by untwist are all well within the capabilities of theAMS 4928 titanium alloy to be used for the blades. Because all critical speeds are out-side the operating range, vibratory stress levels should be low.

t

i 2I

1969021491-012

II. INTRODUCTION

Future aircraft powerplants will require lightweight highly-loaded compressors whichare very efficient over a wide operating range. Pressure ratio per stage can be in-creased considerably above current levels by increasing rotor wheel speed and bladeloadings. In general, such attempts have been discouraging and resulted in severe

aerodynamic penalties. High blade-element losses result from normal shocks at highMach numbers, shock boundary-layer interactions on blade surfaces, and boundary-layer growth due to high blade loadings. Recently-developed design techniques and newconcepts of blade design may be employed to gain this work increase without attendantefficiency loss.

In general, loss '_nhigh Mach-number rotor blading is strongly related to chordwisecamber distribution. The optimum blade shape provides the best balance between shockloss and subsonic diffusion loss. The supersonic curvature should account for axial ve-

locity changes due to annulus convergence, and provide a cancelling system of expan-sion and compression waves across each rotor gap. Channel area between blades must

be adequate to pass the flow, but should be small enough to limit the growth of a separ-ated boundary layer.

A new airfoil termed multiple,circular arc has been designed by NASA. The multiple-circular-arc airfoil is formed by portions of two circular-arc sections, with blade sur-

faces and the mean camber line each mutually tangent at the points where they join.Blades of thir type provide a means of varying blade shape to control camber distrib-ution and minimize overall blade losses. Two recent experimental evaluations of this

airfoil, References* 1to 4 for rotors, and Reference 5 for stators, have provided data

indicating probable success for a stage of this type. In view of this, a high-tip-speedhigly-loaded single stage was designed and will be fabricated and tested under ContractNAS3-10482.

In addition to the basic rotor design configuration, two other rotors were designed toevaluate other advanced concepts. A slotted rotor was designed in which slot dischargeflow is used to cause an oblique shock to reduce the Mach number upstream of the nor-

mal shock and reduce overall shock loss. A tandem rotor was designed with a super-sonic forward blade and a subsonic rear blade. The shock impinges on the forwardblade, whose boundary layer is isolated from the subsonic suction surface by a streamof high-energy air.

This report presents the aerodynamic and mechanical design of the compressor stage,including all three rotor configurations.

*See Page 110 for numbered list of references

" i, 3

1969021491-013

III. AERODYNAMIC DESIGN

Velocity vector diagrams within the flowpath were used to design rotor and stator blade

ele:nents. Rotor and stator blade elements were designed based on their leading andtrailing-edge vector diagrams. Design of the stage flowpath was guided by the aero-

dynamic objectives specified in the contract. Multiple-circular-arc airfoil sectionswere designed for both blade rows. A slotted rotor and a tandem rotor were also de-

signed for the same velocity vectors, although the slotted and tandem rotors are expec-ted to operate with lower losses. The design point for the stage is summarized in Ta-ble 1.

TABLF 1

Aerodynamic Design Point

corrected speed, rpm 11,110

corrected flow, lb/sec (kg/sec) 187.1 (84. 869)

rotor pressure ratio 2.0

stage pressure ratio 1. 936

rotor adiabatic efficiency, % 88.7

stage adiabatic efficiency, % 84.2

The stage design parameters are compared with the approximate values specified by the• contract work statement in Table 2.

TABLE 2

Design Parameters

Final Design Work Statement

rotor inlet tip diameter, inches (meter) 33.0(0. 8382) 30.0 (0.7620) rain.

rotor tip speed, ft/sec (m/sec) 1600(487.68) 1600(487.68)

rotor inlet tip relative Mach number 1. 609 1.6

rotor tip diffusion factor 0.5 at 10% span 0.55from tip

rotor inlet hub-tip ratio 0.5 0.5 max.

rotor tip solidity 1.2 97 1.3

rotor tip inlet-to-exit axial velocity ratio 1. 682 (1.25 at 1.0; 10% span from

_:' tip) .

i i ..... "

1969021491-014

rotor aspect ratio, average length/root chord 1.663 2.0

stator hub solidity 2. 047 2.0

stator root inlet-to-exit velocity ratio 1. 007 1.0

stator aspect ratio, average length/root chord 1. 721 2.0

The stage is designed without an inlet guide vane and the stator exit flow angle is zero atall radii. Rotor exit pressure is uniform spanwise as specified by the contract. Rotordesign prev_ure ratio is 2.0 and was established as a consequence of the intended rotor-

i' tip loading. With the design stator losses the stage average pressure ratio becomesi. 93G.

., The rotor inlet flow per unit annulus area was set at 42 lb/sec/ft 2 (205. 07 kg/sec/m2).! An analysis was made to determine the effect of inlet flow per unit annulus area on

stage efficiency. This study showed that reducing the flow to 40 lb/sec/ft2 (195.2 kg/sec/m2) did not improve the stage efficiency appreciably. The higher value of inletflow per unit annulus area was selected as the more desirable stage because of weightand frontal area considerations.

The rotor inlet hub/tip ratio of 0.5 and aspect ratio of 1.663 are not typical of advancedstages and were set by mechanical design considerations, as discussed in Section V. Nopart-span shroud was used in order to have a maximum of useful blade-element data.The low aspect ratio was necessary from flutter and resonance considerations, and themoderately high hub/tip ratio was required to maintain acceptable steady-state stresses.

The flowpath, including stage inlet and exit ducting, is shown by Figure 1. Figure 2shows the rotor and stator portion of the flowpath in more detail. Flowpath convergencewas used to control levels and radial distributions of meridional velocity throughout thucompressor. Because of the high stage pressure ratio, considerable convergence wasrequired to obtain equal meridional velocity levels at the stage inlet and exit. This con-sideration is desirable from the standpoint of possible applications for a stage of thistype. Studies of possible flowpath contours showed that a meridional angle of 14 degreesover the rotor tip gave the desired rotor-tip D-factor, Overall convergence was obtain-ed with a rotor-hub meridional angle of 20 degrees, with both walls faired to the axialdirection midway through the stator. The spacing between the rotor and stator was setat 1.20 inch (6. 0305 meter) at the hub, to provide sufficient room for interstage instru-mentation. Figure 3 shows inlet and exit meridional velocity distributions for the rotorand stator. Rotor exit meridional velocity is less than at the rotor inlet, particularlyat the tip.

Inlet Mach numbers for the rotor and stator are shown by Figure 4. It should be notedthat the rotor-inlet relative Mach number is supersonic over nearly the entire span. Thestator Mach number is subsonic spanwise with a maximum value of 0.89 occurring atthe hub. _ -

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The chords ofboth therotorand statorwere held constantfrom root totip. Blade-ele-

ment loadingas representedby diffusionfactorisshown in Figure 5. Itisnotedthat

;, boththerotorand statorexhibita trendtowards increasedloadingattheend walls,which isa resultofdesigningforanticipatedhigherlossesintheseportionsofthebla-ding. The bladeelementsbetween 10 and 90 percentofspan are relativelyfreeofthese._;ndeffect_and thediffusionfactorsofboth_m rotorand statorare wellbelow 0.6. The

rotordiffusionfactorincreasesrapidlyfrom 0.5 attenpercentof span from thetip,to

a valueof 0.62 at thetip. The levelof loadingin theneighborhoodo£ therotor tipisconsideredconsistentwith the intentofthecontractualrequirement fora 0.55 rotor-I

! tip diffusion factor.

i The procedureused to calculatethevelocitytrlahgleswas an iterationinvolvingthe

flow-fieldcalculationgiveninAppendix 1 and thelosssystem giveninAppendix 2. Theflow-fieldcalculationproceduresolvesthecontinuity,energy, and radialequilibrium

[; equationsincludingeffectsof streamlinecurvature,slope,and entropygradient. Stag-nationpressuresand temperaturesare enteredspanwisebehind therotorand stator.

The iterationisrequiredbecause thevelocitytrianglesare dependentupon losses

i which, inturn,are dependentupon tho,blade-inletMach number and loading. In gener-i al, ittakesthreepasses beforetheb._._deand aerodynamic calculationsare consistent.

Figures6 and 7 show thespanwisevarih.tionof rotoradiabaticefficiencyand stator

i lossrespectively.Boundary-layergrowkh on theflowpathwallswas estimatedusingtestdatafrom severalsingle-stagerigsof similarsize. Effectiveend wallswithinthe

'. actual flowpath were used to account for the displacement thickness of the wall bound-ary layers. The ratio of actual-to-effective area is 1.005 ahead of the rotor, 1.01 be-

i ween therotorand stator,and i.015 behindthe stator.Area reductionswere obtained

I by locatingeffectivewalls atequalpercentagesofspan insideoftheactualwalls.

Figure 8 shows the rotor inlet and exit relative air angles and the stator inlet absolute[ air angle. The stator turns the flow to axial at the rig discharge _md flow enters the to-[ tor without swirl. The design vector diagrams are presented in Appendix 7 at the rotor

[ and stator leading and trailing edges along streamlines for ten percent of flow ra-every

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1969021491-020

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IV. AIRFOIL DESIGN

A. Rotor

The rotor chord is constant spanwise at 4.4 inches (0. 11176 m) which, with 30 blades,yields a rotor-tip solidity of 1.297. The low rotor aspect ratio (1.66) is not typical of

front stages for engine design, but was considered preferable to use of a part-spanshroud.

The rotor uses multiple-circular-arc airfoil sections designed on conical surfaceswhich approximate streamlines of revolution. Airfoil sections are defined by specify-ing total and front chord, total and front camber, maximum thickness and its location,and leading and trailing edge radii, as shown by Figure 9. Figure 10 shows severalviews of the rotor.

The total chord and thickness ratio were so selected as to provide mechanical integrity.The ratio of front chord to total chord was set to provide a transition point just aft ofthe assumed shock location, discussed in Appendix 2 and shown by Figure 32. Figure11 is an axial projection of the blade and shows the chordwise location of the transitionpoint versus radius. The maximum thickness was so located that for a given total cam-ber, front camber, total chord and front chord, the leading edge wedge angle is theminimum possible without a cusp shape in the front portion of the blade.

The thickness ratio was set at 0.10 at the root to provide mechanical integrity. Thisis somewhat thicker than optimum from loss considerations, but was necessary from astructural point of view. The thickness ratio varies linearly from the root to a valueof 0. 025 at the tip.

Multiple-circular-arc rotor airfoil forward sections were tailorod to satisfy entrance-region requirements and to provide the desired throat width in _e channel betweenblades. The approach and design procedure for the forward section is discussed fullyin Appendix 3. Rearward sections were designed to give the necessary air turning fordesign work input. Deviation angles were calculated using Carter's rule plus an exp-erience factor, as discussed in Appendix 5.

Incidence was set at approximately 1.5 degrees to the suction surface at a point mid-way between the leading edge and the first captured Mach-wave emanation point, PointB' in Figure38. ,Incidenceatthispoint,togetherwithentrance-regionand channel-area considerations,determined leading-edgeincidence.IncidenceatPointB' and at

theleadingedge isshown in Figure 12. Flow channelsbet-veenfrontand rear bladesectionsinbothslottedand tandem designsare downstream of thefirstcapturedMach

wave, where theycannotaffectentrance-reglonflow(Figures13 and 14).

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Figure 11 Axial Projection of Rotor

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Figure 13 Slotted Rotor Blade Section

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1969021491-030

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Figure 14 Tandem Rotor Blade Section

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1969021491-031

The radialdistributionof minimum locala/a* between multiple-circular-arcblades

was setbetween 1.03and 1.05. The dcsiredthroatwas obtainedby adjustingfront

camber as discussedinAppendix 4. Because of thisa/a* criterionand thehighrela-tivetipMach number, thereiva strongsuction-surfacecompression atthetip. Fig-ure 15 shows thedifferencebetween thefree-streamair anglesiftherewere no blad-

ing,and thesuction-surfacemetal angleat theshock impingement location.Assumingthattheflowisattachedto thesuctionsurface,Figure 16 shows thechange inMachnumber at theassumed shock locationdue tothefrontcamber. There isa maximum

increaseinMach number of approximately0.2 at 90 percentfrom thetip,and a reduc-

tionintipMach number of 0.16. The resultantlocalchannela/a* throughtherotorpassage forten annularetream tubesis shown by Figure 17.

Rear-section camber of the multiple-circular-arc rotor was set to obtain air turningcorresponding to the destred work input. Calculations of blade-element deviation re-quiL'_d knowledge of solidity, inlet and exit air angles, incidence, and chordwise loca-tion of maximum camber. The tip solidity was specified to be 1.3, and solidity be-tween hub and tip was determined by stress and vibration criteria. Inlet and exit airangles were determined by a streamline calculation. Maximum camber point is depen-dent on deviation, and was determined by an iteration. The ratio of forward chord tototal chord was held constant, and the ratio cf fl'ont camber to total camber was var-ied to obtain a combination of camber and maximum camber point which produced thedesired exit-air angle. Deviation was set by Carter's rule plus an experience factoras discussed in Appendix 5. Deviation verses span is shown by Figure 18. Rearwardblade sections of the slotted blade are exactly the same as for the multiple-circular-arc blade. Rearward tandem sections have small differences in suction-surface con-

tour extending from the leading edge of the rear blade to approximately 25 percentchord. Deviation angles are assumed to be the same as for the muttiple-circular-arcrotor.

An additional correction was required to compensate for blade untwi_t due to centrifu-gal loading. Static blade angles were set so that the untwls_ at design speed producesthe desired operating blade angles. Untwist angles calcu]ated for the slotted bladewere the same as for the unslotted blade. Untwist at the t_ndem blade tip is approxi-mately 0.1 degree less than that for the multiple-circular-arc blade tip. Rotating, asweli as static blade angles, are shown by Flgurc 19. Multiple-circular-arc rotor ge-ometry on the conical surfaces used for design is summarized in Appendix 7.

The slotted and tandem rotor designs were essentially modifications of the multiple-circular-arc design. The slotted rotor aerodynamic design is described in Appendix8, and the tandem rotor design Is described in Appendix 9,

22

I

1969021491-032

1.8

Figure 16 Rotor Entrance Region Relative Math Number Showing Effects ofAnnulus Convergcnce and F_nt Camber

#

' 24

1969021491-034

|

[[

t '<1.2:

1.2! a) 0% TO 10% FLOW l J b) 10% :ro 20% FLOW

, I_ROM,,, II _ROM,,,% _. --- -- 1.1

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Z-ZLE ,,,., METERS Z-ZLE ,'-- METERS

,.2_ ,.21,,,0%To80%] -_ ]:=,., .....f__ ,.I /

1.31.3 i) 80% TO 90% j) 90% TO 100%

/I1.2 1.2 ------ I •

, XI /1.0 1.0

0 1 2 3 4 1 2 3 4

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Z-ZLE,-.,METERS Z-ZLE"-"METERS

Note: x = a/a* at roLor inlet relative Mach number

Figure 17 Rotor Spanwise Variation of a/a*!

25

1969021491-035

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1969021491-036

2T

1969021491-037

For manufacturing purposes the airfoil sections _,cre redefined o._ planes normal to a linetermed the stacking line, which is defined as a radial line passing through the center ofgravity of the root conical section. A computer program provided a smooth fit of the air-foil properties and produced a set of coordinates for manufacturing purposes. These co-ordinates are tabulated in Appendix 11.

B. Stator

The stator chord is constant spanwise at 3.0 inches (0. 0762 m) and with 44 blades, yieldsa root solidity of 2. 047 al,) _ _ aspect ratio of 1.72.

The stator is a multiple-circular-arc airfoil, as shown by Figure 9, and is designed on

conical surfaces approximating streamlines of revolution. Figure 20 shows several viewsof the stator.

The stator design criteria were very similar to those of the rotor. The front chord wasselected to provide a transition point just aft of the first covered section, as shown byFigure 21, similar to what was done for the rotor.

The stator thickness ratio was set to vary linearly from 0.05 at the root to 0.07 at the tip,which should result in low losses and yet be large enough to provide mechanical integrity,as discussed in Section V, below. Maximum thickness was located in the same manner

as for the rotor, to provide a minimum wedge angle without a cusp front section.

Incidence was set at roughly zero degrees to the suction surface at the leading edge, asdiscussed in Appendix 3. Figure 22 shows the stator suction surface and mean camberline incidence.

The stator throat a/a* was set at 1.05 times the a/a* implied by the inlet Much number,

and is discussed in Appendix 4. Figure 23 shows the local channel a/a* through the statorpassage for ten stream tubes spanwise.

Stator deviations, shown by Figure 24, were calculated using Carter's rule with a correc-tion factor implied by the test results of Reference 5 Appendix 5 discusses the deviationand shows this correction factor to Carter's rule. Figure 25 presents the stator inlet andexit mean camber metal angles spanwise on the conical surfaces used for design.

The stator geometry is summarized in Appendix 7 on the conical surfaces used for design.For manufacturing purposes, the airfoil sections were redefined on planes normal to aline termed the stacking line defined as a radial line passing through the center of gravityof the root section on the plane normal to the stacking line. A computer program provideda smooth fit of the airfoil properties and produced a set of coordinates for manufacturingpurposes. These coordinates are tabulated in Appendix 10.

28

1969021491-038

- Q " .

Convex Surface Concave Surface

Lead ing Edge Trailing Edge

Figure 20 Four Vie VB of Stator

29

32

0.8 1 BEGINNING OF _.

COVERED SECTION

:'

30

TRANSITIONPOINT

[ 28

0.7_i

LLII"M.I ¢_

= 26

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, __ _= 1; 24

i: 0.6m a

1!'

L 22I

II

II' 18

4.0 5.0 6.0 7.0 8.0 9.0AXIAL LOCATION--- INCHESiI i , I

0,1 0.2i, AXIAL LOCATION",.'METERS

Figure 21 Axial Projection of Stator

: 30(

I I mmlul i- _ j lU --I ...... . .......

1969021491-040

31F

I I_ Ili_l I t I ill I |

1969021491-041

1.2 a)O%TO 1-0%FL_ 1.2 b) IO%TO20%PLOVV

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[ -_ 1,1 1.1 -

1.0 1.C ,

1.2 =e) 40% TO 50% J 1.2 f) 50% TO 60% /

' "_ 1.1 J 1.1 _

: 101.0 i . ,

i 1.2, O) 60% T() 70% 1.2 h) 70% TO 80% -!

-_ 1._ _- 1.1 / -I

1.0 1.G

121J)

1.2 i) 80% T61/90% / _'--

% 1.1 _ I..I/i

1.00 1 2 3 0 2 3

Z-ZLE ,-- INCHES Z.ZLE "INCHESL , I I • I I I I I0 0.02 3.04 0.06 0,08 C 0.02 0.04 0.06 0.08

Z'ZLE "" METERS Z-ZLE ,_,METERS

Note: x = a/a* at stator inlet Mach number

: Figure 23 Stator Spanwise Variation of a/a*

32

1969021491-042

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1969021491-043

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34

I m_ aM m i II J film

1969021491-044

C. Stator Slit Desi_n

In order to reduce total pressure losses and attain higher lift coefficients in axial flowturbomachinery, it is necessary to prevent end-wall boundary-layer seraration. Datafrom Reference 5 sh 3w that lossos in the end-wall region limit stator ]oadings. Otherexperimental data6 show that substantial improvements in performance (2 points inefficiency) can be gained by applying boundary-layer control in the corner of a blade orvane surface and the annular wall or casing. Peacock6 reduced compressor cascadelosses significantly by extracting a small amount of flow _rough a slit at the intersectionof the wall apd blade suction surface. His data were analyzed aud slit design criteriawere developed. It was found that most of the attainable benefit was obtained by bleedingoff the flow common to displacement thickness growths on the vane surface and _he wall.This is the flow in which the direct and turbulent dissipat:on of kinetic energy is greatest7.As shown in Figure 26, a slit width greater than the w_c : _, his critical region requiresan increased flow rate to eliminate separation.

With flowrate determinedby slitwidth,many combinationsofslitlengthand flowve-locityare possible.In thisdesign,thvslitextendsfrom 15 percentto85 percentofchord. This isfurtherforward thantheoptimum configurationreportedinReference6,but itprovidessuctioncloserto theminimum pressure point,and covers thearea of

possibleshock impingement duringtransoE:.coperation.Sonicbleedvelocitywas se-lectedtopreventrecirculation. A slitwidth of0.025 inch(0.000635m) was selectedtoprovidean adequatemargin between slitflowcapacityand thecalculatedflowratetopreventseparation.

, 35

ii J m I ............... _.,__!

1969021491-045

0" AT HUB = 2.047

ASPECT RATIO = 1.721

0.0035 /

//

0.0030 . V .J__FLOW CAPACITV /IOF CH()KED /

0,002, ,, /. .,,,,"."e (0.85NOZZLEFLOW_' /

COEFFICIENT1/

0.0020 /' /'_- .....MINIMUM BLEED

Q / / TO PREVENT,.,,Iu_ / SEPARATION/O 0.0015 ......... f I

I.., /

,,,I 0.0010 / /

/ I ESTIMATE PGROWTH OF

J SUCTION SURFACE

/ !! DISPLACE K:ENT

0.0")0s / I THICKNESS I / I "'

, ,:/' DESIGN WIDTHo_ I1 1 1 l0 0.002 0.004 0.006 0.006 0.010 0.012 0.014

SLITWIdTH/CHORD

Figure 26 Slit Flow to Prevent Corner Separation at Intersection of StatGr SuctionSurface and Hub Wzll

36

_____..,,,,_.... m . e _ i i

1969021491-046

V, STRUCTURAL AND VIBRATION ANALYSIS

Two major problem areasthatthe rotorbladvmechanical designwas tailoredtoavoid

were first-modebendingresonance withtwo excitationsper revolutionand staticstressdue tobladeuntwistathighspeeds.

A. Bladeand Disc Vibration

Because circumferential distortion testing requires operation in an environment of twoexcitations per revolution, it was mandatory that there be no 2E resonances in the op-erating range. Airfoil spanwise length was set at the minimum allowed by the contractwith an inlet hub/tip ratio of 0.50. A parametric study of vibration frequency as a func-tion of blade aspect ratio was made, with root maximum thickness-to-chord ratio setat 0.10. An aspect ratio of 1. 663 gave its lowest natural frequency (first bending mode)higher than the frequency of two excitations per revolution at 110 percent of designspeed. To avoid a large drop in frequency from dovetail root and disc flexibility, afirtree attachment was used in conjunction with a large steel disc. The resultant over-all first bending mcde resonance frequency margin at 110 percent of design speed is 7percent, which is considered adequate. The frequency-speed diagram is ,_hown inFigure 27.

B. Blade Stress

Combined centrifugal pull and untwisting stresses were calculated at 110 percent of de-sign speed. A comparison was made to allowable stresses based on AMS 4928 titaniumalloy at 150°F (338.7°K).

There are two spanw e maximum stress locatLons, the root section which has local highstresses at th_ leading and trailing edges, which from past experience is not a problem,and the 90 percent span combined stresses. The combined stresses at 90 percent spanand 110 percent of design speed for the final design with 0.5 _ub/tip ratio and 0.10 t/cratio were calculated to be 74,000 psi (52,029,000 kg/m2), which is well below the 0.2percent yield strength of 116,000 psi (81,560,000 kg/m2).

Gas bending stresses with centrifugal restoration were calculated at design speed for vari-ous tangential tilts (}igure 29). Airfoil stresses were minimized for the combination ofload and no-load conditions. The selected t_ngential tilt is 0.034 inch (0. 000864 m),which results in a maximum bending stress of 3800 psi (2,672,000 kg/m2).

The first bending mode maximum-vibratory-stress _ocation is at the root trailing-edgeposition. The 90 percent span location is predicted to have only 32 percent of vibratorystress in first mode compared to the root trailing edge, as shown by Figure 28

37

1969021491-047

38

1969021491-048

1969021491-049

lOx 108 __ 20 - I jf L3

3

2

3 "_3__ _ TANGENTIAL TI = 0.034 INCH

% (0.000864 METE R)

--10 x 106 --

--20

21

1

x 106 -- __ O WITH LOAD_20

lWITHOUTLOAD 2

1 = LEADING EDGE

2 = TRAILING EDGE

3 = MAX. CONVEX SURFACE CAMBER

I I0.1 0,2 0,3

TANGENTIALTILT ~ INCHES

I I I !0 0.002 0.004 0.006

TANGENTIALTILT ~ METERS

Figure 29 Effect of Tangential Tilt on Rotor Gas Bending Stressf_

_ 40

g

1969021491-050

A Goodman diagram was developed for the t_lade material at 150°F (338.7°K) and is pre-sented in Figure 30. The maximum allowable continuous vibratory stress is 10,000 psi(7,031,000 kg/m 2). Since critical resonances were removed from the operating range,vibratory stress levels should be less than 10,000 psi (7,031,000 kg/m 2) and the rotorshould have satisfactory fatigue characteristics. The rotor and stator stresses are sum-marized in Table 3.

TABLE 3

Rotor and Stator Stresses

Rotor Stator

material AMS 4928 AMS 5613number of airfoils 30 44rotor speed (110 percent),

rpm 12,221 -root centrifugal force, lb 45,600 (20,684 kg)root centrifugal (P/A) ten-

sile stress, psi, 36,000 (25,311,000 kg/m 2) -maximum gas bending stress

at 100 percent speed, psi 3,800 (2,672,000 kg/m 2) 3,815 (2,682,000 kg/m 2)tangential tilt, inches

tip 0. 034 (0. 00964 m) -root 0 (0m) -

10 percent span above plat-form, airfoil P/A +untwist tensile stress, psi 74,000 (52,029,000 kg/m 2) -

0.2 percent yield strength,psi 116,000 (81,560,000 kg/m 2) 110,000 (77,341,000 kg/m 2)

C. Torsional Blade Flutter

The torsional flutter parameter V/bcot, (where b is the blade semi-chord at 75 percent

span and cot is the torsional frequency) was calculated for the rotor blade. For theseblades, the torsional flutter parameter is 0.932, which is well below typical values whereflutter problems occur.

Torsio,,al flutter stability was brought about through minimizing the flutter parameter byincreasing airfoil root thickness. The thickness-to-chord ratio at the airfoil root and tipare G.100 and 0.025 respectively. Th.; thin tio section was set primarily by aerodynamicconsiderations, but sufficient thickness was u,Jed to avoid panel flutter.

:, 41

1969021491-051

-I

t,,,OI-,-

/I•. I ' 8

u. i_

< _ _

t_ r,.'3

__• ' x,c e .2,

- /.1,,Jg,®oO

I$_ "' SS31:I.LSAI:IOJ.VH81^

42

1969021491-052

D. Disc and Attachment ~tresses

A firtree root attachment was used as discussed previously to provide a rigid rOut attach-ment and obtain adequat.e 2E margin. A conventional dovetail type of root attachment would have caused a fifty percent reduction iu 2E margin. Table 4 summarizes the disc and attachment stresses.

TABLE 4

Disc and Attachment Stresses

Assumed Operating Conditions:

N = 12,220 rpm (110 percent of design speed) T = 200°F (366. 5 ole) (no radial temperature gradient) pressure difference across disc = 1.5 psi (1054.7 kg/m2) total disc rim load = 3,845,800 lb (1,744,470.0 kg)

Materials:

disc blade

AMS 6415 AMS 4928

Percent Allowable Location Stress TyPe Stress, psi Stress, kg/m2 Stress

A-disc bore tangential 137,644 96,777.000 98.4 B-disc lug tensile 40,600 28,546,000 31.2 C -blade root tensile 36,000 25,312,000 32.8 D-firtree attachment tensile 19,850 13,957,000 18.0 E-tooth bending tensile 17,350 12,198,000 18.8 F-tooth shear 19,800 13,921,000 30.6 G-tooth bearing 5'7,600 40,499,000 28.6

Total radial centrifugal growth at blade tip = 0.0268 inch (0.000681 m)

c

8

........

A

43

E. Critical Speeds

A critical speed _nalysis was performed for the two-bearing system shown by Figure 3i.

The critical speed for a 6000 lb (2722 kg)thrust is 14,738 rpm, or 32.5 percent critical

speed margin at design speed.Q

44

1969021491-054

i

I

A PPE NDIX I

Flow FieldCalculationProcedure

I

I 45

i

1969021491-055

APPENDIX 1

Flow FieldCalculationProcedurest

• The aerodynamic flowfieldcalculationused inthisdesignassumes axisymmetr'.cflowl

and uses solutionsof continuity,energy, and radialequilibriumequations.These equa- -t tionsaccountfor streamlinecurvatureand radialgradientsof enthalpyand entropy,but

viscousterms are neglected.Calculationswere performed on stationsorientedatantE angle ),withrespecttotheaxialdirection.

The equationofmotion is inthe form of:b

l[

1 aV2m 72 m V_ 1 a p- cos (),-e) + sin(X-e)- -- + - 0!. 2 am Rc r p a r[

I

: ae

t Re - - streamlineradiusof curvaturel amt

[[ Enthalpy rise across a rotor for a streamline _ ts given by the Euler relationsh;pi,

'. AHRoto r = (U2 Vs 2)¢j - (Ul Ve 1) _kI

Weight flow is calculated by the continuity equation|J

I y tip

: f sin O, - e)W=2_r.] Kp Vm y dysin X

y root

where K I£ the local blockage factor and y is the length along the calculation station fro:nthe centerllne to the point of Interest.

46

1969021491-056

APPENDIX 2

Loss System

. 47

1969021491-057

APPENDIX 2

Loss System

The loss :nodel that was developed is an extension of the NASA model which combines ashock loss with a profile loss to obtain total loss 8. To obtain shock loss a normal shock

'. is assun%d to originate at the leading edge and to be normal to a mean camber line atmid-gap as - _)wn by Figure 32. To obtain the Much number of the assumed n r,nal shock,a Prandtl-Me;,er expansion from the free-stream relative Mach number is calculated from

[

; the turning of tt:e flow from the inlet to the suction-surface shock intersection point. The,_ camber of the s_ction surface between the leading edge and the shock inter_ection point' is termed the supersonic suction-surface camber (t$_Sss ). Thus, the flow turns from

L the free-stream condition through an incidence angle iss to the suction-surface leadingI edge and is further turned through the angle 0SSss to the shock interaction point. Three-

dimensional effects were considered small and were ignored.}

t

For a Prandtl-Meyer expansion, Macb "umber and turning are related by:I|

i[

[ = / k-I# f(i)=._ tan"4. (M2-1) - tan"t_M 2-1V k-I V' -iTV

[ ThenI.ul = f(M'l )

i[ Vs =#t +_ssss + tss

M' = rl (#s)S

" , _ M's + M'tI M savg. 2Ii

I Using the general loss model where total losr_ is equal to the sum o_ shock loss and pro-I, file loss, several compressors were analyzed. Aerodynamic conditions were calculated!

: along the leading and trailing edges of both rotors _nd stators with a data reduction pro-igram that uses test data in the form of pressures and temperat,_res as iliput. Thin pro-Igram provides a three-dimensional axtsymmetric compressible flow solution of tbe con-

i tinuity, energy, and radial equilibrium equations. It is discussed in Appendix 1. Span-m

wise values for profile loss (Op) were obtained by subtracting shock los_ from the total!' los;_. A profile loss parameter (_p cosB 2/2¢ ) was calculated and correlated against

diffusion tactor. Table 5 lists the compressors that were analyzed _o obtain this correla-tion of profile loss.

i!

Compressors 1 through 5 were used for the rotor profile loss correlation shown in Figure33. Compxessor 4 was used for the stator profile loss. This compressor was a NASA-sponsored high-Much-number st_tor ret, earch rig in which three different stators wereevaluated. The stator profile loss correlation i[, shown in Figur_ 34.

48

1969021491-058

ROTATION

LEADING EDGE l

SUPERSONIC SUCTION SURFACE CAMBERw _*SSssLINE OF INLET

RELATIVE M,ACHNUMBER. M 1

ASSUMED/ NORMAL SHOCK

/ M ' _ "_ ...MEAN CAMBER LINE AT

"'AVERAGE . _f MID-GAP

Figure 32 Location of Assumed Normal Shock for Loss Model

|, !

t

1969021491-060

TABLE 5

Compressors Analyzed for l>rofilo Loss Correlation

Match

Rotor Stator Pres. Tip Speeds

No. Compressor Design _Series Series Ratio Analyzed

ft/sec {m/seel

1 JTF14F single-s tage straight not - 1790 1630fan entry analyzed (545. 6) (496.3)

1465 1300

(446.5) (396.2)

2 JTF17 two-stage straight not - 1550 1395fan entry analyzed (472.2) (425. 1)

3 JT9D single-stage straight not - 1570 1430fan entry analyzed (478.5) (435. 9)

1287

(392.3)

4 Contract single-stage C.A. C.A. & 1.5 1315 1197NAS3-7614 f._n M.C.A. (410.8) (364. 8)

1075 955

(327.7) (291. 1)

5 Contract single-stage C.A. & not 1.6 1400 1260NAS3-7617 fan M.C.A. analyzed (426.7) (384.0)

' 50

1969021491-061

f_

Lo

\_ ° °\CD c_

CJ

o

¢gCO

0_4

0

o.

_. _ _. _ 8. 8 _ oOo c_ o _ o _

PZ

1969021491-062

r_

I--

0

°_

0

/ °o _ _ _ _ _ o

-_ 52

._.

1969021491-063

APPENDIX 3

Incidence Selection

534

1969021491-064

APPENDIX 3

Incidence Selection

For blade elements with supersonic inlet relative Mach numbers, the _bsolute flow capa-city is established by incidence and entrance-region curvature. The entrance region is

defined as the portion of the blade suction su_-face forward of the first captured Mach wave,- as shown on Figure 38. Consequently the selection of incidence is of major concern.

However, this condition is relaxed for subsomc blade elements and additional latitude isavailable in selecting incidence.

Subsonic Flow

The philosophy used in this design was to set the subsonic incidence at approximately zero

degrees wi+.hrespect to the leading-edge suction surface, except near the blade ends.

For the rotor o'_ly the very root is subsonic. Data from Contract NAS3-7617 that wasreviewed, indicated tidal a leading-edge suction-surface incidence of approximately -2.0

degrees gives minimum blade-element loss for the root. This incidence was faired intothe incidence set for the supersonic portions discussed below.

The entire stator operates subsonicaUy at design. This stator is very similar to thoseused in Reference 5. Figure 35 shows typical mid-span stator loss data indicating thatan incidence of zero degrees to the suction surface is appropriate. Figures 36 and 37

' skew s';ator data for roots and tips respectively. This data indicates" that both the rootand tip suction-surface incidence for minimum loss should be positive. Figure 22 shows

, the selected spanwise variation of i_cidence for the stator.

Supersonic Flow

Efficient alignment of supersonic flow to the suction surface in the entrance region ofrotor tip airfoils was a major design consideration. If the blade suction surface curva-ture upstream of the first captured Mach wave does not correspond exactly to free-stream flow, then alignment with the surface causes expansion or compression wavesshown in Figure 38. When meridional flow is subsonic, the waves formed between theleading edge and the first captured Mach wave propagate forward, and adjust upstreamflow conditions. Compression or expansion from this portion of the suction surface mustbe balanced by a corresponding expansion or compression of equal strength respectively,since each blade must have the same inlet flow conditions. Far upstream of the leadingedge, compression and expansion waves cancel each other, and the flow is axisymmetric.Thus, a shock at the leading edge (Figure 38a) must be followed by an expansion to aneutral point B', where aligned flow angle equals the axisymmetric flow angle, followedby continued expansion to point B sufficient to cancel the entire leading-edge compression.Conversely an expansion about the leading-edge (Figure 38b) must be followed by a sys-

54

I

1969021491-065

0.20

o''o-- '"J SPEED MACH NO. RANGEI,MnP 0 120

1.01 -- 1.06

_ 110i,M 0.94- 1.01

IX _ 100

o,. 0.05 0.87 -- 0.93

--' 0 90 0.78 - 0.85

I.- 0 700 0.59 - 0.66

P" _ 50 0,40 0.44

0 DESIGN POINT 0.94

0 _ I I I--8 -4 0 4 8 12 16 20

INCIDENCEANGLE,SUCTIONSURFACE,iI, DEGREESI _I I I I I I

0 4 8 12 16 20 24

INCIDENCEANGLE,MEANCAMBERLINE,im,DEGREES

Figure 35 Multiple-Ch_ular-Arc Stator B Loss Data for 50 Percent Span fromContract NAS3-7614

55

-i

1969021491-066

0.4U _ I

0.35

_:'a _ o_ 0.3o ......EI,M

° _0

..........eL % DESIGN.J 0.20 ......... SPEED MACH NO. RANGEI--° _ gI,- 120 ;.01 - 1.12110 1,00- 1.06

O 100 0.95 - 0.99

O 90 0,85 - O.92

0.15 0-- 70 0.G5- 0,72

(1) _ 5o o.4;-o49E_ DESIGN POINT 1.07

0,10-8 -4 0 4 8 12 16 20

INCI,DENCEANGLE,SUCTIONSURFACE,ir DEGREESt l l I,, I I I0 4 8 12 16 20 24

INCIDENCEANGLE,MEANCAMBERLINE,im. DEGREES

Figure 36 Multiple-Circular-Arc Stator B Loss Data for 5 Percent Span fromRoot, Contract NAS3-7614

56

1969021491-067

0.35 '

/o=...... _ ....°//

,a 0 1

¢,,1

E _x0

_0 _ _ - ,.....E

_ ,

_" 0.151

h-O % DESIGNI,.- SPEED MACH NO. RANGE

120 0.87 - 0.9O110 0.81 - 0.890.10

100 0.75 - 0.81

90 0.63 - 0.72

70 0.47 - 0.54

II 0 50 o_- o.,oO DESIGN POINT 0.85

0.06 _ i ,-8 --4 0 4 8 12 18 20

INCIDENCEANGLE,SUr'ON SURFACE,i_ DEGREESI I _ I I..... I - I I0 4 8 12 16 20 24 28

INCIDENCEANGLE,MEANCAMBERLINE,im,DEGREES

Y.':"_re 37 Multlple-Ctrcular-Are Stator B Loss Data for 95 Percent Span fromRoot, Contract NAS3-7614

1969021491-068

\ \\

\

%\

BLADE LEADING EDGE

a) CONVEX ENTRANCE REGION (INBOARD OF 10% OF ROTOR SPAN)

AXISYMMETRIC WAVE

__,_,.'_ "_ _ EXPANSION OR COMPRESSION WAVESHOCK WAVE

\

BLADE LEAOIN.O EDGE

COMPRESSION FIELD B_SHOWN AS SHADED AREA

B

b) CONCAVE ENTRANCE REGION (ROTOR TIP)

.j

Figure 38 Wave Pattern for Super-Critical Operation of Curved-Blade EntranceRegions

#

1969021491-069

tern of cancelling compression waves. Thus, a repeating pattern of waves with a periodof one blade gap exists at any radial location, and only one Mach wave in the expansionor compression fields on the leading edge b_s the axisymmetric flow angle.

If entrance-region deslgn of supersonic sections does not accommodate this adjustment,then it must take place through detache# shocks or blade boundary-layer buildup, orcombination of both. This sort of inefficient adjustment can be a source of significantloss in supersonic blade elements.

It was assumed that the neutral point on the blade-suction surface, where ali_ed flowangle equalled the axisymmetric flow angle, was halfway between th_ leading edge andthe emanation point of the first captured Mach wave, as shown in Figure 38. A detailedconstruction of a typical flow pattern verified that this assumption i_ a good approxima-tion.

Engrance-region relative flow angles for free-stream flow were calculated by the flow-field calculation procedure discussed in Appendix 1. These angles were decreased ap-proximately 1.5 degrees to account for increased axial velocity due to blade leading-edgeblockage, and for boundary-layer development on the suction surface. Then the blade-element suction surface was set tangent to this a ,justed free-stream flow angle at pointB' in Figure 38, which is midway between the leading edge and the emanation point ofthe first captured Mach wave, point B. Thus for supersonic flow, the airfoil suctionsurface was 91igned with the axisymmetric flow angle at point B', with cancelling zonesof expansion and compression on either side of this point. The resultant design incidenceangles are shown on Figure 12.

"_ 59

1969021491-070

APPENDIX 4

Stream-Tube Analysisin Channels Between Blades

60

I

1969021491-071

A PPENDIX 4

Stream-Tube Analysis in Channels Between Blades

Supersonic blade element data in References 1 through 4, show minimum loss occurringwhen blade passage a/a* was approximately 1.03. Subsonic stator blade-element datafrom References 5 and 4, and subsonic two-dimensional cascade data show minimum lossoccurring when channel throat area was approximately 1.03 s cos 5 . Thus channel areawas considered of equal importance to incidence angle in designing for operation at mini-mum loss.

A stream-tube analysis was performed to obtain the local value of a/a* through the bladepassage. The calculation used an average passage width obtained by averaging passagewidths of blade channels for streamlines defining the stream-tube annulus, shown byFigure 39. The calculation accounts for convergence, total relative enthalpy change inrotors due to radius changes, and distribution of losses within the channel. The distri-bution of losses within the channel assumed no loss until the first covered section, seeFigure 40. At this point the shock loss described in Appendix 2 was applied. The pro-file loss was linearly applied from the first covered section to the full value at the trailingedge. Thus, at the first covered section the full shock loss was used and at the trailingedge the sum of the shock and profile loss, as shown by Figure 40.

Values of minimum channel a/a* were controlled by adjusting front camber. If a largervalue of a/a* was required the passage was opened by increasing the value of front cam-ber while holding rotor incidence constant at _e B' point, and stator incidence constantat the leading edge, as discussed in Appendix 3. Conversely, if the value of a/a* was toolarge, front camber was reduced. In general, care was taken to design the blade so thatminimum a/a* occurs near the blade entrance region, to obtain gradual diffusion throughthe channel.

, 61

1969021491-072

+1

Z

LE Z TE

O(Z)_n+ 1 + O(Z)_STREAM TUBE AREA - -- ,n x h (Z)

2

._ Figure 39 Channel Area ,

_ 62

1969021491-073

ca

ASSUMED SHOCK LOCATION

n-

MEAN STREAMLINE

lOO I

, _i80 I J

Iso I

= , _40 i _"

_iil,C"

0 J 20 40 60 RO 100

I %CHORD

Figure 40 Loss Distribution through Channel

i

;. 63

1969021491-074

A PPENDIX 5

Deviation System

64

1969021491-075

A PPENDIX 5

Deviation System

Carter's rule plus an experience factor was used to obtain deviation for both the rotor andstator. The form of Carter's rule that was used is given below:

(A_- i)mc_o 1_Carter's rule deviatio,_ =

l m l

mc = 0.92 (a/c) 2 + 0.002 t_'2

a distance to maximum camber point from leading edgec chord

For the rotor the modification to Carter's rule that was used is shown by Figure 41. Datafrom References 1, 2, 3, and 4 is shown in this figure.

For the stator the modification to Carter's rule that was used is shown by Figure 42. Thiscurve is based on data from Reference 5 which evaluated three stators with Mach numbers

and loadings slightly higher than those of the present design.

4

t4

;2

]96902]49]-076

!

I[i L

7-

i i'' 1 '"I

n-i o

° CI =

L \LLI' W _

, _

.. '" C 0_"- _ _

!0 s

o = . o _ ,,u.muJ

; SltidO]O ~ ]1nil S,HIlUVO 9_NOIS]O9 ,

i., I 66

1969021491-077

a.mt.-

o,,"

0I-< ld

o _

it-

_o ® 8 o

_oo-__ 13 _

• 0-_ @

© _ m qr o" u:" LLtU

S33@91Q~ 3111Y&HI.LHV3 _ _ N@ISI@g

I' 67

i

1969021491-078

APPENDIX 6

Aerodynamic Summary

68i

1969021491-079

TABLE 6 Aerodynlmlc Summary - Rotor Inlet

0 10 20 30 40 50 60 70 80 -iffectlveO.O. effective 1.0.

Diame18r inches 16.541 19.086 21.210 23.068 24.743 26.286 27.732 29.105 30.4

meters 0.4201 0.4848 0.5387 0.5859 0.6285 0.6677 0.7044 0.7393 0.7728

reRatio 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0

Temperature Rmo ,;

-:J o

'10

Oiamew Diameter Pressure Rltio Tempenture Ritio M M' V V V' V' Vz Vz Vm Vm Vo Vo

V' 0 v'o

13z 13z (lz

{3' z

f

f

U U

Mm

z 13m 13m (1m

(1m

0

effective 1.0.

inches 19.193 meten 0.4875

2.0 1.250 0.854 0.469

ftJ.ec 995.7 m/sec 303.5 ftJ.ec 546.3 m/.ec 166.5 ftJ.ec 512.3 m/.ec 156.1 ftJ.ec 538.4 m/.ec 164.1 ftJteC 837.6 tn/sec 255.3

ftJ.ec ·92.8

m/~ ·28.3

degrees 58.55 radians 1.022

degrees 10.26

radians 0.179

degrees 17.93 radians 0.313 ftJ.ec 930.3 m/sec 283.6

0.462 0.439

degrees 57.27 radians 0.999

depts 9.n

radilns 0.171

TABLE 7 Aerodynamic Sumnwy • Rotor Exit

10 20 30 40

20.815 12.217 23.530 24.782 0.5287 tl.5c>43 0.59n 0.6295

2.0 2.'0 2.0 2.0

1.230 1.229 1.232 1.235 0.801 0.764 0.735 0.712 0.579 0.620 0.659 0.700 933.6 895.1 865.0 841.2 284.6 272.8 263.7 256.4

674.6 726.1 n5.8 827.7 205.6 221.3 236.5 252.3

590.8 592.0 586.0 579.0

180.1 180.4 178.6 176.5

605:2 598.4 588.0 579.2

184.5 182.4 179.2 176.5

710.9 665.6 634.4 810.0 216.7 ~2.9 193.4 185.9

-298.1 -411.3 -506.1 -591.3

-90.9 ·125.4 -154.3 -100.2

SO.27 48.35 47.27 46.49

0.8n 0.144 0.825 0.811

26.77 34.79 40.82 45.60

0.467 0.607 0.712 0.796

12.55 8.38 4.69 1.28

0.219 0.146 0.US2 0.022 1008.9 1076.9 1140.6 1201.3

307.5 328.2 347.6 366.1

0.519 0.511 0.500 0.490 0.S07 ~ 0.505 0.498 0.490 49.59 48.05 47.18 46.48

0.866 0.839 0.823 0.811

26.22 34.SO 40.72 45.59

0.458 0.602 0.711 0.796

SO 60 70 80 90 100

effective 0.0.

25.987 27.158 28.302 29.427 30.553 31.847

0.6601 0.6898 0.7189 0.7474 O.n60 O.

2.0 2.0 2.0 2.0 2.0 2.0

1.239 1.244 1.251 1.259 1.2n 1.

0.694 0.680 0.670 0.662 0.657 0.651

0.741 0.781 0.819 0.854 0.866 0.742

822.9 809.7 800.7 795.4 795.4 811.1

250.8 246.8 244.0 242.4 242.4 247.2

879.8 929.8 979.1 1025.4 1048.2 924.3

268.2 283.4 298.4 312.8 319.5 281.7

571.9 563.9 556.5 548.0 524.5 379.4

174.3 171.9 169.6 187.0 169.9 115.8

572.2 566.1 562.0 558.4 540.9 395.4

174.4 172.6 171.3 170.2 164.9 120.5

591.5 578.9 570.2 566.4 583.1 708.2

180.3 176.5 173.8 172.6 1n.7 215.9

-688.2 -737.5 -801.7 -860.0 -897.9 .a:s.4

-203.7 -224.8 -244.4 -262.1 -273.7 -254.8

45.96 45.76 45.70 45.95 48.03 81.82

0.802 0.799 0.798 0.802 0.838 1.0'

49.44 52.60 55.24 b'7.SO 59.71

0.883 0.918 0.954 1.004 1.042 1.1

-1.95 -5.04 -8.07 -11.~1 ·14.18 -1

-0.034 -0.088 -0.141 -0.'194 .0.248 -0

1259.7 1316.4 1371.9 1428.4 1481.0 1543.7

384.0 401.3 418.2 434.8 451.4 470.

0.482 0.475 0.470 0.465 0.447 0.318

0.482 0.473 0.465 0.456 0.433 O. 45.95 45.64 45.42 45.41 0.802 0.~7 0.793 0.793

49.43 52.49 54.97 57.0 58.93 64.

0.883 C,.!,16 0959 0.995 1.029 1.1

1969021491-082

?2

i

1969021491-083

A PPENDIX 7

Airfoil Sections on Conical Surfaces

(Multiple-C ircular-Arc Definitions )

73

I

1969021491-084

• • .mo .u+o m q_,t c_c;_c_ ,_o " "

_,, _oSSm |:_ ¢_ooo oo

"/4

I

1969021491-085

,= (_ "_ " _ ' *c; .... o

, i i I:ii! o !

'15

1969021491-086

APPENDIX 8

Rotor Slot Design

j,

, 76

1969021491-087

APPENDIX 8

Rotor Slot Design

A. Aerodynamic Design

A slot was designed for the multiple--circular-arc rotor with the objective of reducingshock losses and improving efficiency. Slot exit flow was used to deflect supersonicflow on the suction surface of the blade, causing an oblique shock upstream of the as-sumed normal shock position of the caannel entrance, as shown in Figure 13. Thisoblique shock has low loss, and reducing the downstream Mach number gives lower

normal shock losses. Recovery of the oblique and normal shock system is higher thanthe recovery of a single normal shock.

Investigations of slot chordwise position showed that placing the intersection of the rearslot wall and the suction surface at the assumed normal shock impingement point gave

optimum results. When the slot was moved forward, the recover_ -_as almost exactlythe same for the oblique and normal shock combination, but flow behind the slot expandedto conform to the blade surface with nearly the original high Mach number. This re-sulted in an area of reduced loss in mid-channel but the original losses on the suction

surface remained. Moving the slot aft reduced the channel area affected by the obliqueshock.

Aerodynamic limits and practical limits of fabrication were determined. Parametricstudies of overall shock recovery were made for slot designs falling within these bound-aries. Results are shown in Figure 43 for the three radial positions defining the slotends shown in Figure 44.

Rotor Slot Design Limits

1) Wedge Angle Limit

The relative Mach number on the suction surface of a blade determines the maximum

wedge angle formed by the slot discharge air for which an attached oblique shock canexist.

2) Compression Limit

At a given radial location, total pressure at the slot exit was assumed to be equal to thetotal pressure at the leading edge, multipled by the recovery of a normal shock at theblade inlet relative Mach number. In order to obtain flow ra_e control, slots were de-signed for choked exit flow which then specified the maximum static pressure at the slotexit for the corresponding total pressure. As the slot angle is increased, the obliqueshock angle is increased with a corresponding rise in the static pressure downstream of

J

77

1969021491-088

78

i

1969021491-089

::ISV:iU3NIA_:]^03_ %

79

i

1969021491-090

8O

i

1969021491-091

34

32

Fi_zre 44 Axial Projection of Slotted Rotor

I

, 81l

1969021491-092

theobliqueshock. The limitingslotangleoccurs when the staticpressure inthe channel

downstream oftheobliqueshockbecomes greaterthanthe staticpressure inthe slot.The assumptionofa choked slotisvalidonlywhen itdischargesintoa regionoflowerstaticpressure. An obliqueshock compression which would unchoke the slotwould makecalcul_tlonsof slotflow,and ofthe obliqueshock system, very complex. This slo_

anglelimitwas foundtobe lessthanthe wedge-anglelimitdescribedinI)and thusre-placesitas an upper limit. R can be foundon Figure4.

i

3) Forward LocationLimit ofObliqueShock!

With a slotexitat theassumed normal shock locationand a givenslotangle,increasing

slotwidthmoves the startingpointoftheobliqueshock towards the leRdin_;edge, sop

thatitcovers more of thechannel. Finally,the obliqueshock justreachesthe leadingedge ofthe adjacentrotorbladeso thatthe entirechannelisaffectedby theobliq4J,vshock.Any furtherincreaseinslotwidthwillcause the obliqueshock to extendinfrontoftheleadingedge ofthe adjacentbladeand change the entireinlet-flowpattern. This limitisshown inFigure 43.

4) SlotLength-to-WidthRatioLimit

For a givenradialstationand slotangle,increasingslotwldtbdecreasesthe ratioof

I slot length to width. For flow guidance capabilities it was decided that the ratio of slott' lengthto widthshouldbe greaterthanthree. Any valuelessthanthisresultsina poorly-

definedthroatwhere flowcontrolisdoubtful.The slotlengthwas measured d_rectlyfrom drawings,_akingintoaccountthe roundingoffofsharp edges. These limitsareshown in Figure43.

• 5) Channel Choking Limit

For a givenradialstationthe channelbetween two rotorbladeshas a givenaverage Machnumber and thusan a/a* associatedwithit. But when slotflowisadded tothe channel

flow,the referencea* willincreaseand a/a* willdecreasetowards one where choking

may occur. This isIx-'tialiyoffsetby the increasedrecovery due totheobliqueshock,_ system. R is desired to keep a/a* from decreasing more tbJn two percent from the

value of 1.03 originally in the channel. Thus:i

i a/a* original - =1.02[ a/a* slot + originalf

This ratio is calculated u_,_ the relattonsh|p:

W slot + original rec°verYori_ina[ a/a* originalX

W°rlg lnal rec°verYoriginal _- slot a/a* slot + original

, 82IB

1969021491-093

!I[

i Thus percent flow rates of the slot are limited by channel choke considerations. Limits

i are shown in Figure 43.I

I 6) Geometric Limits]

[ (a) Slot Intersecting Rotor Blade Leading Edge|

For highly cambered blades and for slots near the leading edge, the slot entrance can cutfoff the leading edge of the blade. These limits are shown in Figure 43.

' (b) Machinable Slot Angle

Slot angles less than 5 degrees with the tangent to the section surface would compromise

the dimensional accuracy of the slot because of machining tolerances. Fabrication limitsare shown in Figure 43.

The shaded area in Figure 43 represents the feasible slot configurations at a particularradius.

Calculation of Relative Mach Number at Shock Impingement Point

Calculations made in the parametric study of overall shock recovery depended on thechannel Mach number distribution between the shock impingement point and the leadingedge of the adjacent blade. The procedure used in determining this distribution was asfollows.

Radial distributions of relative Mach number at the rotor leading edge and at the loci of

shock impingement points on the blade suction surface were calculated using a stream-line analysis computer program. The calculation was made for free flow, with no blade

fleets. Resulting free-stream Mach numbers at the shock impingement points werecorrected for blade blockage and boundary layer effects which increase the axial velocityand decrease the relative air angle. The calculation method is shown below.

_Shcor = /_Shfree flow - ia'

where lay",the blockage correction angle, iz discussed in Appendix 3.

83

I

1969021491-094

or

A t

_f.wM Shcor

free free flowflow

T _ v

_e----_-- U/A t ------_

At =_ speed of sound at absolute stagnation temperature

Ratios of absolute velocity to the speed of sound at absolute stagnation temperature wereused to determine corrected relative Mach numbers corresponding to adjusted axial ve-lo_'.'ties, as follows:

Msh free flow = M'sh free flow cos _ sh free flow

ee flow = Msh free flow from tables

V(_-'t )c ¢ Vm ) ( tanB sh free flOw )or = _ free flow tan_sh cor

Msh found from tables corresponding to(V'm /

cor \At /cor

Msh eorMv =Shco r

cos B sh cor

This procedure uses the constant At based on absolute stagnation temperature as a ref-erence for all geometrical corrections. The resulting relative Mach number M'sh coris the value at the shock impingement point corrected for blockage ant] boundary-layereffects. This corrected free streamline 18 not parallel to the suction surface, so theactual blade-surface relative Mach number was found as follows:

84

I

1969021491-095

Relative air angles which had been adjusted for blockage were subtracted from actualbJade surface angles. Differences were added to the hodograph vector angles of adjust-ed Mach numbers to obtain the Mach number vector angles of flow aligned to the bladesurface. An example is given below:

Osh = -_ - 8*A

B_h cor sh

_* is the metal angle of the blade suction surface at the shock impingement point.sh

is the hodograph vector angle defined by:

_ tan-l_/ k -1 2=_-_- -_-_-(M -1) - tan-l_-i

M'shss, the relative Mach number on the suction surface, is found from the hodographvector angle, _final" The Mach number distribution immediately upstream of the shock

was assumed to b_ linear between M'sh_s and the relative Mach number at the leadingedge, These Mach numbers were uses m the parametric studies of overall shock re-covery and in determining aerodynamic design limits.

Recover_ Calcula£iorm

The channel and slot were broken up into regions, as shown in the sketch. The Machnumber distribution across the channel was divided into Mxa, Mxb_ and Mxav for usein the various regions, Boundary conditions for flow downstream of intersecting shocksrequire that adjacent streams have the same static pressure and flow direction. These

rules were not strictly followed for the intersection of the normal shock and obliqueshock at the channel entrance, since downstream flow is subsonic and since static pres-sure is not constant across a gap. Instead, a normal shock position was assumed, asshown in the sketch, and shock losses were calculated for components of Mach numbernormal to the direction of this shock. Calculations pertaining to oblique shocks upstream

_mm

b 6

85

I

1969021491-096

of the normal shock satisfy the usual boundary conditions. Sonic flow emerging from thethe slot expands to the static pressure determined by the oblique it causes. A methodof characteristics solution of this flow showed that the outer boundary of the flow is on-

ly slightly irregular, and can be considered to be parallel to the rear slot wall, asshown in the sketch. As a result, the slot angle was used as the wedge angle for oblique

M=I 57

N SHOCK M__

52.5____ DIMENSIONS IN INCHESFLOW

_" "--.. BOUNOARY'_

""" I

10° 0.010 RMMA X = 1.15,.,.=

shock calculations. The procedure is as follows:

1) h, slot or wedge angle £ is chosen. Using MXa and this wedge angle, the shock an-gle v is determined from Keenan and Kaye tables.

2) The Mach number normal to the oblique shock is determined from IMxa)n =Mxasin v.

This determines Py/Px from normal shock tables.

: 3) Downstream of the oblique shock, My is found from Keenan and Kaye tables, using£ and v from 1).

4) My n, the normal component of My to the normal shock, is found from

My n = My sin (90 - £). This determines P5/'Py from normal shock tables.

5) Total :recovery across the oblique and normal shock is found from

PS/P x = (PS/Py)(Py/Px)

6) Mach number is found in the slot in the same manner as for we dge angle determin-ation:

a) py/Py found from isentropic tables for Myb) Py/Px found in Step 2c) P6/Px found from normal shock tables for Mxb

d) Psi/Psi = Py/P6 = (Py/Py)(Py/Px)(Px/P6)

86

(

I

1969021491-097

7) Msl is found from isentropic tables for this pressure ratio. P4/Psl is found from

normal shock tables for Msl n sin (90 - fi )t

' 8) Total recovery from Region X to Region 4 across the slot region is!

PI/Px = (Pl/Psl)(P6/Px)

L 9) The recovery across the normal shock with no oblique shock present, correspond_Ji to Mxav. This is the base recovery from which all increases are measured.

10) The area weight factors for Regions 4, 5, and 6 are found by measuring the widthsj of the regions intersected by the shock pattern, and dividing them by the channelI width. Thus the total recovery in the channel with the oblique is found by adding

the products of the recoveries and their respective area weight factors.{

[

11) Recovery increase is found by subtracting the P6/P x recovery corresponding tono oblique shock present, from the new recovery calculated in Step 10.

I

' Efficiency Calculations

IAdiabatic efficiency is given by k-__!l

[T2/T1][ total] k - 1=

T2/T 1 - 1

(1 - R profile) + (1 - R shock) = 1 - R total

Rprofile is constant but the slot improves Rshock

Efficiency was calculated for slotted and unslotted blade sections, using the correspond-

ing Rtotal values. Figure 45 is a plot of estimated rotor efficiency with and withoutslots, versus span.

87

I

1969021491-098

THE DIAMETERS OF THESE STATIONSCORRESPOND TO STREAMLINE

96 LOCATION AT ROTOR INLET OF

28, 30, 31 5 INCHES DIAMETER,

__ C_ RESPECTIVELY

_= j WITH SLOT

_ 88

M.I WITHOUT SLOT _ Huu 84

k-

n,, 800i--0

76

72

16 18 20 22 24 26 28 30 32 34

ROTOREXIT DIAMETER_- INCHESI I I I I I

ROOT 20% 40°/= 60% 80% TIP

Figure 45 Rotor Efficiency vs Potor Exit Diameter

B. Mechanical Design

Vibration Frequency

The coordinates of the slotted blade sections were used in calculating the blade first-mode vibration frequency of 500 cps at 110 percent of design speed. This is shown inFigure 46 to give adequate margin above the first-mode resonance caused by two exci-tations per revolution. Second-mode resonance frequency is 1030 cps.

A bending frequency of 842 cps was calculg_ed for the rear outboard blade cantile,ered

_t a radius of 14 inches where the slot begins, assuming no intermediate or tip bridgesexist. A bending frequency of 1085 cps was calculated for the front outboard blade can-tilevered at a radius of 14 inches. Neither frequency will resonate with 2E or 3E exci-tations per revolution.

Blade Steady-State Stress

At 110 percent of design speed maximum blade stress due to centrifugal and untwisteffects occurs at the root leading edge of the blade and is 85,264 psi, which is below themaximum allowable of 116,000 psi for titanium (AMS 4929). This stress is lower thanthe MCA unslotted blade maximum steady-state stress of 89,400 psi at 110 percent ofdesign speed. The results are summarized below.

J

88/

1969021491-099

I t.=(_) REAR TIP

I TANDEMI SECTION

' 7 II REARTANDEM _ /

SECTION,:1STCANTILEVER MODE

4 TANDEM BLADE- DISK

ATTACHMENT EFFECT

INCLUDED

X

2

1 100%DESIGNSPEED!I

0 2 4 6 8 10 12 14

ROTOR SPEED "-.,RPM X 10 .3

Figure 46 Rotor Blade Resonance Diagram

Allowable Stress

Location Calculated Stress_ psi at 100°Ft 9si

blade root P/A 34,843 69,000

blade root L/E, P/A untwist, 85,457 116,000 (0.2% yield)

tensile

blade 10% from root, concave 76,604 116,000 (0.2% yield)

surface, max. thickness, P/A and

untwist, tensile

89

! , i J

1969021491-100

APPENDIX 9

Tandem Rotor Design

9O

,i

1969021491-101

APPENDIX 9

:' Tandem Rotor Design

A. Aerodynamic Design

A tandem rotor was designed with the objectives of 1) minimizing shock losses by con-touring the supersonic section of the suction surface to minimize supersonic accelera-

', tion, and 2) minimizing the effects of shock boundary-layer interactions, by initiatingthe diffusion process on a subsonic portion of the blade, where the boundary layer is no

f longer affected by the passage shock shown in Figure 14. Tandem blade sections weredesigned by shifting the portion of the multiple-circular-arc blade downstream of the

I normal shock impingement point, tangentially in the direction of rotation. In this wayi blade leadivg and trailing edge planes were kept the same as for the MCA blade and sur-

face contours were kept the same except in the immediate regior, of the normal shock.: Other criteria used in the tandem design are as follows:

i_ 1) The tandem shape was required only from a radius of 10.4 inches to the tip, sincel at that radius the local Mach number was less than 1.35, below which value boundary-t layer separation was unlikely (Reference 9).

2) Near the hub the trailing-edge location of the forward section was placed downstreamof the shock impingement to provide for possible boundary-layer reattachment. At thetip where the shock is very strong and there is little chance of reattachment, the front-section trailing edge was located immediately behind the shock-impingement point. Trail-ing edges of intermediate front sections follow a straight line on a projected view of theblade (Figure 47).

3) The slot width between the two tandem blade sections was made approximately 10times the predicted total boundary-layer momentum thickness buildup on the pressureand suction surfaces at the assumed normal shock location. The slot length-to-widthratio was made equal to or greater than three whenever possible for directional guidanceand flow-rate control.

4) Rear-section leading edge and front-section trailing edge radii were determined from

[ results of a vibration stress test on various tandem-shaped test specimens.If 5) Bridge width and bIend radii were determined as a result of a stress study done on

I the proposed tandem design.

[ 6) The stacking line was determined as a result of tilt stress considerations on the blade.!

i 7) Smooth surface contours were maintained on front and rear sections.

91

1969021491-102

34

32

I

28 l_"

t_w 26Zt.1_z

0 24 ,

22

2O

.2

16

Z AXIS

Figure 47 Axial Projaction of Tandem Rotor

' 92

1969021491-103

B. Mechanical Design_

.Bridge Fatigue Specimen Tes_

Fatigue tests were run on sample tandem blades having bridges at mid-span betweenfront and rear blades. Relative fatigue strengths of four bridge conftguratio_ls were

', determined. A schematic drawing of a test specimen clamped to a shaker is shownrin the sketch, together with sections of four bridge configurations. Specimen No. 1

L was the basic configuration, No. 2 had increased overlap between front and rear tan-I

i dem blades, No. 3 had the bridge extended outside the channel, and No. 4 had !n-L creased radii on blade edges intersecting the bridge. Relative strengths were judged

[h

L' NO. 1 I//

///////////X/////A

, _ IINO. 2

I

I NO. 3

• _2"- L.,_ANO. 4

SECTION AA - 4 CONFIGURATIONS

ion the basis of the vibratory stress measured by strain gages at maximum thicknessat the radial height of the bridge. Stresse_ were not measured at failure points be-cause slight variations in gage position in areas of high stress concentration pro-duced scattered results.

All failures occurred at the airfoil bridge junction. Relative stresses for fatiguefailures at approximately 107 cycles are tabulated below:

Stress or Reference Gage

Configuration for Failure at 107 Cycles, psi

1 3_ 5002 10, 000

3 7,5004 15, 000

From these test_ it was determined that large blade-edge radii as in ConfigurationNo. 4, and long blade overlap as in Configuration No. 2, were effective methods of

, obtaining fatigue strength. Both of these featuLres were incorporated in the tandemblade design.

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1969021491-104

Vibration Frequency

Tandem blade section coordinates were used to determine airfoil section properties.These properties were then used to calculate first--mode vibration frequency _,f the blade.This f_equency was adjusted to include the effect of disc attachment flexibility. At 110percent of design speed the blade-alone frequency was 490.0 cps which is shown in Fig-ure 46 to give adequate margin above the first-mode response caused by two excitationsper revolution.

A bending frequency of 745 cps was calculated for the rearward outboard section canti-levered from a radius of 14 inches where a bridge is located (Figure 47). This frequencyis lower than that of the front section cantilevered at the same radius. The front section

cantilevered at a radius of 10.4 inches where the tandem sections start has a bendingfrequency of 418 cps, close to the 2E value of 408 cps at 110 percent of design speed.The rear section cantileverea at a radius of 10.4 inches has a bending irequeney of 645cps, which is well above the 3E value of 608 cps at 110 percent of design speed. Thesevalues are shown in Figure 46. The assumption of cantilevered blades gives very con-serva*_ve results, since the blades are actually fixed at more than one position, whichwill tend to increase the natural bending frequency. It was concluded tl_t the isolatedt_udem-section frequencies are less critical than the full=blade system.

Blade Steady-State Stress

Combined centrifugal pull and untwisting stresses were calculated at 110 percent of de-sign speed. A comparison was made to allowable s_resses based on the "lse of AMS 4928titanium alloy at 150°F (338°K).

The maximum local stress occurs at the root leading edge of the elude, which from pactexperience is not a problem. The combined stress at 10 percent. _pan at the maximumthickness point was calculated to be 76, 61_ psi comared to 74,000 psi for the MCA bladeat 110 percent of design speed, well below the 0.2 percent yield strength of 116,000 psifor titanium. The results are summarized below.

At 110 Percent of Design SpeedAllowable Stress

Location Calculated Stress_ psi at 100°F, psi

blade root P/A, tensile 41,517 69,000

blade root L/E, P/A 85,457 1i_,000 (0.2% yield)unt,,,__., tensile

blade 10% out from root, 76,604 116,000 (0.2% yield)concave surface, max.thickness, P/A + untwist,tensile

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1969021491-105

Bri.dge Steady-State Stresss

Centrifugal loads, untwist torques, and bending moments due to gas loads and nonradial

stacking were used to calculate shear and bending stresses on the three bridge sections.Thre_ types of stress calculations were made.

"£he first type of calculation was based on the assumption that the bridges equalize theradial elongations of the front and rear spanwise lengths between bridges by absorbing

the difference between centrifagal loads that would occur if front and rear were unre-strained and the loads that wilt occur when the bridges restrain them. The untwist programfor the tandem-blade system gave the qualitative result that centrifugal loads constituted

the major stress for all tandem sections, and was constant for a cross-section at a given

radius. Since elongation is directly proportional to stress, equal elongation of front andrear t3ndem sections was assumed in calculating loads on a spanwise length between bridgesin the restrained condition. Centrifugal loads were determined for the individual frontand rear tandem sections, and the difference between the restrained and unrestrained

centrifugal loads was used