The mechanical design of a two-stage impulse turbine for the liquid hydrogen turbopump of the M-1 engine

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THE MECHANICAL DESIGN OF A TWO-STAGE

IMPULSE TURBINE FOR THE LIQUID HYDROGEN TURBOPUMP

OF THE M-1 ENGINE

GPO PRICE S

B Y CFSTI PRICE(S) S

T . W. Reynolds

Hard copy (HC) z m

(I Microfiche (M F) W 853 July 85 c

Prepared for

National Aeronautics and Space Administration

Contract NAS 3-2555

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A E R O J E T - G E N E R A 1 C 0 R P O R A T I O N

S A C R A M E N T O , C A L I F O R N I A NOTICE This report was prepared as an account of Government sponsored work. Neither the United States, nor the National Aeronautics (NASA), nor any person acting on and Space Administration behalf of NASA: Makes any warranty or representation, expressed or irnplinrl, with r r c p n r t tc thn r i ~ ~ u ~ r ~ ~ y , ~ ~ ~ m l n t m - - = c r . - .

-- i or usefulness of the information contained i n t h i s report, or that the use o f any information, apparatus, method or process disclosed i n thi-s report may not infringe privately owned rights, or Assumes any l i a b i l i t i e s w i t h respect to the use of, or for damages resulting from t h e , u s e o f any infor- mation, apparatus, method or process disclosed in this report.

A s used above, “person acting on behalf o f NASA” includes any employee or contractor o f NASA, or employee o f such con- tractor, to the extent that s u i h employee or contractor of NASA, or employee of such contractor prepares, disseminates, or provides access to, any information pursuant t o h i s employment or contract with NASA, or h i s employment w i t h such contractor.

Requests for copies o f this report should b e referred to: National Aeronautics and Space Admini stration Office of Scientific and Technical Information Attention: AFSS-A Washington, D. C. 20546 NASA CB 54821 AGC 8800-58 TECHNOIXXiY REPORT e THE MECHANICAL DESIGN OF A TWO-STAGE IMPULSE TURBINE FOR THE LIQUID HYDROGE2I TURBOPUMP OF THE M - 1 ENGINE Prepared f o r NATIONAL AERONAUTICS AND SPACE ADMINISTBATION CONTRACT NAS 3-2555 T e c h n i c a l M a n a g e m e n t : Prepared by: NASA LEWIS RESEARCH CENTEB AEROJELI-GENEBAL CORPORATION CLEVELAND, OHIO L I Q U I D ROCKGT OPEBATIONS SACRAMENTO, CALIFORNIA TECHNICAL MANAGER: D. D. Schter AUTHOR: T . W . Reynolds APPROVED: W e W . Uilcox APPROVED: W . E. Watters M-1 Pro ject Manager Manager M-1 T u r b o p u m p Project ABSTRACT of t h e This report describes the mechanical and s t r u c t u r a l design two-row Curtis-turbine f o r t h e hydrogen turbopump of t h e H-1 &&e. The turbine is a lightweight d i r e c t drive Unit with a design point power out- put of 88,150 hp at 13,225 rpm, and a t h e m 1 efficiency of 6 2 . 8 % . Oper- a t i n g temperature gradients from -42O0F t o +1200°F are accommodated by f a b r i c a t i n g t h e housings from t h i n , welded s h e l b of Inconel 718 a l l o y and supporting t h e homing8 near the axial center point of the blade rows by an e x t e r n a l frame, iii TABLE OF COIOTZZUTS pagd

I. s u m

I1 . 1N"RODUCTION

I11 . TEXXNICAL DISCUSSION

A. DEjIGN REQUI-s 1. Performance 2. Interfaces a. Internal Turbopump Interfaces b. External Interfaces 3. Primary bchanical hsijm Objectives B. SELBCTED CONF'IOURATION 1. Description Hethod for AchievinR kchanical Design Objectives 2.

3. Basis for Configuration Selection C. MATERIALSELEZCION 1 . Haterial Requirements 2. Selected Materials D. STRUC!WRAL CRITERIA 1. Static or Non-Rotating Components 2. Rotating Components 3. Vibration and Acceleration Loa& E . COl4K)NE"l! STBBS SUHMRY iv TABLE O F CONTENTS (Cont'd)

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Turbine I n l e t Manifold First-Stage and Second-Stage Rotors a. First-Stage Rotor S t r e s s b. First-Stage Rotor Vibration ZL /" C. second-atage K O t O r i h r e s s d. Second-Stage Rotor Vibration Reversing Row a. S t r e s s 4 1 b . Vibration Rotor Tie-Bolt BearinK Housing S e a l Turbine Exhaust Housings a. Exhaust Cone be Dual Exit Exhaust Cone Support Frame and Turbine Main-Flange Clampe DYNAMIC BALANCING AND ASSEMBLY TECHNIQUES F .

1. Dynamic Balancing 2 . Assembly Techniques G. DESCRIPTION O F COMPONENT FABRICATION 5 1 1. Turbine I n l e t Manifold 2 . First-Stage Rotor and Blades V

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a. Machining Process Techniquea 52

b. Manufacturing Problem Area 54

Second-Stage Rotor and Blades 54 3.

I " Reversing Vane13 55 4 .

Rotor Tie-Bolt 5.

Bearing Housing Seal 56 6 , Exhaust C a n e 57 7.

Exit & h U 8 t H O U 8 W 57 8.

Support Frame and Hain Fl-e Clamps 58

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I V . CONCLUSIONS AHD RFZOMIPIIIATIONS 59 BIBLIOGRAPElY LIST OF TABLES T i t l e Pagt No .

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F i r s t Stage Disc S t r e s s , 10 t o 2% stc 1, LIST O F F I G U R E S T i t l e

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H-1 Ehgine Mockup 1, Turbine Components as Assembled 2 .

First-Stage &tor and Blades 3 .

First-Stage Rotor - Front Side

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First-Stage Rotor - A f t Side

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1 1 Second-Stage Rotor and Blades 6.

Secondstage Rotor - Front Side

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Second-Stage Rotor - A f t Side

8 , Turbine Support Frame 9.

Torus and Nozzle Assembly 10, Nozzle Profile and Support Rings 11.

Reversing Bow, Section and Vane P r o f i l e 12.

Turbine Exhaust Cone Dm1 Exit Exhaust Housing 14.

Gas Temperature Transient at S t a r t u p I n l e t Manifold S t r e s s , S t e a d y a t a t e 16.

I n l e t Manifold S t r e s s , Maximum Thermal Gradient 18. I n l e t Manifold Streas, Proof-Pressure T e s t v i i LIST O F FIGURES (Cont'd) Title

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Inlet Manifold Stress, Line and Shock Loads 31 First-Stage Blade Stress First-Stage Blade Temperature Gradients 21 e First-Stage Blade Vibration Frequencies 22 e Second-Stage Disc Stress Second-Stage Blade Stress Second-Stage Blade Vibration Frequencies Reversing Row Stress I 26 Reversing Row Vibration Frequencies Turbine Rotor, Tie-Bolt and Adjacent Parts 28.

Bearing Housing Seal Assembly v i i i I . SUMHABY The turbine is designed as a flightweight, direct-drive f o r the M-1 l i q u i d It is a f u l l admission, two-row C u r t i s turbine, which is the type hydrogen pump.

t h a t is used extensively i n high pressure steam turbines and Large rocket engine turbOpmpS The The aerodynamic design point power output is 88,150 hp at U,225 rpm.

driving gas has a temperature of 1460"R and a pressure of lo00 psia. T h i s driving gas is obtained from t h e combustion products of hydrogen and oxygen.

The turbine components are fabricated almost e n t i r e l y of Inconel 718, a nickel-chromium alloy. This material possesses many desirable q u a l i t i e s , such a8 high strength at both cryogenic and elevated temperatures, high d u c t i l i t y , good stress rupture q u a l i t i e s , and good weldability using t h e gae tungsten arc weld o r t h e electron-beam methods.

Thermal gradients a r e severe because the turbine soaks a t l i q u i d hydrogen temperature (-423*F) and is then shocked with hot gas at lOOO*F t o 1jXX)OF. The turbine s t a t i c components are a l l supported from a main flange by a space frame t o permit unrestrained thermal movement and is free t o move r a d i a l l y at the turbine end, while it is firmly anchored t o t h e hydrogen pump discharge housing. The arrangement permits the housings t o move f r e e l y with t h e thermal gradients, with a minor a f f e c t upon blade row alignment and clearances. A f l e x i b l e , bellows- only type seal is provided between the f l o a t i n g housings and the fixed turbine r o t o r bearing housing.

The s t a t i c components were designed with a s a f e t y f a c t o r of 1.2 upon t h e 0.2% y i e l d strength, or a s a f e t y factor of 1.6 upon the ultimate strength, which- The turbine blades were designed with a ever is lower at the l o c a l temperature.

1.25 at 110% of operating speed. S t r e s s limits w e r e determined s a f e t y f a c t o r of by p l o t t i n g combined s t r e s s e s on a modified Goodman diagram. The djagram f a i l u r e The average e f f e c t i v e stress in t h e r o t o r discs w a s line w a s used as a l i m i t .

The r o t o r bore stress l i m i t e d by a 1 . 5 s a f e t y f a c t o r upon t h e ultimate strength.

was limited t o less than twice the yield strength at the l o c a l temperature.

This turbine o f f e r s good performance, light weight, and simple mechanical construction. The t o t a l turbine weight, including the support frame and the i n l e t - adapter-elbow, is 960 lb.

11. INTRODUCTION The turbine specified f o r t h e M-1 l i q u i d hydrogen turbopump is a direct- drive, f u l l admission, two-row Curtis-type (velocity staged) machine. T h i s turbine has not been t e s t e d at f u l l s c a l e under hot conditions. However, a single stage t u r b i n e w i t h some components i d e n t i c a l t o the two-row machine w a s tested driving Page 1 t h e hydrogen purnp‘l’to approximately 93% o e s i p speed. The single stage turbine t e s t r e s u l t s are being reported separatelyS2P .

“he pumping system f o r the l i q u i d propellants of the PI-1 engine consists of The turbines are driven with gas-generator-supplied hot two separate turbopumps.

gas obtained from the combustion products of the l i q u i d hydrogen and the l i q u i d The turbines are arranged i n series, with the gas i n i t i a l l y expanded i n the oxygen.

l i q u i d hydrogen turbopump turbine and then, further expanded i n the l i q u i d oxygen turbopump turbine. The exhaust gas from the l i q u i d oxygen turbopump turbine is routed through three heat exchangers; one t o heat hydrogen f o r t h e gimbal actuators, one t o heat hydrogen for tank pressurization, and the other t o heat oxygen f o r tank ?

pressurization. The exhaust gas is then directed t o the lower skirt of the engine Finally, the gas is ejected through a s e t of t h r u s t chamber f o r cooling t h e walls.

small nozzles t o provide thrust augmentation with an approximate specific impulse Figure No. 1 show6 the M-l engine turbopump and gas ducting of 260 lbf-sec/lbm.

arrangement .

The turbine components a r e fabricated almost e n t i r e l y of Inconel 718 , which is

This material was selected because it is an age-hardenable nickel-chromium alloy.

very ductile, it can be e a s i l y welded by the gas tungsten a r c weld or t h e electron- beam methods, and it has high strength a t both cryogenic and elevated temperature (-420’F t o +130O0F).

The design, development, and fabrication of t h i s turbine were conducted by the Aerojet-General Corporation under contract with the National Aeronautics and Space Administration. The aerodynamic design of the turbine is being reported separately(3’ 111. TECHNICAL DISCUSSION A. DESIGN REQUIREMENTS

1 . Performance

The turbine gitator and rotor gas passages were designed t o meet the following aerodynamic c r i t e r i a : I n l e t Gas Total Temperature 1460’F ‘“Regan, P. J., Mechanical Desiffn of the M-1 Axial Flow Liquid Hydrogen Fuel Pump, NASA CR-54823, 15 February 1966

(2)Blakis, R., Iiindley, B. K. Ritter, J. A . , and Watters, W. E. , Initial Test

Evaluation of t h e M-1 E q u i d Hydrogen Turbopump, Including I n s t a l l a t i o n T e s t Procedures and Test Results, N A S A CR-54827, 20 July 1966 I - \ ‘3’Reynolds, T. W., Aerodynamic Design and Estimated Performance of a Two-Stage Turbine f o r the M - 1 Fuel Turbopump, NASA CR-54820, 15 April 1966 Page 2 lo00 psia I n l e t G a s Total Pressure 1134 O R E x i t Gas Total Temperature 214 psia Exit Gas S t a t i c Pressure Shaft Speed 138225 rpm Shaft Power 88,150 hp Shaft Torque 3!j9800 f t lb Gas Flow Rate 99.3 lb/sec Blade/Jet Speed Ratio (U/Co) 0.188 Pressure Ratio ( I n l e t T o t a l t o Exit S t a t i c ) 4.673 23.0-in .

Turbine Bucket Pitch Line Diameter 0.628 Efficiency ( I n l e t Total t o Exit S t a t i c ) The turbine parts were designed t o meet the following mechanical design c r i t e r i a : * 1 3 0 0 0 ~ (1760OR) Gas I n l e t Total Temperature *1175 psia Gas I n l e t Total Pressure l W 0 F (1460OR) Gas Fkit Total Temperature 386 psia Gas Exit S t a t i c Pressure S h a f t Power 119,725 hp Shaft Speed 14,550 rpm S h a f t Torque 43,200 f t l b Run Duration 400 sec Usable Life 10,500 sec S t a r t Cycles 2 1 O-l3,225 rpm, 1.9 sec Starting Acceleration I Page 4 The turbine r o t o r s t o be close-coupled t o the turbine bearing c s t o achieve s h a f t c r i t i c a l speed requirements, The turbine r o t o r weight t o be minimized t o achieve both do s h a f t c r i t i c a l speed requirements and r o t a t i n g part i n e r t i a requirements f o r engine s t a r t i n g .

e. The flightweight turbine weight goal is t o be achieved.

B . SELECTED CONFIGURATION 1. Description The s e l e c t e d configuration is shown i n Figure No. 2. It c o n s i s t s of a two-stage, velocity-compounded (Curtis-type) turbine s e t with a t o r o i d a l inlet manifold; a segmented conical support structure; a conical exhaust housing (turbo- has a bifurcated hemispherical, exhaust pump test u n i t , engine turbopump design seal-welded separable turbine housing j o i n t s .

housing) ; and unique The f i r s t - s t a g e turbine r o t o r is an i n t e g r a l u n i t consisting of a shaft, turbine d i s c o and 80 blades (see Figures No. 3 through No. 5 ) . The s h a f t connects the turbine t o t h e f u e l pump rotor. P i l o t i n g diameters and a s p l i n e a r e provided f o r alignment and power transmission, The turbine s h a f t a l s o c a r r i e s the turbine end bearing spacers and t h e other r o t a t i n g components needed t o permit t h e functioning of a shaft r i d i n g s e a l and a l i f t - o f f seal. These l a t t e r components are parts of t h e power transmission assembly.

The turbine blades a r e shrouded t o provide b e t t e r performance.

The shroud is s p l i t between a l t e r n a t e blades f o r minimum thermal stress and satis- f a c t o r y c o n t r o l of blade vibration, The blades a r e hollow f o r minimum weight and s t r e s s .

The turbine disc is of minimum cross-section t o reduce weight.

The d i s c has a curvic coupling at the center on the s i d e opposite from the shaft.

This coupling is t h e i n t e r f a c e f o r the second-stage turbine.

The second-stage r o t o r is an i n t e g r a l u n i t consisting of a d i s c and 78 blades, which have a t i p shroud, This shroud is s p l i t between each blade, which is hollow f o r minimum weight. The turbine disc is a l s o designed f o r minimum weight. A curvic coupling is provided as an i n t e r f a c e t o t h e f i r s t - s t a g e turbine, Figures No, 6 through No. 8 show the r o t o r and blades.

A turbine tie-bolt fastens the turbine assembly t o t h e pump rotor.

T h i s t i e - b o l t , which is threaded i n t o the pump r o t o r at one end, has a nut and lock washer r e t a i n i n g system that bears against the second-stage turbine.

The turbine housings, including the i n l e t manifold, a r e mounted Page 6 Component Temperature at S t a r t -420'F ( * The original mechanical design requirements were 1250 p s i a i n l e t t o t a l pressure and lj5O'F i n l e t t o t a l temperature. The lower design values presented were deter- mined t o be compatible with t h e engine requirements and permitted t h e use of Inconel 718 forgings which were s p e c i f i e d t o be solution-annealed a t 1 9 5 0 ' F and An 18000~ solution anneal is required t o obtain adequate aged a t 1 2 0 0 ' F t o 1350°F, stress rupture properties f o r t h e higher design values o r i g i n a l l y specified.

2. Interfaces a. I n t e r n a l Turbopump Interfaces The turbine s h a f t has a s p l i n e and two p i l o t i n g diameters, The turbine s h a f t a l s o c a r r i e s t h e which form the i n t e r f a c e with t h e pump r o t o r , journals f o r t h e turbine-end r o l l e r bearing and the shaft r i d i n g seal as well as To achieve predictable and acceptable turbo- t h e mating face f o r the shaft seal.

pump c r i t i c a l speeds, an interference f i t between the pwnp r o t o r and t h e turbine s h a f t w a s specified f o r a l l operating conditions.

b. External I n t e r f a c e s The following gas duct i n t e r f a c e s were specified with t h e i n t e r f a c e location shown i n Figure No, 1, A s i n g l e 8-ino inner diameter i n l e t f o r turbine drive (1) gas (2) A s i n g l e 5-in. inner diameter i n l e t f o r turbine bypass gas f o r eqgine c a l i b r a t i o n and propellant u t i l i z a t i o n control.

(3) Dual 11.4-in, inner diameter o u t l e t s f o r turbine ex-

haus t gas -

3. Primary Mechanical Design Objectives The s t a t i c and r o t a t i n g parts t o be designed t o withstand a.

severe thermal gradients, The thermal gradients received prime consideration because t h e turbine is exposed t o t h e e f f e c t s of very long pump chilldowns t o l i q u i d hydrogen temperature followed by r a p i d s t a r t u p with 1300'F hot gas. This causes the hot gas flow path t o have very l a r g e t r a n s i e n t temperatures. Some parte, such as the bearing housing and the r o t o r hub, are adjacent t o l i q u i d hy- drogen which causes a large thermal gradient throughout t h e e n t i r e operating cycle.

Adequate s t a t i c - t o - r o t a t i n g part clearance t o be provided b .

under a l l operating conditions while maintaining reasonable close tolerances t o minimize leakage effects.

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F i p r e 2 Turbine Components as Assembled Page 7 U ' w , - I , w /

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Figure 3 First-Ltage dotor and Blades Page 8 Figure 5

First-Stage Rotor - A f t Side

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Figure 6 Second-Stage Rotor and Blades Page 11 Figure 7

Second-Stage Rotor - Front Side

Page 12 t o t h e pump discharge housing by means of the space frame shown i n Figure No. 2, This frame consists of t h r e e equally spaced conical s e c t i o n s 9 t o provide openings f o r the i n l e t manifold and instrumentation l i n e s . A support frame section is shown i n Figure No, 9. The housings a r e supported at a s i n g l e main flange, which is positioned near t h e axial center of t h e blades rowso This location allows thermal movements i n a l l directions with the least a f f e c t upon a x i a l clearances between t h e r o t a t i n g and t h e s t a t i o n a r y members, The turbine housing j o i n t con- sists of lightweight flanges secured by clamps, Sealing is accomplished by applying a small weld bead t o t h e flange j o i n t before t h e clamps are i n s t a l l e d .

The i n l e t manifold consists of a constant cross-sectional area t o r u s having a nozzle r i n g in one s i d e as shown in Figure No, 10. The nozzle r i n g has 37 vanes, which can be seen in Figure No, 11, A flow deflector consisting of a f l a t p l a t e shaped t o f i t the i n l e t is i n s e r t e d i n t h e i n l e t neck t o reduce gas turbulence around the protruding outer nozzle shroud. In general, t h e parts are fabricated from sheetmetal f o r minimum weight and t o allow thermal movement re- s u l t i n g from temperature changes, The reversing row has 67 vanes and is i n s t a l l e d between the first-stage and second-stage r o t o ~ s , This reversing row is made up of s i x c i r - cular segments t o permit u n r e s t r i c t e d movement under the varied thermal conditions The vane segments are of sheet- encountered during pump chilldown and operation, metal construction, The reversing vanes a r e shown i n Figure No. 12. The seg- mented r i n g is mounted on its outer diameter by clamping it between t h e i n l e t manifold and t h e exhaust housing, The turbopump There a r e two types of turbine exhaust housings.

t e s t u n i t shown i n Figure No, 13 is a s i n g l e o u t l e t , conical housing which con- n e c t s t o t h e turbopump test stand exhaust ducting. The engine turbopump unit shown i n Figure No, 14 is a hemispherical, dual o u t l e t housing which connects t o t h e crossover duets t h a t carry exhaust gases t o t h e M - 1 l i q u i d oxygen turbopump The dual exit exhaust housing includes a connection f o r t h e 5-in. tur- t.urbine, bine by-pass duct a n d has an access pad which can be removed t o manually torque check t h e turbopump rotor, During engine operation, flow i n t h e turbine by-pass duet. would be controlled t o provide engine c a l i b r a t i o n as w e l l as t o provide pro- pellan t u t i l i z a t i o n control Method f o r Achieving Mechanical Design Objectives 2, The turbine i n l e t manifold is a combined t o r u s , nozzle assembly, and turbine casing, These components are welded i n t o one assembly giving a l i g h t - weight, strong structure that is f r e e t o move with t h e l a r g e thermal gradients.

The housings are supported from a main flange t h a t is located near t h e a x i a l center of t h e blade rowee This permits unrestrained thermal move- ment of t h e s t a t i c turbine housings without s e r i o u s l y a f f e c t i n g blade row align- ment and clearances, The support frame a t t a c h e s on t h e outside diameter of t h e Figure 9 Turbine Support Frame Page 15 .

Figure 10 Torus an?. Nozzle Assembly Page 16

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Fiaure 11 Nozzle T r o f i l e and Support Zings Page 17

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Figure 12 Aeversing How, Section and Vane P r o f i l e Fage 18 I

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Figure 14 3;xit ixhaust Housing Page 20 m a i n flange and extends back t o the pump housing. This frame is f r e e t o move r a d i a l l y at t h e turbine end t o a l l o w f o r contractions and expansions of the To reduce weight and thermal d i s t o r t i o n s , t h e main flange is turbine housings.

held together and attached t o the support frame by means of clamps. A small (.060-in.) weld beam is applied t o t h e flange j o i n t t o seal against hot &as leakage .

R 1 7 . 5 - i n . shaft w a s welded t o the f i r s t - s t a g e turbine r o t o r t o It a l s o s i m p l i f i e s t h e assembly allow c l o s e coupling of t h e turbine-end bearing.

P i l o t i n g journals and a s p l i n e a r e pro- with the pump r o t o r and seal assembly.

The opposite end has a curvic coupling vided f o r alignment and power transmission.

f o r centering and transmitting torque from the second rotor.

Attachment of t h e two hot turbine r o t o r s t o t h e cold pump r o t o r w a s accomplished by we of a long, e l a s t i c tie-bolt. T h i s allows thermal con- t r a c t i o n s and expansions during t r a n s i e n t conditions without overstressing the parts .

The turbine blades were forged t o obtain good n a t e r i a l strength and t o allow f o r an i n t e g r a l shroud on the outside diameter, which improves per- formance and permits blade vibration control. "he blades have i n t e r n a l c a v i t i e s t o reduce weight and improve thermally-induced s t r e s s e s . These blades are attached t o t h e d i s c s by electron-beam welding, which reduces t h e disc r i m thickness, and r e s u l t a n t weight as compared with the mechanical attachment methods.

To permit u n r e s t r i c t e d thermal movement during temperature tran- s i e n t s , the reversing row is divided i n t o s i x segments and retained i n the main housing f'.ange with a groove and tongue arrangement. Leakage around the i n s i d e diameter of the reversing row is controlled by a honeycomb seal that is positioned close t o t h e f i r s t r o t o r rim., 3 . Basis f o r Configuration Selection S e r i e s flow tufbine sets f o r t h e M - 1 l i q u i d hydrogen and l i q u i d oxygen turbopumps were s e l e c t e d over parallel flow turbine sets upon the bases of engine performance and turbopump complexity c r i t e r i a . P a r a l l e l flow turbine s e t s r e s u l t i n e i t h e r an increased gas generator weight flow or an increased number of t u r b i n e stages.

Because of t h e higher power and s h a f t speed requirements of t h e l i q u i d hydrogen turbopump, t h e l i q u i d hydrogen turbopump turbine set w a s s e l e c t e d This assured that turbine r o t o r stress l e v e l s t o b e the higher pressure set.

-.-nuldbe acceptable f o r both the l i q u i d hydrogen turbopump turbine set as well as t h e l i q u i d oxygen turbopump turbine s e t .

Single-stage, two-stage, and three-stage turbines were con- s i d e r c for t h e l i q u i d hydrogen turbopump s e t . The two-stage was selected over Page 21 t h e single-stage using engine performance and weight as t h e basis. The two-stage turbine reduces gas generator weight flow, thereby increasing engine s p e c i f i c impulse performance, Based upon t y p i c a l vehicle payload comparisons, the increase of engine performance f o r the two-stage turbine version more than o f f s e t the engine weight h c r e a s e f o r t h e second stage.

The two-stage turbine w a s s e l e c t e d over a three-stage turbine because t h e payload increase r e s u l t i n g from the increased engine performance Of the three-stage version is nearly o f f s e t by t h e t h i r d s t a g e weight increase, Also, the three-stage turbine turbopump requires an a f t o r outboard bearing t o achieve s h a f t critical speed requirements; whereas, the two-stage turbine turbo- pump achieves shaft c r i t i c a l speed c r i t e r i a with a forward o r inboard bearing arrangement. The a f t bearing would increase turbopump complexity by requiring a d d i t i o n a l bearing coolant c i r c u i t s , bearing coolant c i r c u i t control valving, and a s p e c i a l turbine s t r u c t u r e f o r the support and alignment of the aft bearing, A velocity-staged turbine was s e l e c t e d in preference t o a pres- sure-staged turbine because t h e second-stage s t a t o r blade and diaphragm of the velocity-staged turbine is l i g h t e r and simpler. For t h e s p e c i f i e d operating con- d i t i o n s , turbine e f f i c i e n c i e s f o r t h e velocity-staged and pressure-staged turbine6 a r e e s s e n t i a l l y equal, Impulse blading w a s s e l e c t e d i n preference t o reaction blading f o r t h e f i r s t - s t a g e because the impulse blading gave higher e f f i c i e n c y f o r the design conditions, Also, the impulse blading gives lower turbine housing pres-

sure which minimizes turbine housing weight . The second-stage s t a t o r (reversing

blade row) and the second-stage r o t o r have s l i g h t r e a c t i o n flow t o s u f f i c i e n t l y a c c e l e r a t e the flow t o o f f s e t f r i c t i o n a l losses.

C. MATERIAL SELECTION 1, Material Requirements To meet t h e mechanical design point conditions as w e l l as t h e design considerations, the turbine materials should have the following q u a l i t i e s : a . High s t r e n g t h at operating temperatwes t o permit t h i n cross-sections.

Good elongation c h a r a c t e r i s t i c s over a wide range of tem- b o peratures ( - 4 2 0 ' F t o +1300°F) t o withstand thermal shock f a c i l i t a t e crack-free welding, and permit localized yielding.

c e Good stress-rupture p r o p e r t i e s up t o 130O0F.

d. Corrosion resistance.

Page 22 e. Reasonable machinability and metal-forming q u a l i t i e s , f . Be s u i t a b l e f o r welding by gas tungsten a r c welding o r electron-beam method.

Capable of being annealed f o r s t r e s s relief and heat g.

t r e a t e d f o r maximum strength.

2. Selected P a t e r i a l s A comparatively new material that conformed t o these require- T h i s material, ments was s e l e c t e d f o r most components of t h e M - 1 f u e l turbine.

Its material Inconel 718, is a nickel-chromium a l l o y that is age-hardenable.

propertles are as follows: a. For stress-rupture limited applications, the material is solution-annealed at 1.800~~. It is aged at l325OF and 1150'F with t h e following properties: For bars and forgings, 23 hours l i f e at l W ° F and (1) 75,000 p s i stress.

For sheet and p l a t e , 23 hours l i f e at 1300OF and ( 2 ) 72,500 p s i stress.

of ambient t o above 12W°F, b. For tensile-limited applications It is aged at 1350°F and 1200OF the material is solution-annealed at 1950°F.

with t h e following properties: The room temperature-tensile, ultimate and y i e l d (1) s t r e n g t h s a r e 175,000 p s i and 145,000 psi, with minimum elongations of 1 2 9 6 .

The 1200'F t e n s i l e , ultimate, and y i e l d s t r e n g t h s are (2) 145,000 p s i and 125,ooO p s i , with minimum elongations of l & Rene9 41 was s e l e c t e d f o r the f i r s t - s t a g e r o t o r shaft extensionp mainly because its coefficient of thermal contraction was s l i g h t l y l e s s than Inconel 718. !Chis f a c t o r permitted tolerances in the turbine journals s o they would assemble i n t o t h e pump r o t o r with a s l i g h t l y loose f i t at room temperature.

At -420°F, t h e pump r o t o r w i l l contract around the %ne' 41 shaft producing an Also, t h i s material has good mechanical properties and i n t e r f e r e n c e - f i t j o i n t .

can be welded t o the 718 turbine rotor.

Another f a c t o r i n s e l e c t i n g Rene' 41 f o r the s h a f t concerned t h e turbine bearing inner race f i t . The inner race is of & I C s t a i n l e s s s t e e l and had a s i g n i f i c a n t l y l e s s e r contraction at -420°F than 718, but it is closer This allowed a smaller roon temperature t o t h e contraction r a t e of Renee 4 1 .

i n t e r f e r e n c e - f i t .

D . STRUCTURAL CRITERIA

1 . S t a t i c or Non-Rotating Components The s t a t i c and non-rot.ating components were designed with t h e following minimum c r i t e r i a : a, Proof pressure ( y i e l d ) z 1.2 ( f a c t o r of s a f e t y ) times Po, where Po equals the m a x i m u m s t a t i c pressure expected t o e x i s t a t any time.

b . B u r s t pressure (ultimate) = 1,6 ( f a c t o r of s a f e t y ) times PO.

c o The s a f e t y f a c t o r s from nominal working pressures f o r flex- i b l e hoses, tubing, and pressure f i t t i n g s with diameters less than 1.5-in. a r e 2.0 f o r proof and 4.0 f o r burst.

Pressure vessels a r e designed t o withstand 500 operating do pressure cycles.

e. Proof and burst pressures f o r tests at ambient temperatures are adjusted t o c0mpensat.e f o r material strengths at operating temperatures.

Proof pressure is that test pressure t o which an item is subjected f o r a minimum of two minutes without deformations t h a t adversely a f f e c t t h e rocket engine.

I n addition t o t h e above s a f e t y f a c t o r s j , a f a c t o r of 1.15 f , s h a l l be used f o r a l l f i t t i n g s , mounts, and b o l t s which e i t h e r c o n s t i t u t e the only load path o r whose f a i l u r e would i n t e r f e r e with t h e proper functioning or i n t e g r i t y of any engine componentD The section a r e a s were determined using t h e material strength a t t h e maximum expected operating temperature, For most parts, t h e ultimate strength was used as the l i m i t i n g s t r u c t u r a l c r i t e r i a . The margin of safety is defined as follows: M . S, = Allowable S t r e s s - 1 Calculated S t r e s s In general, the calculated s t r e s s includes the applicable s a f e t y factor.

The fatigue margin of s a f e t y is applied t o components having a For f l u c t u a t i n g stress and is determined by use of a modified Goodman diagram.

these calculations, t h e a l t e r n a t i n g stress is assumed t o be 30% of the m a x i m u m mean stress. The f a t i g u e margin of s a f e t y is c a l c u l a t e d as follows: M, S o = Allowable S t r e s s From Goodman Diagram 1.25 x Calculated S t r e s s Page 24 _____ This calculation includes a 1.25 s a f e t y factor.

2 . Rotating Components

Rotating components were designed with t h e following minimum c r i t e r i a : Design Yield s 1.0 (Factor of Safety) times l i m i t load, a.

where l i m i t load is t h e maximum predicted load which t h e system may experience under s p e c i f i e d operating conditions.

Design Ultimate = 1.5 (Factor of Safety) times l i m i t load.

b.

C. Components subjected t o elevated temperatures s h a l l be de- signed using material properties a t the estimated maximum temperature environmento d. The f a t i g u e margin of s a f e t y is applied t o components having a f l u c t u a t i n g s t r e s s and is determined by use of a modified Goodman dia- The a l t e r n a t i n g s t r e s s is assumed t o be 3& of the maximum mean s t r e s s .

g r a m .

The f a t i g u e margin of s a f e t y , which includes a 1.25 f a c t o r of s a f e t y , is cal- culated as follows: M. S . = Allowable S t r e s s from Goodman Diagram 1 . 2 5 x Calculated S t r e s s The turbine blade l i m i t i n g s t r e s s was determined by adding c e n t r i f u g a l , gas bending, and assumed a l t e r n a t i n g stress a t 110% of operating speed t o determine the f a t i g u e margin of safety. Margins of s a f e t y included t h e e f f e c t of t r a n s i e n t thermal stresses.

The turbine r o t o r design w a s based upon symmetrical sections of constant stress, except a t t h e neck ( j u s t below the blade platform) and the bore.

is s u f f i c i e n t l y reduced t o permit inducer blade shedding p r i o r t o The neck a r e a The bore s t r e s s was limited t o less than twice the y i e l d s t r e n g t h d i s c f a i l u r e .

a t t h e l o c a l temperature. The average tangential s t r e s s i n t h e d i s c w a s l i m i t e d by a 2.0 s a f e t y f a c t o r upon t h e ultimate strength. Transient thermal-induced stresses were superimposed upon the combined s t r e s s e s t o obtain calculated mar- g i n s of safety.

3. Vibration and Acceleration Loads Operation a t a point which will e x c i t e t h e n a t u r a l frequency o r some harmonic of a component should be avoided unless the vibration-induced load is below t h e design c r i t e r i a .

The engine-induced, f l i g h t acceleration loads within the turbine are as follows: design a, Longitudinal b o Lateral c e Angular d o Gimbal snubbing E. COEIPONENT STRESS SUMMARY 1. Turbine I n l e t Manifold The maximum s t r e s s e s produced from the shock loading and from the 2fi g vibratory l i n e loads do not occur at the same locations as the maximum s t r e s s e s from gas pressure and thermal gradients. Shock load produces a moment on t h e i n l e t nozzle, This load occurs during an emergency shutdown when t h e end of t h e i n l e t duct is blown-off t o rapidly r e l e a s e turbine supply pressure. The c o n t r o l l i n g s t r e s s e s a r e ’ t h o s e produced from gas pressure and thermal gradients.

Figure No. 1 5 shows the gas temperature t r a n s i e n t t o which the t o r u s and nozzles are exposed.

by use of an Aerojet-General Cor- The s t r u c t u r e is analyz poration computer program (Job No, 1040) @) The analysis shows t h a t yielding w i l l not be the c r i t i c a l mode of f a i l u r e , Fatigue w i l l cause f a i l u r e after 2200 starts.

A tabulation of i n l e t manifold stresses with 1175 p s i i n t e r n a l pressure and a steady-state temperature of 1300°F is shown i n Figure No. 16.

The most severe stress case with 1175 psi i n t e r n a l pressure and t h e maximum transient thermal gradient is shown i n Figure No, 170 The calculated s t r e s s e s developed using the required f i x t u r i n g f o r hydrotesting, which imposes d i f f e r e n t loads at some points, are tabulated i n Figure No. 18.

A tabulation of t h e s t r e s s e s with 1175 p s i i n t e r n a l pressure, a temperature of 1300°F, with 2)/2g l i n e loads applied, and t h e blow-off valve shock load imposed, are shown i n Figure No. 19.

‘ ‘ ) T h i s program wa6 devised from information i n t h e p a p e r 9 A h m e r i c a l of t h e Equations of Thin S h e l l s of Revolution, by Radkowski, P, Po Davis, R. M . , and Boldul, El. R., Avco Corp.

Page 26

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1 I M A X

t I Figure 18 I n l e t Manifold S t r e s s , Proof-Pressure Test Page 30 + Figure 19 I n l e t Maaifold Stress, L i n e and Shock Loads Page 31 2, First-Stage and Second-Stage Rotors a. First-Stage Rotor S t r e s s The f i n a l r o t o r configuration was the r e s u l t of a stress a n a l y s i s program wherein an i n i t i a l design w a s analyzed, revised, and re- This process was repeated u n t i l an optimum configuration was de- analyzed.

The optimum configuration yielding a turbine which is of minimum veloped.

weight while s t r u c t u r a l l y adequate, The stress a n a l y s i s considered the following s t r e s s pro- ducing items: centrifugal; pressure; thermal; vibration; and thermal fatigue.

The f i r s t s t e p i n t h e analysis was t o determine the c r i t i - c a l time period i n terms of maximum thermal stresses, This was found t o be 230 The s e c from t h e time equal t o zero i n t h e pump start and operation cycle.

r e s u l t s of t h e analysis show t h e following: Local yielding w i l l occur i n the turbine r o t o r bore (1) as a r e s u l t of the combined c e n t r i f u g a l and thermal stresses, The maximum t h e o r e t i c a l stress l e v e l is 275,000 p s i at 230 s e c running time. This stress l e v e l occurs i n a l o c a l a r e a of the 1,875-inO diameter bore through t h e turbine Table No, I gives a tabulation of t h e d i s c bore stress from 10 sec t o disc.

230 sec.

TABLE I FIRST STAGE D I S C STRESS CONDITION *MAX. TANG. YIELD ULTIMATE ELONGATION (SZC) STRESS (KSI) LOCATION (KSI) (KSI) % 10 150 Ten, 165 230 15 Bore 60 240 Ten. Bore 165 230 15 100 220 Ten, Bore 165 230 15 120 265 Ten. Bore 165 230 15 230 275 Ten, Bore 165 230 15 *The average tangential stress is 65 K S I at 14,550 rpm.

The t h e o r e t i c a l l o c a l s t r e s s value (calculated (2) e l a s t i c s t r e s s ) i s not considered t o be detrimental s i n c e it is less than twice Page 32 ( 5 ) t h e material y i e l d stress value .

During the f i r s t s t r e s s cycle, l o c a l p l a s t i c flow w i l l When occur but i n subsequent cycles all l o c a l deformation will be e l a s t i c .

cycling is purely e l a s t i c , t h e low-cycle, p l a s t i c s t r a i n fatigue failures are not probable.

B u r s t speed is 21,800 r p m .

( 3 ) Margin of s a f e t y , which includes a s a f e t y f a c t o r of (4) 1.5, is a.49 based upon an average stress of 65,000 psi.

"he r o t o r blades were t r e a t e d i n a separate analysis.

The stress analysis included centrifugal forces9 pressure load, thermal e f f e c t s , v i b r a t i o n , and thermal fatigue. The a l t e r n a t i n g stress was assumed t o be 30% of mean stress. The blades were shown t o be adequate i n a l l respects (see Figure The margin of s a f e t y is +.72 at l m ° F based upon a m a x i m u m stress of No, 20).

59,000 p s i which includes a s a f e t y f a c t o r of 1.2. The highest thermal stress i n t h e blade occurs 1.25 sec a f t e r start (see Figure No. 21) and is calculated t o momentarily reach 110,000 psi.

The turbine shaft stress analysis w a s a case of ?roving The basic assumption was that only 25% of t h e s p l i n e t h e s p l i n e t o be adequate.

The shaft is adequate because t e e t h were engaged and carrying the f u l l torque.

c a l c u l a t i o n s give t h e following margins of safety: Cycle bending = 0,62 based on a modified Goodman diagram and stress of 63,800 psi which includes a s a f e t y f a c t o r of 1.25. The c y c l i c bending stress includes the e f f e c t of an assumed a l t e r n a t i n g stress of 10% of t h e s t a t i c stress plus a stress concentration f a c t o r of 1.505 applied t o t h e c y c l i c stress.

Cyclic shear = +0.18 based on a modified Goodman d i a - The c y c h c gram, and a stress of 52,500 p s i which includes a s a f e t y f a c t o r of 1.25.

s h e a r stress includes an assumed a l t e r n a t i n g s t r e s s of 10% of the static s t r e s s p l u s a stress concentration factor of 1.505 applied t o t h e cyclic stress.

Compression = +.30 based upon a stress of 90,500 p s i which includes a s a f e t y f a c t o r of 1.1.

The shaft stress values are conservative because only t h e The cryogenic room temperature material mechanical properties were used.

(5)Criteria from Section I11 of t h e ASME Boiler and Fkessure Vessel Code f o r Nuclear Vessels, p.6, Library of Congress Catalog Card No. 56-3934 1963 Figure 20 ,

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TEUPERATURE t N OR

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Figure 21 First-Stage Blade Temperature Gradients Page 35 mechanical properties as well as t h e ultimate t e n s i l e and y i e l d s t r e s s e s a r e considerably higher, b. First-Stage Rotor Vibration Calculated values f o r resonant points of the blades re- vealed no problem areas. The resonant speeds associated with t h e first harmonic of the 37 nozzles passing frequency and e x c i t i n g the f i r s t a x i a l and t a n g e n t i a l bending modes of t h e blades a r e as follows: Tangential = 10,400 rpm Axial = 11,600 rpm The blades a r e considered adequate because these vibration values are s u f f i c i e n t l y removed from t h e pump design operating point so as t o cause no problem. These frequencies w i l l be passed through quickly during the start t r a n s i e n t and shutdown phases of pump operation. Figure No. 22 shows all of t h e blade vibration frequencies.

The primary s t i m u l i a r e the 37 nozzles up- stream of the blades.

The turbine d i s c was analyzed with respect t o axial vi- Preliminary calculations revealed t h a t t h e two nodal diameter mode brations.

shapes gave the lowest c r i t i c a l speed values. Assumed values of r a d i a l and tar- g e n t i a l s t r e s s representative of expected s t r e s s were used i n t h i s a n a l y s i s t o obtain t h e minimum c r i t i c a l speeds, The r e s u l t s are as follows: S t a t i c = 14,370 rpm f o r a r a d i a l s t r e s s of 100,000 p s i and a t a n g e n t i a l s t r e s s of 50,000 psi.

10% overspeed = 18,000 rpm.

These c r i t i c a l speed values9 when compared w i t h t h e pump opera- t i n g speed range, show t h e turbine d i s c design t o be s a t i s f a c t o r y .

c e Second-Stage Rotor S t r e s s The stress a n a l y s i s f o r t h i s turbine r o t o r w a s made con- c u r r e n t l y with the f i r s t - s t a g e turbine rotor. A l l criteria were t h e same and included centrifugal forces, pressure load, thermal e f f e c t s , v i b r a t i o n , and thermal fatigue. Alternating stress w a s assumed t o be 30% of t h e mean s t r e s s .

The f i n a l design is the r e s u l t of repeated analyses w ? - . - T - i n d i s c p r o f i l e s were a l t e r e d as problem areas or l i g h t l y - s t r e s s e d a r e a s W . ' uncovered. The f i n a l configuration is considered optimum. The following i i the calculated values from the f i n a l analysis.

Burst speed is 22,loO rpm.

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Figure 22 First-Stage Blade Vibration Frequencies Page 37 The margin of s a f e t y , which includes a f a c t o r of s a f e t y of 1.5, is +O.53 based upon an average tangential s t r e s s of 66,000 psi. F i e w e No. 23 shows a tabulation of t h e disc s t r e s s e s , which include thermal s t r e s s e s .

A maximum combined t h e o r e t i c a l stress of 205,000 p s i occurs i n t h e l o c a l a r e a on t h e surface of the 1.562-in. diameter bore through This s t r e s s l e v e l is not considered detrimental t h e center of t h e turbine disc.

because it is less t h a n twice t h e y i e l d stress value f o r t h e material.

The curvic coupling stress values proved t o be r e l a t i v e l y low. The maximum stress is 55,600 psi. The minimum margin of s a f e t y is +0.37 including a s a f e t y f a c t o r of 1.2.

The blade s t r e s s analysis yielded t h e following r e s u l t s : stress of 61,440 psi.

(1) Maximum combined Margin of s a f e t y t o y i e l d point at 1 3 0 0 ' F is +.38, (2) including a s a f e t y f a c t o r of 1.2.

Margin of s a f e t y t o rupture is +.22 including a ( 3 ) s a f e t y f a c t o r of 1.6.

Figure No, 24 summarizes t h e blades s t r e s s e s .

The blades a r e considered s a t i s f a c t o r y i n a l l respects.

d. Second-Stage Rotor Vibration The vibration analysis was made using t h e same c r i t e r i a as The primary s t i m u l i are t h e 67 upstream used f o r t h e f i r s t - s t a g e turbine rotor.

reversing vanes. Secondary s t i m u l i are the 37 nozzles upstream of t h e f i r s t rotor. The calculated resonant frequencies f o r t h e blades expressed i n terms of pump revolutions per minute a r e as follows; F i r s t t a n g e n t i a l = 1100 F i r s t a x i a l = 2100 F i r s t t o r s i o n a l = 3100 Second t a n g e n t i a l = 10,100 Second a x i a l = 14,900 Second t o r s i o n a l = 12,100 F

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T RA\ LING ED GE lBACKSIDE I I I Figure 24 Second-Stage Blade S t r e s s Page 40 These resonant frequencies present no problem when com- Figure No. 25 shows a p l o t of the pared with t h e pump operating speed band.

various vibration frequencies.

As in the The turbine disc axial vibration w a s analyzed.

case of the f i r s t stage turbine, only t h e two nodal diameter modes were inves- t i g a t e d because this provided t h e lowest c r i t i c a l speed values. T h i s a n a l y s i s revealed t h e following: S t a t i c c r i t i c a l speed is 14,430 r p m with 100,OOO p s i (1) r a d i a l and 50,OOO p s i t a n g e n t i a l loads.

A t 14,550 rpm, the disc axial c r i t i c a l speed is (2) 17,640 rpm with 100,OOO psi radial and 50,OOO p s i t a n g e n t i a l loads.

These c r i t i c a l speeds are s a t i s f a c t o r y when compared with t h e pump design operating speeds.

3. Reversing Row a. S t r e s s S t r e s s a n a l y s i s of t h e reversing vane segments considers The thermal-induced loads were loads r e s u l t i n g from gas pressure and velocity.

The analysis shows the minimum made negligible by design of the component.

The margin of s a f e t y t o be +lo? and the maximum stress t o be 27,400 psi.

stresses a r e tabulated i n Figure No. 26.

b. Vibration Vibration analysis of the reversing vanes shows the highest resonant frequency, expressed in terms of pump speed, t o be t h e t h i r d t a n g e n t i a l When compared All other modes a r e below t h i s frequency.

mode a t 10,000 rpm, with t h e pump operating speed band, t h e reversing vanes present no d i f f i c u l t y as Figure No. 27 shows a l l of t h e vibration frequencies.

p e r t a i n s t o vibration.

4. Rotor Tie-Bolt To prevent j o i n t separation during all conditions of thermal and dynamic loading, a b o l t pretorque of 11,500 in.-lb is required (see Figure With this pretorque, and during the worst period of thermal gradients, No, 28).

This includes stress concentration f a c t o r s up t o t h e margin of s a f e t y is + O . l g .

2.78, a f a c t o r of s a f e t y of 1.2, and a thread c o e f f i c i e n t of f r i c t i o n of 0 . 1 5 , The natural frequency of t h e t i e - b o l t , with the center piloted, "his is well above the operating speed.

occurs a t 52,500 rpm.

Page 41 Figure 25 5 e c ond-S t age Blade Vib rat i on Frequencies Page 42 tl.7 t2.O INS\CNIFI CANT Figure 26 Reversing Row Stress Page 43

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2 I 4

6 8 I O

Figure 27 Reverslng Aow Vibration Frequencies Pzge 44 Figure 28 Turbine Rotor, Tie-Bolt and Adjacent Parte Page 45 5 . Bearing Housing Seal The most severe s t r e s s , pressure, and thermal gradients c m - tained i n t h e s e a l components, plus the thermal movements of t h e i n l e t daiif016, were considered.

The most severe s t r e s s occurs 10 sec a f t e r s t a r t - u p a t the junction with the end of t h e bellows and t h e inner support cone (see Figure No.

This stress is caused by bending. The apparent computed e l a s t i c s t r z s s 2 9 ) .

is 443,000 psi. Yielding w i l l occur and t h e problem becomes one of low cyCli-C f a t i g u e because the primary membrane s t r e s s is low.

A low c y c l i c f a t i g u e a n a l y s i s of the bellows revealed f a i l u r e w i l l occur after 520 cycles. However, when a f a c t o r of s a f e t y of 1.6 is applied A cycle is des- t o t h e t o t a l s t r a i n , t h e f a t i g u e l i m i t reduces t o 222 cycles.

cribed as a hot gas run of a t l e a s t 10 sec.

6. Turbine Exhaust Housings Two d i f f e r e n t exhaust housings a r e used on t h e M-1 f u e l turbine, A n exhaust cone frustum is used when the turbopump is t e s t e d apart from t h e For complete engine t e s t s p a dual e x i t hemisphere exhaust housing is engine, used.

a. The Exhaust Cone (Figure No. 13) The s t r u c t u r a l evaluation of t h e exhaust cone is based upon the following extreme operating conditions at thz t e s t stand.

(1) G a s temperature - 1025OF

bhximum I n t e r n a l Pressure - 386 p s i a

( 2 )

External applied loads - 400,000 in-lbs bending

( 3 ) moment e The margin of s a f e t y , based upon the ultimate s t r e n g t h of the m a t e r ' a l at lO25'F, with a 1.6 f a c t o r of s a f e t y , is 1.17 f o r hoop stress and 4.5 Pn buckling from t h e bending moment.

b. Dual Exit Exhaust Housing (Figure No. 14) The s t r e s s a n a l y s i s f o r t h e t u r b i n e exhaust housing i n - cludes t h e following loads: (1) I n t e r n a l Pressure ( 2 ) Inertial-induced l i n e loads on t h e cross-over duct Page 46

UTER SUPPORT

RING

DIA

I k O N E FRUSTUM

12.67 DIA

I N N E R SUPPORT

Figure 29 Bearing Housing Seal Assembly Page 47 flanges and on the propellant u t i l i z a t i o n valve flange, Vibration-induced l i n e loads at a l l flanges.

(3) The material strength at l O O O ' F was used t o determine margins of safety. Thermal-induced s t r e s s e s were not determined. They were considered t o be minor because of the t h i n walls and the type of construction, which does not r e s t r a i n t h e housing, The f a c t o r of s a f e t y f o r pressure loading is a f a c t o r of 1.6 based upon an ultimate s t r e n g t h of 155,000 p s i at 1000'F.

The f a c t o r of s a f e t y f o r i n e r t i a loading is a f a c t o r of 1 . 5 based upon an ultimate s t r e n g t h of 155,000 p s i at 1 0 0 0 ' F .

The maximum combined stress occurred i n t h e spherical- shaped wall at a point approximately midway between the propellant u t i l i z a t i o n valve flange and the junction of t h e t r a n s i t i o n t o t h e ll.4-in. diameter cross- over flange, The minimum margin of s a f e t y at t h i s point was calculated t o be +0.17.

Support Frame and Turbine Main-Flange Clamps 7.

TFigures No. 2 and No. 9 ) The s t r e s s a n a l y s i s was performed t o determine the s t r u c t u r a l i n t e g r i t y of t h e frame with loads as imposed at the test stand. These loads include the t e s t area turbine exhaust adapter l i n e and its bellows. It was a spring rate of 4500 lb/in. in the axial assumed t h a t t h e bellows would have and transverse directions and would deflect one inch in each direction during The maximum expected pressure i n t h e manifold is 1175 psi and in operation.

386 psi. The maximum expected temperature is 1NO'F.

t h e exhaust cone, With these conditions, the margin of s a f e t y , based upon t h e material y%eld strength with a 1.0 f a c t o r of s a f e t y , is 1 . 0 9 and based upon the ultimate strength with 1 . 5 f a c t o r of s a f e t y , t h e margin of s a f e t y is 0.97. The b o l t s attaching the frame t o t h e pump discharge housing (AS4013, P/N 712329-311) have a margin of s a f e t y of 0.60 based upon t h e y i e l d s t r e n g t h and 0.18 based upon ultimate strength.

The b o l t s a t t a c h i n g the half-clamps t o t h e frame ends and the paired clamps, have a margin of s a f e t y based upon y i e l d s t r e n g t h of 0.36, w i t h a b o l t pretorque of 1190 in.-lb.

main flange j o i n t margin of e a f e t y , with a f a c t o r of s a f e t y The 1.6 based upon the material ultimate strength, is 0.28.

of Page 48 DYNAMIC BALANCING A N D ASSpfBLY TECHNIQUES F .

1. Dynamic Balancinq The dynamic balancing of the turbine r o t o r s is accomplished in a series of steps, starting with t h e f i r s t - s t a g e turbine and ending with the complete turbopump r o t a t i n g assembly. All balancing is performed on a Gisholt U type machine. The engineering dynamic balance requirements are based upor, limits t h a t can be readily achieved with the U type machine. The balancing sequence is as follows: The first-stage turbine rotor is balanced using a two a .

plane system.

The second-stage turbine r o t o r is i n s t a l l e d on the bal- b.

anced f i r s t - s t a g e turbine r o t o r and the assembly is checked f o r the degree of unbalance. The second-stage rotor is removed and rotated r e l a t i v e t o its initial position. The amount It i s then reassembled t o the f i r s t - s t a g e rotor.

of unbalance is again determined. This sequence is repeated as often as neces- sary t o ascertain the assembled condition that provides the least unbalance.

The assembly is then balanced t o the drawing requirements with corrections made t o the second-stage only. The rotors a r e match-marked.

“he balanced turbine rotor assembly is i n s t a l l e d i n a C.

balanced pump rotor assembly. The condition of least unbalance is determined by repeated disassembly, rotation, and reassembly. The complete rotating assem- bly, having t h e smallest degree of unbalance, is balanced using a two plane system. A l l components: pump r o t o r , f i r s t - s t a g e turbine r o t o r , second-stage turbine r o t o r , and tie-bolt are match-marked and i d e n t i f i e d with a common s e r i a l number.

2. Assembly Techniques The turbine components are assembled t o the f u e l pump i n the following sequence (use Figure No. 2 86 a reference).

The turbine frame is bolted t o t h e pump discharge housing.

a.

b. The turbine i n l e t manifold is attached t o t h e turbine frame by means of an assembly fixture.

A gaging tool, which simulates t h e turbine shaft and disc, C.

is i n s t a l l e d i n t h e pump rotor. Measurements a r e made from gaging surfaces on t h e t o o l t o various surfaces on the i n l e t manifold and turbine bearing housing.

These measurements are recorded and t h e gaging t o o l is removed.

d. The recorded measurements are used t o calculate the fol- lowing: The shim thickness required at t h e turbine frame (1) discharge housing i n t e r f a c e t o c o r r e c t l y position the turbine i n l e t manifold, r e l a t i v e t o t h e turbine bearing housing.

The spacer length required t o c o r r e c t l y position (2) the f i r s t - s t a g e turbine r e l a t i v e t o the turbine bearing housing f o r proper lift- off seal operation.

e. The turbine frame shims are i n s t a l l e d i n the turbine frame discharge housing i n t e r f a c e and t h e b o l t s a r e torqued.

The i n l e t manifold position is checked r e l a t i v e t o the f .

The diaphragm s e a l , which forms the closure between turbine bearing housing.

the i n l e t manifold and t h e turbine bearing housing, is i n s t a l l e d and the gas tungsten a r c w e l d s are made.

The f i r s t - s t a g e turbine r o t o r is i n s t a l l e d as follows: go W a r m gaseous nitrogen is flowed through t h e turbine (1) bearing housing t o raise t h e temperature of the pump rotor.

The turbine shaft is placed i n l i q u i d nitrogen.

(2) When the temperatures of the components become sta- ( 3 ) b i l i z e d , the f i r s t - s t a g e r o t o r is positioned by a l i g n i n g s p l i n e t e e t h and dy- This operation namic balancing match marks, then i n s e r t e d i n t o t h e pump rotor.

is completed as rapidly as possible.

The turbine t i e - b o l t is i n s t a l l e d with s u i t a b l e f i x - (4) t u r e s t o permit applying an axial load t o t h e turbine equivalent t o the f i n a l assembly preload. The assembly is allowed t o r e t u r n t o room temperature.

The reversing row is i n s t a l l e d with t h e anti- ( 5 ) r o t a t i o n s l o t s aligned with t h e lugs i n t h e i n l e t manifold main flange.

( 6 ) The f i x t u r e is removed from the turbine bie-bolt.

The second-stage r o t o r is i n s t a l l e d , taking care t o a l i g n curvic coupling teeth and dynamic balancing match marks.

(7) The turbine t i e - b o l t is positioned a x i a l l y . The nut and lock r i n g a r e i n s t a l l e d and torqued t o drawing requirements.

The f i x t u r e , which has been positioning t h e i n l e t (8) The exhaust housing manifold, is removed and the exhaust housing is i n s t a l l e d .

is held i n position with a minimum number of clamps t o provide access t o the f h g e f o r welding. The turbine i n l e t manifold and exhaust housing j o i n t is sealed by gas tungsten a r c welding. After t h e welding is completed and leak checks are made, all clamps a r e i n s t a l l e d and torqued t o drawing requirements.

The assembly is completed by adding appropriate ( 9 ) closures as required t o protect the unit.

G. DESCRIPTION O F COMPONENT FABRICATION 1. Turbine I n l e t Manifold The t u r b i n e i n l e t manifold is a completely gas tungsten a r c t o r u s , nozzle r i n g , and turbine casing (see welded (GTAW) assembly composed of Figure No. 1 0 ) . Inconel 718 alloy(*) is used f o r a l l components. P a r t s are After t h e subassemblies a r e stress relieved a f t e r each weld is completed.

welded i n t o t h e final assembly, the complete manifold is heat t r e a t e d by solu- tion-annealing and aging.

“he t o r u s is formed from sheet material i n t o doughnut-shaped half-shells. These half-shells are then GTAW welded together on t h e outside A n eight inch dia- diameter and i n s i d e diameter forming a complete doughnut.

meter i n l e t neck is i n s e r t e d a t an eighteen degree angle, using double thickness material f o r f i f t e e n degrees t o each s i d e of the center, which is the reinforced A conoseal flange with twenty-four b o l t holes is welded on hole cutout area, the upstream end of the inlet neck f o r attachment of t h e helium start l i n e , the hot gas by-pass line, and t h e emergency blow-off valve.

On t h e downstream s i d e of the t o r u s , a 4 . 5 - i n . wide s l o t is machined out on a 23,O-in. mean diameter f o r i n s e r t i o n of t h e nozzle ring. This nozzle r i n g has flange extensions t h a t match the t o r u s w a l l s , which f a c i l l t a t e GTAW welding t h e two p a r t s together.

The nozzle ring is made from two forged r i n g s and thirty-seven forged nozzles, a l l GTAW welded together. The nozzles a r e rough-forged into approximate shape in a long section; then, several nozzles a r e cut from each The outside p r o f i l e is ground t o close tolerances and forms the gas section.

The i n s i d e of the nozzle is e l e c t r i c a l discharge machined t o give a passage.

This reduces thermal s t r e s s e s and a l l o u s the low w a l l thickness of 0.125-in.

pressure gas t o c i r c u l a t e between t h e outside diameter and i n s i d e diameter giving b e t t e r pressure distribution and more even thermal gradients.

A turbine casing, which contains t h e main flange is then welded t o an extension on the outer r i n g , completing the nozzle r i n g assembly ( s e e Figures No. 10 and N o . 1 1 ) .

(6)Frick, V., H u n t V . , Inouye, F. T., and Janser, Go R. , Summary of &perience

Using Inconel 718 on M-1 &Kine, Aerojet-General Report No. 880-37, 1 March 1966 (To be published as a NASA Contractor Report) The t o r u s downstream diameters are now c a r e f u l l y machined t o match t h e nozzle ring extensions and t h e two parts a r e fastened together by circumferential welds. These welds a r e penetrant inspected as w e l l as radio- g r a p h i c a l l y inspected. The part is then heat-treated, proof t e s t e d , and l e a k tested. The last process is t o f i n i s h machine a l l flanges and indexing surfaces.

Some welding problems were experienced with t h e 718 alloy.

The main problem was from contamination, (ice., occlusions i n t h e weld rod, i m - p u r i t i e s i n the helium and argon gas s h i e l d s , dust contaminated weld booths, m d improper gas shielding of t h e weld). Weld shrinkage and warping, which is not r e s t r i c t e d t o 718 alloy, was a l s o a problem. Proper f i x t u r i n g and improved welding procedures eliminated the l a t t e r problem. Optimum weld j o i n t confim- r a t i o n caused some concern and required the use of sample weld j o i n t s t o a r r i v e a t t h e b e s t solution.

2. First-Stage Rotor and Blades The turbine r o t o r is characterized by a s h a f t extending from one s i d e of the turbine disc. This s h a f t is i n s t a l l e d i n t o t h e pump r o t o r for- ming a r i g i d j o i n t f o r t h e transmission of power. The t u r b i n e s h a f t has a s p l i n e and two piloting diameter, which form the i n t e r f a c e t o t h e pump rotor.

The turbine s h a f t also c a r r i e s t h e journal f o r t h e turbine end r o l l e r bearing.

Figure No, 3 shows a cross-section of the r o t o r , s h a f t , and blades.

The turbine disc s i d e opposite from t h e s h a f t has a curvic (see Figure No. 5). This coupling serves as both t h e power trans- coupling mission and locating device f o r t h e second-stage turbine r o t o r .

The turbine d i s c is of minimum thickness consistent with s t r u c t u r a l i n t e g r i t y , The d i s c contour w a s derived from an elaborate stress a l l operating loads but a l s o a l l thermal a n a l y s i s which included not only The f i n a l d i s c s t r e s s e s generated during chilldown and start t r a n s i e n t phases.

shape represents the optimum configuration which satisfies both functional and weight c r i t e r i a .

of t h e turbine d i s c has 80 blades. Each blade The periphery integral t i p shroud for increased performance. As a r e s u l t of vibration has an a n a l y s i s , t h e blade shrouds are joined by welding at 40 plaues, t h e a l t e r n a t e shroud j o i n t s being t h e ones welded. The blades are hollow t o reduce weight, The which, i n t u r n , reduces r o t a t i o n a l loads imposed upon t h e turbine disc.

This method blades a r e attached t o t h e turbine d i s c by electron-beam welding.

of f a b r i c a t i o n produces t h e minimum weld a r e a , which, i n turn, r e s u l t s i n t h e minimum d i s c r i m size t o accept t h e blades.

a. Machining Process Techniques L The blades were machined from Inconel 718 die forgings.

Page 52 c The machining process used several d i f f e r e n t techniques.

The concave a i r f o i l surface was made by electro- (1) This uses an electrode which duplicates t h e desired chemical milling, ECM.

Material is removed passing e l e c t r i c current from t h e electrode through shape.

N o spark is involved, a conductive salt solution t o t h e metal being worked.

(2) The remaining a i r f o i l surfaces and shroud contours were made by conventional blade milling operations, "he blade platform and shroud platform configurations (3) were machined by electro-chemical grinding, ECG, Blade c a v i t i e s were produced by t h e E M process.

(4) Each blade w a s completed by welding a cap i n t o the end of t h e cavity. This cap w a s necessary f o r adequate shroud strength.

The turbine ahaft is made from a Rene' 41 forging, which was s e l e c t e d not only f o r its strength, but a l s o f o r its coefficient of thermal This difference i n thermal con- contraction, which is smaller than Inconel 718.

t r a c t i o n permits t h e Rene' 41 shaft t o have a s l i g h t l y loose f i t ( a t room tem- perature) i n t o t h e Inconel 718 pump r o t o r , yet have an interference f i t a t pump This feature makes assembly less d i f f i c u l t and produces operating temperature.

a r i g i d j o i n t a t cryogenic temperatures, which is e s s e n t i a l f o r s a t i s f a c t o r y power transmission as w e l l as f o r maintaining the close running clearances within t h e pump.

The turbine disc is manufactured from an Inconel 718 for- "his material provides both t h e high strength required and t h e necessary ging.

This material is capable d u c t i l i t y at both cryogenic and elevated temperatures.

of accepting t h e large thermal gradients involved during t h e pump start cycle.

As mentioned previously, t h e turbine components are joined by electron-beam welding. This method provides a very narrow weld and heat a f f e c t e d zone with mechanical properties equal, or nearly equal, t o the parent material. The narrow weld and the necessary metal-to-metal f i t of the parts being joined make possible a j o i n t configuration, which requires t h e least mass of material, thus producing a part of m i n i m u m weight.

All welds are inspected by dye penetrant, radiographic, Sorne The w e l d s were made with a m i n i m of d i f f i c u l t y .

and u l t r a s o n i c methods.

d e f e c t s did occur because of the electron-beam passing out of t h e j o i n t , but defects were removed i n subsequent machining operations, leaving a weld of t h e s e high quality.

A l l machining other than that mentioned above w a s accon- p l i s h e d by conventional methods, page 53 .

b. Manufacturing Problem Area Only one real problem w a s encountered during manufactu- I n i t i a l This was the welding of the cap which closes t h e blade cavity.

ring.

Aerojet-General welding r e s u l t e d i n poor weld penetration and cracked blades.

Welding Jbgineers found the problem w a s caused by: (1) Too much heat was used during the welding.

The surfaces t o be joined were not adequately cleaned.

(2) Poor weld technique was used i n method of weld appli- (3) cation as well as the sequence of operation.

(4) The welding equipment was i n need of repair.

The necessary corrective a c t i o n s were undertaken by t h e Aerojet-General Corporation Welding Engineers and they a l s o supervised t h e a c t u a l welding of t h e caps i n t o each blade. The welding proceeded s a t i s f a c t o r i l y under these conditions .

Electron-beam No other problem a r e a s were encountered.

Full-scale sample welds were made p r i o r t o welding proved t o be trouble-free.

Weld j o i n t preparation, as regards fit joining the a c t u a l turbine components.

and surface f i n i s h , w a s established from t h e weld j o i n t samples.

3 . Second-Stage Rotor and Blades The second-stage turbine r o t o r is e s s e n t i a l l y a disc having 78 blades on its circumference (see Figures No. 6 through No. 8 ) . The blades have a t i p shroud, which is segmented i n t h a t each blade has a section of t h e shroud as an i n t e g r a l piece. The shroud segments are not joined between blades as on the f i r s t - s t a g e turbine. Figure No. 6 shows a cross-section of t h e d i s c and the blade profile.

The blades are reduced i n weight by means of an i n t e r n a l cavity, machined from t h e blade t i p inward. This c a v i t y is covered with a t h i n cap t h a t is welded i n t o place. Blades are machined as i n d i v i d u a l items from Inconel 718 d i e forgings .

The The turbine d i s c is machined from a n Inconel 718 forging.

d i s c cross-sectional shape is t h e d i r e c t r e s u l t of stress a n a l y s i s aimed at pro- a part of minimum weight. One s i d e of t h e disc has a curvic coupling, ducing which serves as the i n t e r f a c e t o t h e f i r s t - s t a g e t u r b i n e rotor. The coupling not only positions t h e turbine r o t o r properly, but a l s o t r a n s m i t s t h e power developed.

The opposite s i d e of the turbine has a hub contour, which has a s u i t a b l e face f o r the turbine tie-bolt i n s t a l l a t i o n and a grooved area t o a i d i n disassembly.

4. Reversing Vanes The reversing vane row is comprised of s i x segments. There are f i v e segments having 1 1 vanes each and one segment having I 2 vanes f o r a t o t a l of 67 vanes per row. Each segment is a welded assembly made up of an The inner shroud is divided s o t h a t i n n e r shroud, vanes, and an o u t e r shroud.

t h e r e a r e two blades per shroud segment. The s l o t forming this division is These s l o t s are required t o stepped s o as t o i n t e r l o c k t h e adjacent blades.

A minimize t h e thermal s t r e s s e s r e s u l t i n g from the expansion of the segments.

honeycomb seal is i n s t a l l e d on t h e leading edge of the inner shroud between the f i r s t - s t a g e turbine r o t o r and the reversing row t o reduce gas leakage t o a mini- Figure No. 12 shows a cross-section through one segment as w e l l zs t h e mum.

vane p r o f i l e and interlocking inner shroud.

The vanes a r e formed sheet metal u n i t s f o r m i n i m u m weight.

The outer shroud segment is a hollow continuous member. A key The vane segment is mounted i n t h e turbine s l o t is provided f o r anti-rotation.

Where necessary, sheet metal cover-plates housing by means of the outer shroud.

are used t o prevent gas flow through a r e a s other than t h e vane passages.

The reversing vane segments were fabricated as a complete ring.

T h i s f a b r i c a t i o n The r i n g w a s cut apart t o form t h e required s i x segments.

method simplified t h e machining, tooling, and alignment of t h e many parts re- quired i n the weld assembly.

The outer shroud w a s machined from an Inconel 718 r5ng forging.

Machining t h e part as a r i n g made it possible t o achieve the tolerances required t o make t h e end product dimensionally acceptable.

The airfoil-shaped holes required t o accent the vane ends were machined on the inner diameter of t h e shroud by e l e c t r i c a l discharge machining, EDM.

The over- The vanes were formed from Inconel 718 sheet metal.

l a p and weld j o i n t were made a t the t r a i l i n g edge of the vane. The f i n a l a i r f o i l surface w a s completed by machining t h e weld deposited material t o t h e desired The inner shroud w a s formed i n t o a complete r i n g using Inconel 718 contour.

The holes required t o accept t h e vane ends s h e e t stock as the basic material.

were pierced by the E D M method.

The two shrouds were f i x t u r e d , t h e vanes i n s e r t e d , and the u n i t Then, the assembly was heat-treated t o the full-aged condition. After welded.

completing a l l remaining machining work, t h e r i n g w a s cut i n t o s i x segments by The EDM electrodes do not a c t u a l l y touch the work piece; therefore, it was EDM.

possible t o use a very t h i n electrode t o produce the shroud s l o t s and segments from a s i n g l e r i n g assembly.

~ Excessive d i s t o r t i o n occurred when t h e f i r s t f u l l r i n g w a s cut i n t o t h e s i x vane segments. Investigation revealed t h a t t h i s d i s t o r t i o n re- s u l t e d from t h e heat-treating operation. The r i n g weld assembly had not been r e s t r a i n e d i n a f i x t u r e during t h e heat treat cycle. A f i x t u r e was constructed The vane segments cut from t h e second and used f o r t h e second weld assembly.

unit exhibited some d i s t o r t i o n , but a l l parts were acceptable f o r use.

5. Rotor Tie-Bolt The function of the tie-bolt is t o hold the two turbine r o t o r s It must withstand the s i g n i f i c a n t thermal movements securely t o t h e pump rotor.

In a d d i t i o n , it must hold t h e f o r c e s from the gyro- of t h e r o t a t i n g components.

scopic gimbaling loads, the a x i a l acceleration loads, and the curvic coupling TO ob- To accomplish t h i s , a very e l a s t i c b o l t was required.

separation loads, t a i n t h i s e l a s t i c i t y , the b o l t was made as long as possible.

A The b o l t has a diameter of 1.38-in. and is 29.25-b. long.

The function and positioning 0.5-in. hole is r i f l e - d r i l l e d through t h e center.

The tie-bolt is threaded i n t o t h e pump of t h i s part is shown in Figure No, 28, r o t o r a t one end. The other end has a nut and lockwasher r e t a i n i n g system, which l o c a t e s against the second-stage r o t o r bore extension. To maintain concentricity with t h e turbine bores and reduce t h e i n i t i a l unbalance, p i l o t i n g shoulders are provided on the tie-bolt at the center and at each end. The center shoulder a l s o damps out the tie-bolt harmonic vibration. The material is Inconel 718 a l l o y , aged f o r high strength, 6. Bearing Housing Seal k f l e x i b l e closure, between t h e hot i n l e t manifold and the cold bearing housing, was required t o contain t h e hot gases i n t h e turbine cavity. The closure had the requirement f o r accommodating l a r g e deflections i n both t h e r a d i a l and axial d i r e c t i o n s and a l s o t o r e t a i n t h e hot gas with zero leakage .

The flexible portion of the closure resembles a t o m 6 cut i n h a l f circumferentially, This w a s formed from 0.040-in. t h i c k 718 material. It is supported by a stiff r i n g on the outside diameter and by a cone frustum on t h e i n s i d e diameter. Figure No. 29 shows t h e bellows and support r i n g s , with a s t a t i c pressure l i n e attached. Three equally-spaced s t a t i c pressure t a p s a r e provided f o r performance measurement.

The bellows is spun from 0.040-in. t h i c k sheet i n t o its t o r - o i d a l shape with extensions protruding f o r attachment t o supporting members.

The outside support r i n g and the i n s i d e support cone frustum are forged from 718 material t h a t is machined t o required dimensions. Close toleranced i n t e r f a c e surfaces have excess m a t e r i a l f o r removal a f t e r welding and heat treatment.

The three subassemblies are GTAW welded together.

The bellows r i n g is supported a x i a l l y and r a d i a l l y by t h e outside diameter and i n s i d e dia- meter support rings; therefore, the GTAW weld is primarily a seal. Flexible hoses and adapters are GTAW welded t o t h e static pressure o u t l e t s . The complete assembly is now heat-treated. To determine mechanical i n t e g r i t y , a hydrostatic proof test at 462 p i g and a helium leak check at 270 psig are accomplished.

I n t e r f a c e dimensions are then finished machined, 7 . Exhaust Cone The exhaust cone is e s s e n t i a l l y an adapter between the turbine diameter upstream flange and t h e test stand exhaust duct., In addition, t h e l a r g e mates with the turbine i n l e t manifold main flange t o form t h e attachment point f o r t h e support frame and clamps, which supports t h e complete turbine assernbly, I n s t m e n t a t i o n p o r t s are provided f o r t o t a l pressure, s t a t i c pressure, and t o t a l temperature i n two axial positions and s i x t a n g e n t i a l positions. Provi- s i o n is made for o r i f i c i n g the downstream end t o control turbine backpressure.

The exhaust cone is a 32 degree cone frustum with flanges at each end. Figure No. 13 shows a cross-section of t h e cone.

The cone frustum is r o l l e d from 0.125-in. Inco 718 sheet and t h e seam is GTAW welded. Flanges f o r each end are rough machined, then GTAW welded t o t h e cone ends. Eighteen instrumentation p o r t s are GTAW welded t o t h e A l l w e l d s are penetrant and radiographically-inspected f o r quality.

cone surface.

The welded assembly is solution-annealed and aged t o obtain a combination of stress relief and maximum material strength.

To determine s t r u c t u r a l i n t e g r i t y , t h e part is proof t e s t e d a t The f i n a l operation consists of f i n i s h 425 psig and leak t e s t e d at 200 p i g , machining the flanges and indexing surfaces t o the f i n a l dimensions.

G T A W welding presented some i n i t i a l problems of gas porosity T h i s was corrected by improved cleaning, voids and small cracks i n t h e weld area.

b e t t e r i n e r t gas coverage, and smaller weld passes with cooling between passes.

8. D u a l E x i t Exhaust Housing The dual exit exhaust housing is used when t h e f u e l turbopump Figure No. 14 assembly operates as part of the complete M - 1 engine system.

shows a cross-section of t h e exhaust housing and Figure No. 1 shows it assembled w i t h adjacent components.

The exhaust housing is a 26.5-in. diameter hemisphere with a flange on the open end t h a t mates with t h e turbine i n l e t manifold main flange.

This forms t h e attachment point f o r t h e support frame and clamps which support t h e complete turbine assembly. Two 90 degree elbows of 11.4-in. diameter, spaced 120 degrees apart, protrude from the hemisphere-shaped exhaust housing.

These elbows provide attachment points f o r the two cross-over ducts, which carry The 4.63-in. diameter t h e exhaust gas and by-pass gas t o t h e oxidizer turbine.

hot gas by-pass port and gas deflector is placed i n t h e center of t h e hemispheri- c a l end of t h e exhaust housing.

The exhaust housing s h e l l and protruding elbows a r e formed from 0.20-in. thick Inconel 718 sheet material. All flanges a r e machined from Inconel 718 r i n g forgings, and GTAW welded t o t h e s h e l l .

All flanged j o i n t s use double conical seals, except t h e l a r g e The main flange is sealed by a narrow (.060-in.) circum- diameter main flange.

f e r e n t i a l GTAW weld.

The 26.5-in. diameter hemisphere is formed from a f l a t plate.

The 90 degree t r a n s i t i o n elbows a r e formed i n t o 180 degree half-shells, then Cutouts i n the hemisphere are reinforced by welded together t o form the elbow.

Flange hubs, slope gradually (15 degrees t o 17 degrees) t o a t r a n s i t i o n rings.

smooth t r a n s i t i o n with t h e housing w a l l . A l l forgings are u l t r a s o n i c a l l y in- spected. Excess material remains on all flange surfaces t o allow machining t o f i n a l dimensions a f t e r welding, heat-treatment, and proof t e s t .

A f t e r a l l parts a r e welded together, the assembly is solution- annealed and aged, which stress r e l i e v e s t h e part and brings t h e material s t r e n g t h up t o desired l e v e l s e The part is h y d r o s t a t i c a l l y proof t e s t e d at 560 p s i g and helium leak t e s t e d at 275 psfg. If proof and leak t e s t s a r e s a t i s f a c t o r y , The t h e part is penetrant-inspected and welds a r e radiographically-inspected.

assembly is then machined t o f i n a l dimensions.

Support Frame and Main Flange Clamps 9.

The support frame extends from t h e outside diameter of the main flange t o the pump discharge housing. The frame c o n s i s t s of three conical segments, with 72 degrees of a r c t o each segment. The segments are free t o move The r a d i a l l y t o accommodate t h e thermal movements of the t u r b i n e main flange.

turbine i n l e t t o r u s and the exhaust housing a r e a l s o free t o move, both r a d i a l l y fixed point. The frame material is Inconel and a x i a l l y , from the main flange 718, which gives lightweight with high strength. It a l s o provides the same CO- e f f i c i e n t of thermal expansion as the turbine components.

The support frame, clamps, and turbine main flange can be seen The frame is made from Inconel 718 i n t h e i r assembled position i n Figure No. 2.

Flanges a r e sheet stock, 0.188-in. thick, r o l l e d i n t o a 20 degree cone frustum.

Figure No. 9 shows a s i n g l e welded t o each end forming the attachment points.

segment of the frame with an end view and a cross-section.

The main flange end of the frame segments are s l o t t e d a x i a l l y for 6-in. and then every 2.3-in. in t h e t a n g e n t i a l direction. This gives t h e required r a d i a l f l e x i b i l i t y f o r the turbine thermal expansions. Two 1/2-in.

high circumferential reinforcing r i b s are welded t o the conical section t o give t h e required s t i f f n e s s t o t h e segments during handling and i n s t a l l a t i o n . The It is then cut i n t o f i v e seg- frame is fabricated as a complete cone frustum.

All heat treatment ments, three of which are used i n each turbopump assembly.

and f i n i s h machhing is accomplished before t h e frame cone is cut i n t o segments.

The clamps that mate with t h e frame ends are machined from a forged r i n g that approximates t h e finished shape. Material is Inconel 718, aged Between t h e frame segments, two clamps, by heat-treating f o r maximum strength, one on each s i d e of t h e turbine main flange, are bolted together giving con- tinuous support t o the flanged joint.

IV . CONCLUSIONS AND RECOMMENDATIONS

The electron-beam welding of the f i r s t - s t a g e r o t o r shaft and the blades of both r o t o r s t o the d i s c s w a s a s a t i s f a c t o r y method f o r attaching these parts It a l s o lessened the r o t o r weight w i t h a minimum of d i s t o r t i o n and machining.

r i m thickness, as compared with mechanical attachment methods.

because of reduced were required on t h e blades and very complex However, very close tolerances w a s needed t o position the blades i n preparation f o r the welding process.

tooling a small development programe it would be less expeneive t o use simple tooling For and hand-fit each blade, 718 material, which was f a i r l y new at t h e time, caused The use of Inconel some welding problems when using t h e gas tungsten arc welding method. The main problem w a s weld contamination ( L e e , occlusions i n the weld rod, impurities i n t h e helium and argon gas shields, and dusty o r d i r t y weld booths). Additional problems encountered included weld psses that were too l a r g e with improper pass causing weld cracking, part d i s t o r t i o n , and weld oc- cleaning between each It w a s found that optimum weld j o i n t configurations had t o be de- clusions.

This required t h a t sample weld j o i n t s be made t o veloped f o r each component.

as w e l l as the equipment.

obtain the best solution and t o check out t h e welder The design achieved t h e t a r g e t weight goal of 960 l b with a l l s t a t i c parts successfully passing proof t e s t requirements. However, the turbine w a s not operated at design conditions.

The separable, seal-welded, housing j o i n t s proved f e a s i b l e and j o i n t s of this type are recommended f o r hot gas flange applications because t h e seal welded j o i n t has less w e i g h t and less thermal stress than a comparable j o i n t with con- ventional seals.

BIBLIOGRAPHY Bartholf, Lo W., S t r u c t u r a l Analysis of M - 1 , Mod I FTPA Turbine Tie-Bolt 1 .

P/N 286134, Aerojet-General Report No, SA-FTPA-164, 24 June 1965 Bartholf , L . W . , H i l t z , J . P. , and Smithers, 0. L., S t r u c t u r a l Analysis 2.

of t h e M - 1 Fuel Turbine Rotors, Aerojet-General Report No. FTPA-102, 20 November 1964

Bartholf, L . W e and Smola, C. R. , S t r u c t u r a l Analysis of FTPA Turbine

3.

Rotor P/N 286109 f o r M-1 Application, Aerojet-General Report No.

FTPA-124, 13 July 1965 Blakis, R., Lindley, B, K., R i t t e r , J. A., and Watters, W. E., I n i t i a l 4.

Test Evaluation of t h e M - 1 Liquid Hydrogen Turbopump Including I n s t a l - l a t i o n , Test Procedures, and Test iiesults, N A S A Report CR 54827, 20 J u l y 1966

Frick, V., Hunt, V , , Inouye, F, T., and Janser, G. R e , Summary of a-

5.

erience Using Inconel 718 on M - 1 Engine, Aerojet-General Report No.

$800-37, 1 March 1966 (To be published as a N A S A Contractor report) Goudreaus G. L e , Smithers, 0. L., M - 1 Fuel Test Manifold, Test Manifold 6.

Brace K i t , and Turbine I n l e t Manifold, Aerojet-General Report No.

SA-FTPA-127, 2 1 April 1965 S t r u c t u r a l Analysis of PI-1 Fuel Turbine H i l t z , J . P o , Smithers, 0 . L.

I n l e t Manifold with Nozzle Assembly, Aerojet-General aeport No. FTPA-SA- 129, 9 March 1965

Radkowski, Po P. ~ Davis, R. M e , and Boldul, M e R, , A Numerical Analysis

8.

of the Equations of T h i n S h e l l s of Revolution, Avco Corporation Mechanical Desim of the M - 1 Axial Flow Hydrogen Fuel b e , Regan P. J.

9.

N A S A Report CR 54823, 15 February 1966 Aerodynamic Desim, Model I1 Turbine M-1 Fuel Turbopume 10, Reynolds, T. W e Assembly, NASA Report CR 54820, 15 April 1966 1 1 . Severud, L. K . , S t r e s s Analysis of the M-1, Mod I1 Turbine Exhaust Housing, P/N 286144, Subjected t o I n t e r n a l Pressure and Line Induced -ding, Aero jet-General Report No. FTPA-112, 31 December 1964 12. Severud, L. K O , S t r e s s Analysis of the M - 1 , Mod I1 Turbine I n l e t Manifold - Exhaust Housing J o i n t , Aerojet-General Report No. FTPA-120, 31 December 1964 Page 60 BIBLIOGRAPHY (Cont d) Smithers, 0. Le, Structural Analysis of M-1 Fuel TPA Bearing Housing Seal 1 3 .

Assembly P / N 286116, Aero jet-General Report No. SA-FTPA-174, 26 July 1965

Toms, R. M . , M-1 Fuel Aunp Turbine Blade Resonant Speeds , Aero jet-General

14 0 Report No. FCPA-108, 2 December 1962 Sponseller, R. Lo and Tom, R. M . , M-1 Fuel 3mp Turbine alade Stress Analysis, Aero jet-General Report No. F T P A - ~ ~ ~ , 31 December 1964 ASME, Section 111, ASME Boiler and Pressure Vessel Code for Nuclear 16 .

Vessels, pe 6 , Library of Congress, Catalog Card No. 56-3934, 1963 .