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USAAVLABS TECHNICAL REPORT 64-68K
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HEAVY-LIFT TIP TURBOJET ROTOR SYSTEM VOLUME XI
ENGINE DESIGN
October 1965
I. S. HIT IIIITIII IITEIIEl IIIIIITIIIES
Fill EISTIS. fllfillll
CONTRACT DA 44-177-AMC-25(T)
HIUER AIRCRAFT COMPANY, INC CONTINENTAL AVIATION AND ENGINEERING CORPORATION
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Disclaimers
The findings in this report are not to be construed as an official Depart- ment of the Army position, unless so designated by other authorized documents.
When Government drawings, specifications, or other data are used for any purpose other than in connection with a definitely related Government pro- curement operation, the United States Government thereby incurs no re- sponsibility nor any obligation whatsoever; and the fact that the Govern- ment may have formulated, furnished, or in any way supplied the said drawings, specifications, or other data is not to be regarded by implication or otherwise as in any manner licensing the holder or any other person or corporation, or conveying any rights or permission, to manufacture, use, or sell any patented invention that may in any way be related thereto.
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Destroy this report when it is no longer needed. Do not return it to the originator.
Task 1M121401D1441?. Contract DA 44-177-AMC-25(T)
USAAVLABS Technical Report 64-68K Oktober 1%S
HEAVY-LIFT TIP TURBOJET ROTOR SYSTEM
VOLUME XI
ENGINE mmcn
CAE Report No. 942
Prepared for
Hiller Aircraft Company, Inc.
By
Continental Aviation and Engineering Corporation
For
U. S. ARMY AVIATION MATERIEL LABORATORIES FORT EUSTIS, VIRGINIA
(ff. S. Anqr Tmspoitatu» Research Ommmi «hm r«port inmparoil
PREFACE
This report, prepared by Continental Aviation and Engineer - ing Corporation covers the design of the Continental Model 357-1 tip turbojet engine. This work was accomplished under Hiller purchase order HAC-l-64, Contract DA 44-177-AMC-25{T).
The Continental Model 357-1 tip turbojet is a modification of the J69-T-29 turbojet engine (CAE Model 356-7) and is designed to operate at the tip of a helicopter rotor.
This report is one of a series of three documents submitted to fulfill the requirements of Hiller purchase order HAC-l -64, Article 1 Statement of Work, Item (3) (e). The reports include:
"Continental Model 357-1 Tip Turbojet Engine - Engine Design, " Heavy-Lift Tip Turbojet Rotor System, Volume XI, CAE Report No. 942, U.S. Army Transportation Research Command * Fort Eustis, Virginia, October 1965.
"Continental Model 357-1 Tip Turbojet Engine - Fuel Pump and Control System Design, " Heavy-Lift Tip Turbojet Rotof System, Volume XII, CAE Report No. 943, U.S. Army Transportation Re- seat ch Command^ Fort Eustis. Virginia. October 1965.
"Continental Model 357-1 Tip Turbojet Engine - Preliminary Model Specification. " Heavy-juft Tip Turbojet Rotor System, Volume XIU. CAE Model Specification No. 2253. U.S. Army Transportation Research Command.*Fort Eustis, Virginia, October 1965.
♦Changed to U. S. Army Aviation Materiel Laboratories in March 1965.
in
CONTENTS
Page
PREFACE iii
LIST OF ILLUSTRATIONS vii
LIST OF SYMBOLS xiii
SECTION ONE. SUMMARY 1
SECTION TWO. INTRODUCTION 5 Operating Environment 5
SECTION THREE. MECHANICAL DESIGN 11 Aerodynamic Flow Path 11 Rotor System 11 Static Structure 38 Accessory Drive 49 Seals 55 Engine Weight and Center of Gravity 57
SECTION FOUR. SYSTEMS DESIGN 61 Fuel System 61 Lubrication System 64 Starting System 7 3 Anti-Icing 76 Electrical System 78
SECTION FIVE. AEROTHERMODYNAMIC DESIGN 80 Summary 80 Discussion 81
DISTRIBUTION 99
ILLUSTRATIONS
Figure Page
1 Continental Model 357-1 Tip Turbine Engine Mockup . 1
2 Continental Model 357-1 Tip Turbojet Performance Characteristics 2
3 Continental Model 357-1 Tip Turbojet Engine Installation Drawing 4
4 Continental Model 357-1 Tip Turbojet Engine Cross Section 6
5 Engine Load Directions 8
6 Forces Acting Upon Gas Turbine Rotating Elements . . 9
7 Aerodynamic Flow Path - Continental Model 357-1 . . 12
8 Rotating Assembly - Continental Model 357-1 13
9 Typical Blade Stress - Continental Model 357-1 Engine 19
10 Axial Compressor Blade Stress Summary - Continental Model 357-1 Engine 20
11 Axial Compressor Blade Interference Diagram - Continental Model 357-1 Engine 21
12 Modified Goodman Diagram for Continental Model 357-1 Axial Compressor Blade 22
13 Inducer Blade Stress Summary - Continental Model 357-1 Engine 24
14 Inducer Blade Interference Diagram - Continental Model 357-1 Engine 25
vii
Figure Page
15 Modified Goodman Diagram for Continental Model 357-1 Inducer Blade 26
16 Radial Compressor Radial Blade Stress Summary Continental Model 357-1 Engine 27
17 Radial Compressor Rotor Blade Interference Diagram - Continental Model 357-1 Engine 28
18 Radial Compressor Tangential Stresses - Continental Model 357-1 Engine 29
19 Radial Compressor Radial Disc Stresses - Continental Model 357-1 Engine 30
20 Modified Goodman Diagram for Continental Model 357-1 Radial Compressor Blade and Disc 31
21 Turbine Blade Stress Summary - Continental Model 357-1 Engine 32
22 Turbine Blade Interference Diagram - Continental Model 357-1 Engine 33
23 Modified Goodman Diagram for Continental Model 357-1 Turbine Blade 34
24 Turbine Disc Steady-State Stresses - Continental Model 357-1 Engine 36
25 Modified Goodman Diagram for Continental Model 357-1 Turbine Disc 37
26 Continental Model 357-1 Engine Static Structure .... 40
27 Continental Model 357-1 Engine Front Bearing Support Frame Stress Summary 45
28 Continental Model 357-1 Engine Support Ring Stresses 46
29 Continental Model 357-1 Compressor Cover Stresses . 47
vui
Figure Page
30 Continental Model 357-1 Rear Bearing Support Stresses 48
31 Continental Model 357-1 Accessory Drive Train. ... 51
32 Continental Model 357-1 Bearing Locations 54
33 Continental 357-1 Engine Oil Seal Balance Pressure . 56
34 Continental Model 357-1 Rear Bearing Oil Seal .... 56
35 Continental Model 357-1 Fuel Pump and Fuel Control Labyrinth Seal 57
36 Continental Model 357-1 Engine Starting Fuel System . 61
37 Continental Model 357-1 Engine Main Fuel System 63
38 Continental Model 357-1 Engine Fuel Tube Seal .... 64
39 Schematic Diagram of Continental Model 357-1 Lubrication System 65
40 Estimated Heat Rejection, Oil Flow Rate, and Temper- ature Rise - Continental Model 357-1 Engine 66
41 Continental Model 357-1 Lubrication System Charac- teristics Versus Engine Speed 67
42 Continental Model 357-1 Lubrication Oil Tank Capacity 68
43 Estimated Pressure Drop in Oil Distribution Lines - Continental Model 357-1 Engine 71
44 Estimated Pressure Drop in Scavenge Lines - Continental Model 357-1 Engine 72
45 Continental Model 357-1 Air Starting System 74
46 Continental Model 357-1 Axial Compressor Starting Characteristics 75
ix
Figure Pag<
47 Continental Model 357-1 Estimated Windmill Starting Envelope 76
48 Continental Model 357-1 Engine Inlet Anti-Icing System 77
49 Schematic Diagram of Continental Model 357-1 Electrical System 78
50 Continental Model 357-1 Engine Cross Section 83
51 Performance Comparison - Baseline and Two-Stage Diffuser 84
52 Two-Stage Radial Diffuser Configuration 85
53 Thermodynamic Cycle and Entropy Diagram 86
54 Estimated Compressor Performance - Continental Model 357-1 Engine 86
55 Estimated Engine Time Constant Versus Engine Speed - Continental Model 357-1 Engine 87
56 Inlet Screen Configuration 89
57 Instrumentation of Air Inlet for Distortion Test ... . 90
58 Schematic Diagram of Inlet Distortion Test Instrumentation 91
59 Installation of Static Pressure Probes and Total Pressure Rakes on Adapter Ring - Model 357-1 Inlet Distortion Test 91
60 Installation of Adapter Ring - Model 357-1 Inlet Distortion Test 92
61 View of Distortion Screen in Model 357-1 Engine Irlet. 93
~VPHR>P"LII " immsm** rr*—-mmm^m
Figure Pag«
62 Model 357-1 Engine Inlet Distortion - Screen Pattern "A" 94
63 Model 357-1 Engine Inlet Distortion - Screen Pattern "B" 95
64 Parameter Change Versus Inlet Distortion at Design Speed - Model 357-1 Engine 97
65 Effect of Inlet Distortion on Thrust - Model 357-1 Engine 98
66 Effect of Inlet Distortion on Fuel Flow - Model 357-1 Engine 98
n
SYMBOLS
u) engine speed
& helicopter rotor speed - radians per second
0 rotor blade pitching velocity - radians per second
g acceleration field
cr stress - pounds per square inch
Fn net thrust - pounds
Wf net thrust - pounds per hour
Wa airflow - pounds per second
^z/^i pressure ratio
Xlll
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SECTION ONE. SUMMARY
This report presents the design of the Continental Model 357-1 tip turbojet engine designed under Hiller Contract HAC-1-64.
The Model 357-1 is a modification of the J69-T-29 engine (Continental Model 356-7). designed to operate at the tip of a helicop- ter rotor.
Figure 1 shows the basic arrangement of the 357-1 engine. It incorporates a straight-through flow path with an integral jet nozzle with the minimum number of accessories necessary to operate the en- gine and features an integral oil tank and optional oil cooler, air start- ing, and provisions for an anti-iced intake.
RIOHT SIDE VIEW
Figure 1. Continental Model 357-1 Tip Turbojet Engine Mockup.
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The static sea level performance targets for the fully qualified production Model 357-1 engine are summarized below:
Condition
Military
Normal
Pounds Thrust
1700
1375
SFC
0.99
0.98
The aerothermodynamic design of the engine is based on the proven J69 engine with an airflow of 29. 5 pounds per second, 5. 7 pressure ratio, and a turbine inlet temperature of 1560oF. The per- formance of the engine is shown in Figure 2.
2000
1800
= «< 22,000
20,790
20,130
19,140
17,300
9, 500 (IDLE)
0 200 400 600 800 TRUE FLIGHT SPEED - V - KNOTS
Figure 2. Continental Model 357-1 Tip Turbojet Performance Characteristics.
The complete engine, including oil tank and cooler, is 47. 95 inches long with a maximum diameter of 27. 80. The installation drawing is shown in Figure 3.
Preliminary design of the fuel control and fuel pump has been accomplished primarily to meet the requirements of, and to provide the necessary response and safety features for, a tip turbojet engine in the hover mode of operation.
Sufficient growth potential is available in the basic engine as both the low cycle temperature and the pressure ratio of the axial compressor can be increased. Increase in cycle temperature with the present compressor design would enable the engine to easily achieve a design point thrust of 1840 pounds at a specific fuel consumption of i. 1. A slightly modified axial compressor and a new combustor con- figuration will enable the engine to be rated at 1900 pounds thrust at 1.12 specific fuel consumption. Both of these aerodynamic configu- rations have already been tested and the performance verified. Fur- ther increases in performance could be realized by aerodynamic de- velopment of the centrifugal compressor and a new axial compressor designed for a higher pressure ratio and mass flow. Without adding stages, it should be possible to develop this basic engine to 2000 pounds of thrust at approximately 1. 1 specific fuel consumption. These modifications plus the possibility of adding stages, which would give as high as a twenty-five percent increase in thrust, should en- able the basic engine to meet higher horsepower requirements.
SIDE VIEW
Figure 3. Continental Model 357-1 Tip Turbojet Engine Installation Drawing.
SECTION TWO. INTRODUCTIO?:
The CAE Model 357-1, Figure 4, is a turbojet engine rated at 1700 pounds static sea level thrust, designed specifically for rotor tip application.
The engine employs the straight-through aerodynamic flow path of the Continental J69-T-29 engine. Air enters an annular intake at the front of the engine and passes successively through a single- stage transonic axial compressor, a two-stage stator, a single-stage centrifugal compressor, radial and axial diffusers, an annular com- bustor, a turbine inlet nozzle, a single-stage turbine, and an integral fixed-area jet nozzle.
The engine utilizes a maximum diameter structural shell with the two primary mounting points near the center of gravity and a steady rest mount at the rear of the engine.
The engine employs a circulating lube system with an integral oil tank and optional engine-mounted oil cooler. The lube system is designed to operate in the lg vertical field static and a 235g horizon- tal field peculiar to tip turbojet operation.
OPERATING ENVIRONMENT
The operating environment for a turbojet engine mounted at the tip of a helicopter rotor is unique. The engine is subjected to continuous loading of considerable magnitude, both static and alterna- ting, not encountered in any other application. The three basic load directions are shown in Figure 5. The direction "R" is normal to the helicopter rotor axis and directed radially outward, and is the di- rection of the predominant load. The direction "F" is tangent to a circle drawn through the tip of the helicopter rotor blades. The di- rection "V" is parallel to the helicopter rotor axis. There is a con- tinuous g field in the "R" direction with a superimposed alternating g field. The g fields in the "F" and "V" directions are alternating. A summary of the g fields in the three basic directions is presented in Table 1.
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TABLE 1 1 TIP TURBOJET g FIELD SUMMARY
HOVERING Percent
Direction Continuous Alternating Frequency* of Time
R 235 + 10 I 5 + 20 0.2 2
1 1 6 1
F + 5 1 5 + 2 + 3
0.2 6
2 1
! v i + 40 1 5 + 20 1 2
+ 10 0.67 1
CRUISE Percent
Direction Continuous Alternating Frequency* of Time
R 195 + 10 1 5 + 20 0.2 2 + 1 + 5
6 1
1 50
| F + 5 1 5 + 2 0.2 2 + 3 6 1
1 3 3 50
V 1 + 40 1 5 + 20 1 2
+ 10 0.67 1 7 15 2 49 + 25 2 1 1
* Per revolution of helicopter rotor.
mm»
Figure 5, Engine Load Directions.
The hovering g field is applicable for 7 percent of the opera- ting time and the cruise g field for 93 percent of the operating time. The engine is capable of an overspeed to 259g ior a total time of 30 minutes per 1, 000 hours of total flight time with one minute per oc- currence.
The angular velocity summary is presented as follows:
In plane velocity Q
Hovering flight (7 percent):
Normal operation 10.8 radians per second. Overspeed 11.3 radians per second for 1 minute per occurrence for a total of 30 minutes per 1, 000 hours of total flight time.
Cruise flight (93 percent):
Normal operation 9. 8 radians per second. Harmonic variation of + 0, 25 radians per second at a frequency of 1 per helicopter rotor revolution for 50 percent of the cruise flight time.
Pitching velocity ♦
50 percent of cruise flight:
Maneuve ring flight:
+ 1 radian per second at a fre- quency of 1 per helicopter rotor revolution.
+ 3 radians per second at a fre- quency of 1 per helicopter rotor revolution for 2 percent of the total flight time.
The complete operating environment specified in MIL.-E-5007B are applicable except that the g fields and operational attitudes pecu- liar to operation at the tip of a helicopter rotor shall apply.
The operation at the tip of a helicopter rotor results in high rotor bearing loads due to the centrifugal field, a continuous gyro- scopic couple, and an alternating Coriolis force parallel to the engine rotor axis. A simplified diagram illustrating these forces are shown in Figure 6. Due to rotation u about the engine axis, any point "P"
ENGINE ROTOR AXIS HELICOPTER ROTOR AXIS
Figure 6. Forces Acting Upon Gas Turbine Rotating Elements,
np*
knows a constant velocity Vj and a constant centrifugal force CFj. Due to rotation 0 about the helicopter axis, the same point "P" knows a variable velocity V2 and a variable centrifugal force CF2. Due to the variation in velocity V2 in the direction parallel to the engine axis, point "P" knows an acceleration force FC. commonly known as the Coriolis force. The summation of the Coriolis forces produce a gyroscopic couple with resulting reactions Fj and F2' The Coriolis force, FC, changes direction above and below the "OR" axis. The alternating Coriolis force, FC, produces alternating stresses in all rotating blades, discs, and shafts with a frequency of one cycle per engine revolution.
10
^äjgj.
SECTION THREE. MECHANICAL DESIGN
AERODYNAMIC FLOW PATH
The Continental Model 357-1 engine incorporate« a straight- through aerodynamic flow as shown in Figure 7. The flow path incor- porates a minimum number of parts and the least amount of aerody- namic obstructions to ensure minimum cost and maximum perform- ance. Air enters the engine axially through a cast aluminum housing (No. 1, Figure 7). From the inlet housing the air passes through a 1.6:1 pressure ratio transonic axial compressor rotor (2), which is machined from a titanium forging. The air from the axial compressor is diffused through a two-stage, precision-cast aluminum stator (3). From the stators the air is directed into a centrifugal compressor (4), which is composed of a machined titanium inducer and a machined titanium radial impeller. From the centrifugal compressor stage the air passes through a fabricated stainless steel radial (5) and axial (6) diffuser. From the axial diffuser the air flows through the combustor housing (7) to the annular combustor (8) where fuel is added and com- bustion takes place. From the combustor the hot gases enter the tur- bine inlet nozzle (9) which is fabricated from precision-cast vanes and forged, fully machined shrouds. The gases then pass through a forged, fully machined turbine (10). From the turbine the gases then exit through the short integral jet nozzle (11).
ROTOR SYSTEM
The Model 357-1 rotor group, Figure 8, has two independently supported shaft assemblies, coupled by a relatively flexible extension of the turbine shaft.
The compressor shaft assembly consists of an axial compres- sor (1) and inducer (2) weldment, on which a radial impeller (3) is piloted and attached. The axial compressor-inducer shaft is made of two parts and electron beam welded. The intermediate section be- tween the compressor and inducer has been aerodynamically contour- ed to form the inner boundary of the airflow path. The axial compres- sor is cantilevered from the shaft and extends partially over the front bearing journal (4).
11
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ZI
c n
> ft H o a •< s
o
P)
ET
i
n o 3 ri-
3 <& 3
O a n >—•
Figure 8. Rotating Assembly - Continental Model 357-1.
The turbine shaft assembly consists of a turbine rotor (11) and intermediate shaft (12) electron beam weldment to which a shaft exten- sion (13) is bolted and piloted. The shaft extension is coupled to the inner hub (15) of the compressor shaft. The extension is designed to transmit torque and axial load from the turbine to the compressor shaft and to be relatively flexible in bending. The turbine shaft exten- sion also serves as a fuel transfer tube, the fuel slinger (14), and the turbine shaft forward bearing journal. The rear turbine shaft journal is an extension of the turbine rotor.
ROTOR SUSPENSION
The compressor shaft assembly is straddle-mounted between two roller bearings. The rear bearing (6) is mounted on the radial compressor extension (5) and is housed in the forward section of the center support housing. The front bearing (7) is axially located so the compressor shaft radial load is divided approximately equal be- tween the front and rear bearings.
13
The engine thrust bearing (8) is mounted on the compressor shaft journal, forward of the front radial bearing. Both the thrust bearing and the compressor shaft front bearing are mounted in a steel housing, which is piloted and bolted to a cantilevered section of the front bearing support. The thrust bearing mounting may be made with* a controlled radial spring rate by cutting a number of longitudinal slots in the steel housing. The soft suspension of the thrust bearing will ensure proper load sharing between the two front bearings; that is, the thrust bearing will absorb all the axial load and only a small portion of the radial load.
The turbine shaft assembly is supported at the center and rear by roller bearings and at the front by the compressor shaft. The for- ward turbine shaft bearing (9) is housed in the rear section of the cen- ter support housing. The rear bearing (10) is housed in the engine rear bearing support.
The roller bearings are mounted in spherical housings that per- mit the bearings to align themselves in the direction of the deflected or misaligned shaft. This mounting arrangement prevents edge loading the rollers and minimizes prestressing the shafts.
The main shaft bearings have been specifically designed for the helicopter rotor tip application. The roller bearings are manufac- tured from consumable electrode vacuum melt (CEVM) M-50 tool steel with a minimum hardness of Rockwell C-(0. The thrust bearing is manufactured from CEVM 52100 steel with a minimum hardness of Rockwell C-58. The use of vacuum melt results in higher quality steels. M-50 tool steels have hardness qualities superior to normal 52100 steels, particularly in high temperature applications. These improved quality and/or hardness properties have resulted in bearing steels with fatigue life several times that calculated for standard bear- ings. This increase in life expectancy is indicated in the bearing load- life tabulation.
The rotor thrust bearing is a 60 mm bore, split inner race, radial contact groove ball bearing. The bearing has a complement of 14 balls, 5/8-inch in diameter, having a normal operating contact angle of 25 degrees. The inner and outer race curvatures are 52 and 53 per- cent, respectively. The bearing has a special flanged outer race that is piloted and clamped in the housing. The single-piece bearing cage is machined from AMS 6415 steel, heat treated to Rockwell C-34 maxi- mum, and silver plated 0. 001- to 0. 002-inch thick. The cage is
14
outer-race land piloted. The pockets of the cage arc flatted in the axial direction to minimixe Coriolis loading by the balls. The ball component is trapped in the cage by deformed tangs on the outer diam- eter. The cage~ball assembly is removable from the outer race. The front half of the inner race has a puller groove. The gearing is manu- factured to ABEC-S standards.
Three of the rotor radial bearing are 60 mm bore roller bear- ings and one is a 45 mm bore roller bearing. These bearings have wide inner rings without lands. The rollers are accurately guided by lands on the outer race. The bearings have a one-piece, machined, silver plated, steel cage which is piloted by the inside diameter of the outer race lands. The rollers, cage, and outer race are »«inseparable. The outside diameter of the outer race is ground spherical to close dimensional and concentricity control. The 60 mm bore roller bear- ings have twelve 20 mm rollers and the 45 mm bore roller bearing has ten 19 mm rollers. The bearings are manufactured by RBEC-5 standards,
ROTOR BEARING ANALYSIS
The tip turbojet environment produces a severe, but not wholly unknown, combination of bearing loads. The three types of loads are:
1. Radial Loads 2. Continuous g Field Operation 3. Continuous Gyroscopic Forces
Current aircraft engines with planetary or epicyclic reduction gearing have bearings operating under radial loads in a continuous g field. The g field in some cases is greater than the tip turbojet en- vironment.
Current aircraft engine bearings are subjected to intermittent gyroscopic forces resulting from aircraft flight and ground maneuver turns. The gyroscopic forces result in overturning couples on the bearing retainer and rolling elements. The effects of these gyroscopic couples, although not amenable to analytical evaluation, should not prove insurmountable to develop.
The main bearing radial reactions are the rest 't of the accel- eration forces on the engine rotor in the radial g field and the gyro- scopic field. These reactions, centrifugal and Coriolis, act in planes
15
90 degrees to each other. The steady-state radial loads have been cal- culated on a duty cycle basis of 93 percent operation at normal rating (20, 790 engine r. p. m. and 640 feet per second helicopter rotor tip velocity) and 7 percent operation at take-off or hover (22. 000 engine r. p. m. and 700 feet per second helicopter rotor tip velocity). The equivalent load is based upon the above duty cycle in proportion to the roller bearing life function.
The basic dynamic load ratings of the roller bearings have been calculated using AFBMA standards. The basic load rating is 30, 200 pounds per million revolutions for the 60 mm roller bearing and 23, 200 pounds per million revolutions for the 45 mm roller bearing.
The thrust bearing load is assumed to be constant (1270 pounds) at both normal and take-off rated conditions. The bearing load is as- sumed to be pure thrust and, therefore, the equivalent load is 1270 pounds. The basic dynamic load rating is 10, 580 pounds per million revolutions, using AFBMA standards.
The equivalent rotor speed of 20,860 r. p. m. has been calcu- lated as the sum of the products of time and speed at take-off and time and speed at rated engine power.
Table 2 presents the load and life summary for the main shaft bearings.
16
TABLE 2 ] MODEL 357-1 MAIN SHAFT BEARING SUMMARY
Function Thrust
Bearing
Front Comp. Roller
Bearing
Hear Comp. Roller
Bearing
Forward Turbine Roller
Bearing
Rear Turbine Roller
Bearing j
Cent. Load at 235g
3842 3902 2500 3321 j
Cent. Load at 195g
3188 3238 2075 2756
' Equiv. Cent. 1 Load
3233 3284 2105 2796 I
I Gyro. Load at 235g
2278 -2533 2083 -1827
Gyro. Load at 195g
1971 -2191 1801 -1580
Equiv. Gyro. Load
1992 -2215 1821 -1597 1
Equiv. Load (pounds)
1270 3797 3961 2783 3220
Materials j Life Factor
5 10 10 10 10
BJO Life Rating (hours)
2315 7980 6950 9350 13,720
17
W m^w
ROTOR STRUCTURAL ANALYSIS
The Model 357-1 engine rotor is subject to all the forces nor- mally encountered in a jet engine plus the acceleration forces pro- duced by the engine masses rotating about the helicopter rotor centerline.
The engine rotor is subject to two significant primary stresses normally encountered in any jet engine. These stresses are:
1. The centrifugal stresses due to rotation about the engine ?.xis.
2. An alternating stress due to circumferential vari- ations in aerodynamic loadings.
The frequency of these alternating stresses is generally a func- tion of strut and stator vane spacing adjacent to the rotating part. Ro- tation about the helicopter rotor centerline produces two significant accelerations, which are termed secondary in order to differentiate them from the normal engine acceleration forces. These accelera- tions are:
1. The centrifugal field.
2. The Coriolis or gyroscopic forces and moments.
The centrifugal field is relatively constant and is directed ra- dially outward from the helicopter rotor centerline. The engine rotor blading changes orientation in the centrifugal field once per engine revolution and experiences an alternating stress.
The Coriolis forces are variable in magnitude and are directed parallel to the engine axis. The rotor blading, discr, and shafts ex- perience an alternating stress due to these alternating forces.
The two secondary stresses are alternating in nature with a frequency of once per engine revolution.
A typical blade stress graph is shown in Figure 9, illustrating the mean blade stress, the cyclic gas bending, and the cyclic Coriolis. The mean stress shown is a summation of the P/A stress due to the primary centrifugal field and any steady-state gas bending stresses. The effects of blade twist and the restoring force due to the primary
18
IBSa«g^
| J- CYCUC GAS BENDING
\ Lr MEAN STRESS PLUS MAXIMUM CYCUC"
>
\ /T psA^a
WS&k. ||. CYCUC CORIOUS PLUS
w uunu'' ^^iiiiftt L SECONDARY CENTRIFUGAL i i i i i i i i i i
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EäflKkJ K-ä &''?-^H^K^ä p 7 ^ M MEAN -
"" STRESS " r "^g ■^-^^^MKES wm r
___t:" IM JC | MEAN STRESS MINUS MAX UM CYCJ 7
LL 1 1 1 i 1 1 i 1 1 1 1
- Ü [ u ■KJtM.1 u i»*- &%£• T ^ r&JU *** u u TIME
Figure 9. Typical Blade Stress - Continental Model 357-1 Engine.
centrifugal field has been fully considered in the blade stress an- alysis.
Particular emphasis has been placed upon providing generous airfoil root fillets. These fillets have been designed to conform to an ideal elliptical shape to allow a smooth load transition from the blade to the rim with low stress concentration.
Axial Compressor
The axial compressor is an integrally bladed rotor, fully ma- chined from a titanium forging. The blade mean stress and cyclic stress summary is shown in Figure 10, and the interference diagram in Figure 11. The modified Goodman diagram is shown in Figure 12; the allowable cyclic stress is 85 percent of the 10° endurance limit, and the allowable mean stress is 85 percent of the 0. 2 percent yield.
The tangential stresses in the rim section vary slightly in the axial direction; the maximum stress is 52, 600 p. s. i.
19
40
35
30
.- 25
w CO U 20 ai H (A
15
10
jl f L
1 V
k>—, f 1 N t 1 \
1 1 1 1 1—MEAN PLUS r
1 N ̂ . x MAXIMUM CYCUC
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| \
| \
K ■„,-^
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i N \
■ s \ i k \
1 ^ i \
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8 \ \ s \ \
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i i i i : MEAN MINUS \ w «i
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; 1 ; i " 1
1 1 p I
: |
I \ b
I \ i 1 \ •
1 |
1 i \ 1 1 .! 1 i
5 6 RADIUS - INCHES
Figure 10. Axial Compressor Blade Stress Summary - Continental Model 357-1 Engine.
20
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2600
2400
2200
2000
1800
1600 u ■
Z 1400
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0 2 4 6 8 10 12 14 16 18 20 22 ENGINE SPEED - r. p. m. IN THOUSANDS
Figure 11. Axial Compreieor Blade Interference Diagram - Continental Model 357-1 Engine.
21
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The material is AMS 4928 titanium with an ultimate of 130, 000 p. s. i.. a yield of 120,000 p. s. i., and an endurance limit of 53,000 p. s. i. with adequate structural margin.
Inducer Compressor
The inducer compressor is an integrally bladed rotor, fully machined from a titanium forging. The blade mean stress and cyclic stress summary is shown in Figure 13, and the blade interference dia- gram is shown in Figure 14. The modified Goodman diagram is shown in Figure 15; the allowable cyclic stress is 85 percent of the 10° endur- ance limit, and the allowable mean stress is 85 percent of the 0. 2 per- cent yield.
T+ie maximum stresses in the rim are 46, 500 p. s. i., tangen- tial; and 28, 000 p. s. i., radial.
The material is AMS 4928 titanium with an ultimate of 130, 000 p. s. i., a yield of 120, 000 p. s. i., and an endurance limit of 53,000 p. s. 1 with adequate structural margin.
RADIAL COMPRESSOR
The radial compressor is an integrally bladed rotor fully ma- chined from a titanium forging. The blade stress summary is shown in Figure 16. The blade interference diagram in Figure 17 is based on experimental data. The disc stress summary is shown in Figures 18 and 19. The modified Goodman diagram for the blade and disc is shown in Figure 20.
The material is 6AL-6V-2Sn with an ultimate of 150, 000 p. s. i., a yield of 140, 000 p. s. 1., and an endurance limit of 62, 500 p. s. i. with adequate structural margin.
Turbine
The turbine is an integrally bladed rotor fully machined from a nickel alloy forging. The blade mean stress and cyclic stress sum- mary is shown in Figure 21, and the blade interference diagram in Figure 22. The modified Goodman diagram for the blade is shown in Figure 23; the allowable cyclic stress is 85 percent of the endurance limit and the allowable mean stress is 85 percent of the 1, 000-hour
23
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Figure 13. Inducer Blade Stress Summary - Continental Model 357-1 Engine.
24
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Figure 14. Inducer Blade Interference Diagram -• Continental Model 357-1 Engine.
25
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Figure 22. Turbine Blade Interference Diagram - Continental Model 357-1 Engine.
33
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•CYCUC/* CYCLIC ALLOWABLE - PERCENT
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stress rupture stress. The disc mean stress and the mechanical property summary for Udimet 700 material is shown in Figaio 24. The modified Goodman diagram for the disc is shown in Figure 25.
The stress and vibration levels for the disc compared to the material properties shew an adequate structural margin for the turbine.
CRITICAL SPEEDS AND SHAFT STRUCTURAL ANALYSIS
The principal source of mechanical vibration in a gas turbine engine is unbalance in the rotor system. This unbalance may excite critical speeds in the rotor itself or induce resonance in any combina- tion of structural elements that may be tuned to the recurring fre- quency of the rotor. If sufficient flexibility exists in a rotor system, locked-in bending moments caused by the uneven distribution of unbal- ance along the length of the rotor may cause ihr, shaft to deflect at high speeds. This deflection will destroy the balance that was pre- viously obtained at lower balancing speeds.
The lightweight, high bending modulus rotor system, combined with the high spring rate rotor suspension system and the short spans between bearing supports, has resulted'in a design that permits rea- sonable bearing loads and shaft deflection under the environmental con- ditions. This combination of rotor and supporting system compliances has also minimized the effect of rotor unbalance and critical speeds.
The shaft flexural critical speeds are above the engine opera- ting range. These values a^e computed using finite spring rate at the bearing »upports and giving full consideration to the gyroscopic stiff- ening effects of high mass moment of inertia components in the rotor system.
The critical speeds and the bearing support spring rates are given below.
Critical Speeds:
Compressor Shaft = 41,000 r. p.m.
Turbine Shaft = 38, 300 r. p.m.
35
a;
(A tn U DC »-«
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RADIUS - INCHES
Figure 24. Turbine Disc Steady-State Stresses - Continental Model 357- 1 Engine.
36
;-smm*. ii"i
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«CYCLIC/ »CYCLIC ALLOWABLE - PERCENT
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o o m o c i 0
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^.
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Spring Rates:
No. 1 Bearing Support = 5.67 x 10 lb/in.
No. 2 Bearing Support = 108 lb/in.
No. 3 Bearing Support * 108 lb/in.
No. 4 Bearing Support = 6. 52 x 105 lb/in.
The torsional critical speed of the turbine is 9, 300 r. p. in., placing it slightly below the engine idle speed.
The maximum cyclic compressor shaft stress is 18, 100 p. s. i. The endurance limit of the AMS 4928 titanium shaft material is 53, 000 p. s. i., with adequate structural margin.
The maximum cyclic turbine shaft stress is 11,600 p. s.i. The endurance limit of the Waspaloy shaft material is 50, 000 p. s. i., with adequate structural margin.
STATIC STRUCTURE
The application of the Model 357-1 engine has required that particular emphasis be placed upon the rigidity of the static structure. Maximum section moduli with minimum weight have been obtained by the use of annular structures without longitudinal joints. Minimum shear loads and bending moments have been effected through the use of lightweight materials and short length multipurpose components.
Components subject to temperature gradients have been de- signed with gradually varying sections to keep thermal stresses with- in the material capabilities. Flanges necessarily represent abrupt changes in section, and all have been isolated from the high tempera- ture areas by suitable conical or cylindrical transition sections. These sections are designed to provide compliance between the rela- tively stiff flange and the thermally cycled hot parts. Maximum ther- mal stresses are experienced during temperature transients. Flange materials have been chosen that will provide yield points above the high transient stresses as well as suitable stress rupture properties at stabilized temperature levels.
38
Care has been taken to make welds only in areas of gradual section changes. Struts and bosses have extended lips tapering to approximately sheet thickness at the weld areas. Annular flanges have conical or cylindrical extensions gradually tapering to sheet thickness at the weld areas. These extensions provide essentially identical heat conduction sections at the weld in order to realize maximum physical properties in the weld regions. The welds, being in areas with gradual section change, are subjected to minimum ther- mal stresses.
The basic structural support of the Model 357-1 engine, Fig- ure 26, is the center support ring (i). The ring directly supports the radial loads of the center bearings and the center support members. The front bearing support and the attached accessories are cantile- vered from the center support ring by the compressor housing. Like- wise, the rear bearing support and the exhaust nozzle are cantilevered by the combustor housing. The primary engine support lugs are an in- tegral part of this center support ring.
The center support lugs, located 60 degrees apart on the in- board side of the engine, support the centrifugal force of the engine, the precession moment of the engine rotor, the vertical (flapping) ac- celeration forces, the shear forces resulting from the chordwise (lead and lag) acceleration of the helicopter rotor tip, and the shear force of the engine thrust.
An engine stabilizing lug is located on the inboard side of the rear bearing support ring (20). This lag is designed to take reactions in line with the helicopter rotor blade, allowing free motion in other directions. The basic loads supported by this lug are the reactions resulting from moments of the engine forces about the center support lug. These include the centrifugal force acting through the engine cen- ter of gravity normal to the helicopter rotor a*is, and the chordw.ise acceleration forces and the engine thrust acting parallel with the en- gine rotor axis.
The front bearing support (2) is a one-piece aluminum casting consisting of the front support ring, five radial struts, and a front bearing housing. The support is provided with cored and drilled pas- sages for the accessory drive shafts, pressure and scavenge oil, fuel supply, overboard drains, anti-icing, and labyrinth seal pressuriza- tion air.
39
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The front bearing housing supports the compressor shaft front radial bearing (3). the engine thrust bearing (4). the internal acces- sory gear drive (5), the fuel transfer labyrinth seal (6), and the inlet nose cone (7).
The front bearing housing and the front support ring aie con- nected by five struts, crossing the airflow path. The boundary sur- faces form a smooth contour for the air entry into the axial compres- sor (8). Three of the struts shroud the accessory drive shafts to the fuel pump, fuel control, and the oil pump. The orientation of these ac- cessories in the centrifugal field has dictated the location of the struts. The remaining struts are spaced for load distribution and structural rigidity of the support.
The front support ring furnishes the mounting and accessory drive pads for the engine accessories. A cylindrical extension serves as the axial compressor shroud, while forming a closure with conical extensions of the compressor housing (9) to create the anti-icing and air start annul!. This extension also traps the axial Stators (10), and prevents their rotation by a slot and tang engagement. The outer sup- port ring is piloted and bolted to the front flange of the compressor housing.
The axial Stators (10) are precision-cast aluminum to provide close dimensional and form control. They are split in two circumfer- ential segments, doweled, and machined as an assembly. The stators are nonstructural elements, thus keeping distortion to a minimum and stress levels low.
The titanium compressor housing (9), is fabricated of an inner and outer shell joined at both ends. The outer shell is integral with the center support ring and forms the primary front structural casing. The inner shell forms the shroud for the radial inducer (11), the outer boundary of the inter-compressor air passage, and the pilot for the axial compressor stators. The cavity formed by the closure of the two shells serves as the engine oil tank (12).
The compressor housing is redundant with respect to the trans- mission of axial and radial loads. However, the structural members supporting the inner shell are axially flexible relative to the conical memWr transmitting the load to the outer shell. The major portion of the internal gas loads is transmitted through the inner shell, result- ing in axial deflections which do not affect critical clearances.
41
A steel weldment consisting of the cast axial diffuser segments and the forged, fully-machined, center support diaphragm (13), is piloted in the center support ring and trapped by the ccmbustcr hous- ing (17). The center bearing housing (14) is bolted to the inside diam- eter of the center support diaphragm. This cast and machined hous- ing supports the radial loads of the compressor shaft rear bearing and the turbine shaft front bearing.
Th* radial compressor shroud (15) and the radial diffuser as- sembly (16) are trapped in the compressor housing by the center sup- port diaphragm. These components are mutually interlocked, and the center support diaphragm assembly is prevented from rotating by a slot and tang engagement with the combustor housing. The differen- tial pressure acting on these components is transmitted as an axial load through the radial diffuser and the radial compressor shroud to the conical segments of the compressor housing.
The combustor housing (17) is a steel casing with a welded flange at each end. The front flange is bolted to the center support ring. The rear flange is bolted to and supports the rear bearing sup- port (20), exhaust nozzle (21), combustor shell (18), and turbine inlet nozzle assembly (19).
The rear bearing support (20) is a welded assembly of forged and machined steel. The support consists of an outer ring, five sup- port struts, and an inner housing. The support houses the turbine shaft rear bearing and supports the exhaust nozzle through a conical isolation section. The turbine shaft rear bearing radial load and the axial gas loads exerted on the exhaust nozzle are transmitted from the rear bearing support through the combustor housing to the center support flange.
The exhaust nozzle (21) is a sheet metal weldment consisting of the outer duct, five vanes, and the inner cone. The inner cone is supported from the outer duct by the vanes. The vanes also shield the support struts, oil supply and scavenge lines, and the labyrinth seal air supply from the turbine exhaust gases.
The exhaust nozzle is required to support only its own weight in addition to a small axial load. However, deformation of the nozzle in the g field will critically affect engine performance. The plug-type nozzle has been designed for low shear loads and bending moments at the support structure with high conical rigidity at the critical flow areas.
42
The turbine inlet nozzle (19) is a welded assembly consisting of an outer ring, 25 hollow vanes, and an inner cone. The assembly forms the boundary passage for the combustion gases, directing the flow into the turbine. The assembly is enclosed by relatively cool compressor discharge air, which passes through the hollow passages of the vanes and through the inner combustor shell assembly to the combustion chamber.
The inner cone supports the inner combustor shell assembly and serves as labyrinth seal shrouds at the rear of the fuel slinger and at the turbine inlet. The inner combustor shell consists of a conical support and a perforated dish which directs air into the pri- mary combustion zone and absorbs thermal radiation. The two parts are welded at their inner diameters and are free to expand in radial and axial directions.
The outer ring of the turbine inlet nozzle has a cylindrical extension which serves as the turbine shroud. The ring is flagged at each end. The front flange pilots and is bolted to the outer com- bustor assembly.
The outer combustor (18) is fabricated in two pieces: the outer shell and the swirl vane. The swirl vane, which directs pri- mary airflow into the combustion chamber, is located axially to the outer combustor shell by spring clips. The vane is piloted at the in- ner diameter and restrained from axial deformation by the center bear- ing housing rear labyrinth seal. This construction allows free radial expansion of the vane, which is subject to large thermal gradients. A conical support is welded to the front of the outer combustor shell. This support is perforated and slotted to provide passages for the pri- mary combustor air and the center bearing housing oil supply and scavenge lines.
The bolted assembly, consisting of the outer combustor shell and the turbine inlet nozzle, is attached to the combustor housing rear flange through a conical transition segment. This assembly is radially supported at the front end by the outer combustor shell' support pilot- ed by a cylindrical extension of the center support diaphragm (13). Therefore, the assembly is radially supported at both ends and is free to expand axially in the forward direction.
The radial support required for the outer combustor shell and turbine inlet nozzle assembly-is for its own distributed load in the g
43
WET"-
field. The pressure differential across the turbine inlet nozzles and the turbine inlet labyrinth seal results in an axial load that is trans» mitted through the combustor housing to the center support ring.
The general structural philosophy has been to use large-diam- eter, continuous structure with suitably compliant sections that per- mit tightly bolted joints. This eliminates the necessity for expensive sliding joints, which are subject to fretting, high frequency vibration, and higher manufacturing costs.
Front Bearing Support Frame
The front bearing support frame is a cast aluminum ring and spoke structure. The stress and spring rate summary is shown in Figure 27. The material yield point is 33, 000 p. s. i. and the ultimate is 43,000 p. s.i.
The maximum stresses are below the yield in both the 235 and 259 g condition. Localized yielding may occur at the 367g over speed
condition but this is considered satisfactory.
Center Support Ring
The center support ring is machined from a titanium forging. The ring is stiffened internally by the center main bearing diaphragm.
The support ring stresses are essentially symmetrical above and below the horizontal center plane of the engine. The highest stresses exist in the 0- to 90— degree quadrant, and these values are tabulated in Figure 28. These values are shown for the 235, 259» and 367g conditions.
The ring material is 6AL-6V~2Sn titanium with an ultimate of 150, 000 p. s. i. and a yield of 140, 000 p. s. i.
The ring stresses are well below the yield for the 235 and 259 g condition. Localized yielding may occur at the 367g overspeed condition but this is considered satisfactory.
Center Bearing Support (Compressor Cover)
The center bearing support is fabricated from cast steel vanes along the OD and a fully machined steel forging in the center.
44
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9 FIELD
ENGINE SUPPORT RING
DEC FIELD BENDING STRESS
k.s.i. HOOP STRESS
k.s.i. COMBINED STRESS
k.*.i.
0 235 29 367
51.2 56.4 80.0
39.5 43.5 60.1
90.7 99.9
140.1
10 235 259 367
33.9 37.4 5X0
37.6 41.4 58.8
71.5 78.8
1118
20 235 259 367
12.8 14.1 2D.0
35.0 38.6 54.6
47.8 5Z7 74.6
30 235 259 367
22.7 25.0 35.4
19.1 21.1 29.9
41.8 46.1 65.3
40 235 359 367
9.6 10.6 15.0
49.3 54.3 77.0
58.9 64.9 92.0
50 235 259 367
14.7 16.3 230
48.9 53.9 76.3
63.6 70.2 99.3
60 235 259 367
27.6 30.4 43.0
46.1 50.8 72.1
73.7 81.2
115.1
65 235 259 367
28.8 31.7 45.0
44.8 49.4 70.0
73.6 81.1
115.0
70 235 259 367
26.9 29.6 42.0
43.9 48.4 686
70.8 78.0
110.6
80 235 259 367
16.6 18.3 26.0
40.1 44.2 62.6
56.7 62.5 88.6
90 235 259 367
5.1 5.7 8.1
38.3 42.2 598
43.4 47.9 67.9
igure 28. Continental Model 357- -1 Engine Support Ring Str
46
The material is 17-4 PH stainless steel with a yield of 170. 000 p. s. i. and an ultimate of 190, 000 p. s. i. The maximum stresses pre- sented in Figure 29 show a substantial structural margin. These stresses are computed from a combination of the air presst load and the g field load.
STRESS INT DIRECTION
STRESS k.$.i ^g 367g
1 TANGENTIAL RADIAL
4U 56.6 38.8 54.9
2 TANGENTIAL RADIAL
60.6 67.0 1.2 11.2
Figure 29. Continental Model 357-1 Compressor Cover Stresses.
Rear Bearing Support
The rear bearing support is a welded spoke and ring structure. The material is Inco 718 nickel alloy with 127, 000 p. s. i. yield and 147, 000 p. s. i. ultimate at 1, 000oF. The stresses presented in Fig- ure 30 show an adequate structural margin.
A summary of the Model 357-1 structural steel housing stresses is presented in Table 3. Table 4 summarizes the combustor and tur- bine inlet nozzle stresses at 259g.
47
9 FIELD
REAR BEARING SUPPORT
•svnuT THERMAL STRESS
lt.s.i. 235g STRESS
k.s.i.
THERMAL STRESS PLUS 2599 STRESS
k.s.i. 367g STRESS
lt.s.i.
1 13.3 20.2 21.0 25.3 2 19.1 20.3 20.6 22.3 3 28.5 20.0 19.3 14.9 4 28.5 25.2 24.7 21.7 5 19.1 25.4 26.0 29.7
Figure 30. Continental Model 357-i Rear Bearing Support Stresses.
TABLE 3 STRESS SUMMARY FOR STRUCTURAL SHEET HOUSINGS
Longitudinal Stress or
, k. s. i. Material Properties, k. s. i.
Item Field Tension Comp. Ultimate Yield Buckling
Compressor 259 23.7 14.1 150 140 23.7 Housing* 367 29.6 20.0
Combustor 259 22.6 17.9 147 98 35.3 Housing 367 28.4 23. 7 at 600oF at 600oF at 600oF
* Hoop stress, IZ, 100 p. s.i.
The compressor housing and the combustor housing are large diameter, high inertia, structural beams. The relatively thin sheet metal construction produces a beam critical in buckling. The stress summary indicates considerable tensile yield strength margin and
48
compressive stresses below the material buckling limit. The combus- tor housing physical properties are given at the maximum expected temperature for the parts.
1 TABLE 4 | ! COMBÜSTOR AND INLET NOZZLE (TIN) | STRESS SUMMARY AT 259g
1 Item Stress
(k. s. i.) Type
Material Yield
(k. s. i.)
Operating Temp.
(0F) 1
Combustor Flange 29.5 Bending 35.0 1200 1
Combustor Shell 12.4 % Beivling ."v C 1200 j
Combustor Shell 11.0 Tangential 35.0 1200
TIN Inner Shell 5.7 Shear 21.0 1000 j
1 TIN Vane 3.3 Shear 21.0 1200 j
The combustor structure is a large-diameter shell supported laterally at its forward and aft sections. In order to withstand the high g loads, the forward support was incorporated..
The stress summary and material properties indicate an ade- quate margin for the combustor and turbine inlet nozzle in the 259g field. The 367g stresses are not presented due to the fact that the combustor and turbine inlet nozzle would be completely contained within the external supporting structure in the event of their failure.
ACCESSORY DRIVE
The accessory drive train, Figure 31, consists'of four spur gears and four bevel gears, which provide the three drives tabulated below:
Drive
Oil Pump Fuel Pump Fuel Control
Drive Horsepower
Total
49
1 The r. p. m. signal will be provided from the fuel control.
Referring to Figure 31, a high speed spur (1) originates from the center of the engine, drives a compound spur-bevel gear (3) through a compound idler (2). The bevel gear (3) drives three radial- ly located bevel gears (5). (6), and (7). The three gears drive through three radial tower shafts (8). (9). and (10) to furnish the three acces- sory drives.
Operation in the tip turbojet environment requires the particu- lar orientation of accessories shown. The fuel control axis must be horizontal, the fuel pump axis vertical, and the oil pump in the lower outboard quadrant.
The three accessory drive pads are located on the outside di- ameter at the front frame of the engine and are designed to suit the en- gine requirements. Accessory drive capabilities are tabulated below:
Drive
Oil Pump Fuel Pump Fuel Control
Torque Rating Gear (lb-in.) Ratio Continuous Static Rotation*
0. 2984 0. 2984 0. 2984
24.2 77.6 2.4
Counterclockwise 310 Counterclockwise 25 Counterclockwise
* Rotation is expressed viewing the mounting pad.
The gear teeth have been analyzed and designed against the three general types of failure: pitting, bending or tooth breakage, and scoring. The first two are essentially fatigue-type in character and are functions of the total number of stress cycles as well as magnitude of dynamic load. Scoring is usually caused by local friction welding and will occur on the first application of maximum speed and torque.
The design criteria used to evaluate the gears against each of the characteristic types of failure are: Hertz stress for fitting, root beam stress for bending, and PVT factor for scoring.
The load used in computing Hertz and beam stress is composed of the total transmitted load plus a dynamic increment load. This dy- namic increment results from inertia forces developed in the gear train by instantaneous accelerations caused by inaccuracies in spacing,
50
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thickness, and tooth profiles. A perfect involute will transmit uni- form angular velocity, but any deviation from this form will cause a change in velocity which, over a very short time interval, represents a high acceleration of the gear mass and all other rigidly attached inertias in the train.
The method used in calculating the dynamic increment is based on Buckingham's works. Although various sources have indicated that the dynamic load as calculated by Buckingham's equation is larger than the actual dynamic load encountered, most authorities agree that it is the best method for comparing different designs. Continental's experience indicates that aircraft-quality, hardened and ground gears can operate up to 230, 000 p. s. i. wear (Hertz) stress with little or no development.
Tooth beading stresses are calculated by applying the dynamic tooth load to the Lewis beam stress equation. This formula determines the equivalent stress that would exist on a constant strength parabola inscribed within the tooth, tangent at the base, and with its apex just intersecting the line of action of the applied tooth load. The stress limit that Continental has successfully permitted on previous designs is 100,000 p. s. i.
The static bending stress is computed as above but using only the static tooth load at the static torque rating.
The PVT scoring factor is a measure of the instantaneous rate of heat generation. It combines the Hertz contact pressure and the sliding velocity at the tip of the tooth with the length of the line of action from the pitch point to the tooth tip.
The safe limit for the scoring factor using very accurate case- hardened gears lubricated with a medium weight petroleum oil is 1, 500. 000 (dimensionless). Continental's experience with MIL-L-7808 turbine oil indicates that no decrease in scoring limit is necessary. The gears in this train have all been designed so that no PVT factor exceeds one-third the permissible safe unit.
The accessory gear stresses are presented in Table 5. Ac- cessory bearing loads and B-10 lives are presented in Table 6,
Figure 32 is a schematic diagram of bearing location for the Model 357-1 engine.
52
TABLE 5 MODEL 357- ■ 1 ACCESSORY GEAR STRESSES
Wear Bending Static Stress Stress Bending
Mesh (P- »- i-) (p. s.i.) Stress
High Speed Spur 226,100 65.500 33.000 Low Speed Spur 223, 000 63,800 40.700 Oil Pump Bevel Fuel Pump Bevel
159,000 203. 000
44.000 71.600
6.800 87.400
Fuel Control Bevel 141,000 34,600 700
TABLE 6 MODEL 357 -1 BEARING LOADS AND B- 10 LIVES
No. Radial Thrust Equiv 1 (Ref. Speed Load Load Load Life > 1
Fig. 32) (r. p.m.) (lb.) (lb.) (lb.) (Hours*)
1 22.000 41.3 0 41.3 Over 104 1 2 10.645 63.2 0 63.2 4850
f 3 10,645 ' 4.3 0 4.3 Over 104
4 5.207 198.0 49.7 198.0 1280 5 5,207 118.0 0 118.0 6207
104 6 6.565 6.9 0 6.9 Over
1 7 6,565 2.7 28.4 60.6 Over 104
8 6,565 0 49.4 76.5 8820 9 6.565 76.0 0 76.0 6500
io4 10 6.565 6.0 32.3 68.6 Over 11 6.565 7.7 32.0 54.9 6360
i 12 6,565 7.7 0 7.7 Over 104
13 6,565 205.0 0 205.0 8290 14 6.565 121.0 29.5 131.2 8470
104 15 6,565 24.7 0 24.7 Over
* All accessory drive bearing lives are based on continuous 100 per- cent speed and load.
53
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SEALS
Labyrinth seal« are used exclusively in the Model 357-1 engine because:
1. They are not subject to the wear problems that con- ventional face and lip seals would experience in the high g field.
2. They are considerably less sensitive to deflection.
3. They are virtually maintenance free.
4. They provide a far simpler engine assembly.
Straight, constant diameter, labyrinth seals have been used throughout the engine for simplicity, economy, easeof assembly, and their immunity to axial tolerances and thermal growth.
The seal element proportionality and overall configuration is based on previous engine experience as well as published literature. The seal configurations used have been individually idealized for sealing requirements, space utilization, radial and axial thermal growth, and to control the resultant rotor thrust load.
Lands of labyrinth seals of the rotating shaft structure are in- tegral with the shaft to preclude the possibility of severe shaft damage in case of severe rub.
All the engine oil seals are designed with a low pressure differ- ential and in such a manner that air flows into the oil compartments. The front and center bearing oil compartment balance pressure, Figure 33, is provided by holes in the compressor shaft that expose the air side of the oil seal to the axial compressor discharge pressure. The air passages are large relative to the expected airflows to assu the same pressure gradient across each oil seal.
Throttled compressor discharge air is provided between the seal lands at the rear bearing oil compartment, Figure 34. The air- flow through the labyrinth seal alcng the back face of the turbine serves to purge the hot exhaust gas from the oil compartment. The balance of the air provides a relatively cool airflow into the rear bearing oil compartment.
55
ENOINiOIL OIL SEAL BALANCE PRESSUtE
ENGINE OIL
Figure 33. Continental 357-1 Engine Oil Seal Balance Pressure.
ENGINE OIL
Figure 34. Continental Model 357-1 Rear bearing Oil Seal.
THROTTLED COMPRESSOt DISCHARGE AREA
56
THROTTLED COMPRESSOR- DISCHARGE AIR
Figure 35. Continental Model 357-1 Fuel Pump and Fuel Control Labyrinth Seal.
Throttled compressor dis- charge air is provided between the seal lands for the fuel pump and the fuel control labyrinth seal, Figure 35. This air supply pro- vides a continuous airflow into the front bearing compartment.
The air flows through the labyrinth seals into the engine oil compartments, flows through the oil scavenge line to the oil tank, and is exhausted through the oil tank vent to the atmosphere. The designed airflow from the air side into the oil compartment produces low oil seal leakage.
ENGINE WEIGHT AND CENTER OF GRAVITY
The breakdown of engine weight and center of gravity for each major section is presented in Table 7,
57
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SECTION FOUR. SYSTEMS DESIGN
FUEL SYSTEM
This section of the report describes only the fuel system con- tained in the engine. The fuel pump and fuel control are discussed in Heavy Lift Tip Turbojet Rotor System, Volume XII, CAE Report No. 943.
The starting fuel is controlled by a solenoid valve. The start- ing fuel. Figure 36, is injected through two starting fuel nozzles,
^^^^^m^J!i^m£m:^^s^^^
STARTING FUEL SOLENOID STARTINO
FUEL NOZZLE
Figure 36. Continental Model 3 57-1 Engine Starting Fuel System.
61
placed alongside two igniters. The flame from th<; starting fuel noz- zle spray passes through louvers in the primary swirl vane into the combustor and ignites the main fuel flow issuing from the fuel dis- tributor. After the main fuel flow is ignited, the starting fuel flow is turned off.
The g field is felt as fuel pressure prior to the fuel leaving the spray nozzle. After leaving the spray nozzle, the fuel is not sub- jected to the g field and should produce a spray pattern similar to a conventional engine.
The metered engine main fuel. Figure 37. enters the engine at the fuel control mounting pad (A) and passes through driUed pas- sageways to the rotating fuel tube (B). From the fuel tube the fuel is sprayed through 18 holes in the fuel slinger (C) into the combustion chamber.
A labyrinth fuel tube seal. Figure 38, is used. This type of fuel tube seal has been employed successively on a previous Continental engine. Fuel enters a low pressure fuel nozzle (i) and sprays into the rotating fuel tube (2). High pressure compressor dis- charge air is provided at the labyrinth seal annulus (3), from where it leaks past two lands of the seal and enters the rotating fuel tube at (4). Any fuel leakage past the labyrinth seal enters the annulus (5) and is drained overboard. The oil from the main thrust bearing is sealed from the fuel system by a three-land seal (6).
The fuel inside the small rotating fuel tube is subjected to a force of 3400g at full engine r.p.m. The fuel at the fuel slinger is subject to 38, 500g. These high g fields should sufficiently over- power the helicopter g field to ensure that fuel tube and fuel slinger operation will be essentially the same as it is in existing Continental engines, and it can be predicted on this basis.
62
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Figure 38. Continental Model 357-1 Engine Fuel Tube Seal.
LUBRICATION SYSTEM
GENERAL DESCRIPTION
Figure 39 is a schematic of the lubrication system for the helicopter rotor tip-mounted gas turbine engine. The system con- sists of an engine driven pressure pump assembly, an integral oil tank, an engine-mounted heat exchanger (optional), a pressure oil distribution manifold, and a scavenge oil manifold. The system is a "hot tank" system, where the hot scavenge oil is returned directly to the oil tank. The oil is gravity fed from the oil tank to the pres- sure pump assembly where it is pressurized, filered, and regula- ted and then passed through an air-oil heat exchanger, and from there distributed to the various lubrication points on the engine. The oil flow at each lubrication poim is controlled by an orifice sized in relation to the available pressure. The estimated heat rejection oil flow rate, and temperature rise at each bearing location is shown in Figure 40: The oil scavenge at each bearing location will be accom- plished by the head difference between the bearing sump oil level and
64
© 1. OIL TANK 4. FILTER BYPASS VALVE 7. ENGINE BEARING SIMPS
2. PRESSURE PUMP 5. PRESSURE REGULATOR VALVE
FILTER HEAT EXCHANGER
Figure 39. Schematic Diagram of Continental Model 357-1 Lubrica- tion System.
the tank oil level, plus a pressure boost from the labyrinth seal air leakage. Figure 41 shows the engine maximum heat rejection to the oil throughout the engine speed range.
SYSTEM COMPONENTS
Oil Tank
The lubrication oil is contained within a tank integral with the titanium compressor housing assembly. By utilizing the compressor housing weldment, the tank is an integral part of the engine structure. An inner wall is added near the rear of the tank to prevent the oil from coming directly in contact with the hot radial compressor housing. The use of the engine structure provides a large volume oil tank with low weight. The oil tank volume of 24 quarts completely surrounds the en- gine compressor. The design oil capacity is 6.0 quarts.
65
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FRONT SECTION BEARINGS AND ACCESSORY DRIVE
BEARING DESCRIPTION
ORIFICE NUMBER
ORIFICE LOCATION
FLOW AT23S9
.35 FRONT REAR SIDE
COMPRESSOR BEARING
FRONT SIDE .33
THRUST BEARING
REAR SIDE .35
ACCESSORY BEVEL DRIVE GEAR BEARINGS SPRAY .34 ACCESSORY DRIVE BEARINGS
SPUR GEAR SPRAY .24
TOTALS I.S3 FUEL PUMP DRIVE PAD BEARING
OUTBOARD SIDE OF BEARING .17
FUEL CONTROL DRIVE PAD BEARING
OUTBOARD SIDE OF BEARING .11
OIL HEAT TEUP.
REJECTION RISE
OPERA PRESSI
40
40
40
40
40 3*5 «4
IS 40
IS 40
ORIFICE ORIFICE OPERATING OPERATKG ORIFICE
PRESSURE *•« (33St) (in. Sm.)
76.9 .045
74.0 .045
7*.| .045
•0.5 .037
Tf.S .037
ffl.3 3- 040
IN SERIES
24.0 2..040
IN SERIES
CENTER SECTION BEARINGS
REAR COMPRESSOR BEARING
• FRONT SIDc REAR SIDE
.31«
«33
FRONT TURBINE BEARING
10
11
FRONT SIDE REAR SIDE
.«33
.31« TOTALS l.lfl
40 •0 .043
40 •0 .0*2
40 •0 .0t2
40 •0 .043 45«
REAR SECTION BEARINGS
REAR TURBINE BEARING TURBINE SHAFT COOLING TOTAL ENGINE TOTAL
12 FRONT SIDE .4 40
13 REAR SIDE .4 40
14 TURBINE SHAFT .4« 40
1.2» 34« «9
4.99 im «4
•0
•0
•0
.040
.053
Figure 40. Estimated Heat Rejection. Oil Flow Rate, and Tempera- ture Rise - Continental Model 357-i Engine.
66
(b o 60 e 1200
i i H i > O50 9 •* 1000 • n
2 1 M
<40 2 4(0 Ü 0 u H ^30 Ü
M 6001 H ^ < M tf 01 uzc H 400 0.
§ «10 X 200
a. u L OIL TO MIL .L-
\—1— ■ 7808
1— 1— t 1— T ^
DESIGN FLOW = 5.0 g. p. m. r SYSTEM MAXIMUM TEMPERATURE = 350°] 1 1 1 1 1 1 1 1 1 1 1
6 0 1 1 1 1 i 1 1 ! 1 L_ TITMPFIIATttBF. f.AIN r« ̂ 'A 1 i ! ! ! 1 ! 1^ ^
• ! 1 1 ^e-1 >-
^*-
• i...< ^r
M • 4 0
U r^ / 1
^ vi ^
i 3 2.0
^ s r
^Z X^
\A ^ 1
^1 <* S*
A y s r ,,
ft ^ —. b—, j L- Ü
20 40 60 80
ENGINE SPEED - PERCENT 100
Figure 41. Continental Model 357-1 Lubrication System Character istics Versus Engine Speed.
Figure 42 shows the oil tank capacity at various oil levels measured from the engine canterline. With an oil volume of six quarts and a system flow rate of 5. 0 g. p. m., the oil dwell time is 18 seconds. The 18-second dwell time, in conjunction with the 235g field and large tank volume, provides excellent deaeration charac- teristics. The oil tank is located around the periphery of the com- pressor. The connection for the filler, vent, sight gage, drain, and tank outlet are provided by welded bosses at the required location.
The tank outlet is located at 45 degrees below the horizontal on the outboard side of the engine near the major diameter of the oil tank. The tank outlet fitting is shaped so as to be near the lowest point of the tank in both g fields.
67
;"*"«P'
Ü 2 W
O ci u, Q U a m < u S J u > M
10.0
9.0 (/) u
Ü 5 «.o
w z 3 sC u H z u ü
7.0
6.0
5.01-
rpn M . in rxj 1 i [i ü ^_a^__ ^m^mm — i MI II UM I 3|M m nwi J ■ m i im Mi HB^^M 1
^^^_ ■•^■H ^_^_ ■MMMBM M^MH MÜMMM ___._ -MMBH ^H^^ ^PMM«
100 200 300 400
TANK CAPACITY - CUBIC INCHES
Figure 42. Continental Model 357-1 Lubrication Oil Tank Capacity.
The sight gage is located on the inboard side of the oil tank for visual oil quantity check. The sight gage is mounted in the tank skin so that the oil is not trapped in the gage. At the 6-inch oil level, the sight gage is capable of indicating jf 0.625 inch of oil, which cor- responds to plus 1. 0 quart or minus 1. 2 quarts.
The oil tank fill port is located 30 degrees from the horizontal on the inboard side. The fill port has a 1-3/8-inch diameter opening, which is ample for hand filling the tank. The fill port cap is sealed with an "O" ring and locked in place by a commercial, hand-actuated locking device.
The tank vent is an AND 10050-16 boss located 30 degrees from the horizontal on the inboard side of the engine. The tank vent s 15 inches from the oil level in the 235g field and 11.2 inches in
the one g field. The high location for the tank vent should provide a very low oil entrainment at the vent. A static air-oil separator and vent line can be provided to suit the installation.
The tank drain plug is located directly at the bottom of the oil tank. The 3/4-inch diameter drain opening will provide adequate flow area to permit the oil to drain out in a short period.
68
Pressure Pump Assembly
The pressure pump assembly contains a gerotor pump element, filter, filter bypass valve, and a pressure regulating valve. The pump- ing element is sized to displace 0. 26 cubic inch per revolution and at 6565 r.p. m. wiU produce 7.4 g.p. m. at 100 percent volumetric effi- ciency. In order to deliver 5. 0 g. p. m. to the engine, a minimum volumetric efficiency of 67. 5 percent is needed. Past experience indi- cates that better than 90 percent volumetric efficiency may be obtained with these pumps at standard sea level conditions. The volumetric displacement is adequate to allow for pump flow loss ^ xe to pressure, temperature, speed, and wear. Since the pump housing has a high co- efficient of thermal expansion, a ring that has the same coefficient of expansion as the pump elements has been installed in the housing to control the clearance variation with operating temperature. The pump inlet "kidney" is located so that a smooth flow results from the oil tank outlet line.
The filter is a standard pleated construction with the filter media backed up with a coarse-weave screen, a support spring, and a perforated tube to provide for rigidity in the longitudinal direction and high level crush strength. The filter rating is 64 microns at 100 percent effective. The effective area of the filter is 31 square inches, resulting in a low pressure drop of 0.6 p. s.i. at 100 F with sufficient area margin for dirt build-up.
The valves are oriented so that the valving action is parallel to the g field (ref. Figure 39). The force on the valves due to the weight of the valve and spring acted upon by the high g load will in- crease the valve operating pressure, which is initially set by the valve spring. This characteristic is used to advantage in the pres- sure regulating valve to provide the increased pump discharge pres- sure required for operation in the high g field. The weight of the valve is controlled such that, in conjunction with the increasing g load, the desired system pressure is maintained. The lg pump output pressure is set at 40 p. s. i. and at the 235g condition the pump pressure increases to 176 p. s. i.
The filter bypass valve is a hollow stainless steel ball for low weight to minimize the bypass pressure increase in the high g field. The lg bypass pressure differential of 10 p. s.i. increases to 16 p. s.i. in the 235g field.
69
Pressure Distribution System
The lubrication oil for the accessory drive and front engine bearings will be carried down an inlet strut and distributed by five orifices. Lubrication oil for the fuel pump drive pad bearing and the fuel control drive pad bearing is carried around the outside of the front support ring by cored holes to the respective drive pads. The center bearing and the rear bearing lubrication oil is carried to the respective locations by external lube lines. Figure 43 shows the pressure drop in the oil distribution lines at 75° and 250oF.
Scavenge System
At each lubrication point, a sump area is provided to collect the oil within the bearing cavity. The engine rotor bearings have suf- ficient scavenge area around the bearing to ensure that oil level build- up next to the labyrinth seal is not possible. The sump drain lines are located on the outboard side below the horizontal to provide for oil scavenge in the right angle g fields. Figure 44 shows scavenge line pressure drop at various flow rates and temperatures.
During start up, and in the absence of the high g field, the oil flow rate at engine idle is reduced to 70 percent normal due to a reduced orifice operating pressure. When the helicopter rotor r.p.m. Is increased, the oil flow increases and the scavenge flows are given a boost by the increasing g field. At 235g an oil level difference of one Inch is equivalent to 7. 9 p. s. 1.
HEAT EXCHANGER
Provisions have been made for an engine mounted air-oil heat exchanger to accommodate the engine heat rejection of 1200 B. t. u. pe; minute. A fuel-to-oil heat exchanger must be ruled out on the basis that the engine heat rejection would boil the fuel in the low pressure fuel system employed on this engine.
The cooler mounting pad is provided on the outboard side of the engine front frame and uses six l/4-inch studs to support the cooler. The inlet and exit ducting would be provided by the airframe manufacturer.
70
LOCATION
1-2 PUMP TO HEAT EXCHANGER
23 HEAT EXCH.
2-3 HEAT EXCH. BYPASS
3-4 HEAT EXCH. TO FRONT BEARING
3-5 HEAT EXCH. TO FUEL PUMP PAD GEARING 5-6 FUEL PUMP LOCATION TO Fl.TING
6-7 FITTING TO FUEL CONTROL PAD BEARING
6-8 FITTING TO CENTER FITTING
8-9 CENTER FITTING TO CENTER BEARINGS 8-10 CENTER FITTING TO REAR BEARING
S
BYPASS ^-r 1 j I 6 7 \ | 1 s—□ 2 3 8 m 1 □ □ □
4 9 10
..uCc.-7c LIME FLOW PRESSURE PRESSURE
.37S
.203
.375
.375
.188
.320
.194
.194
21
1.53 V49
10.5 3.47
4 3.30 28
8.8 11 .33
3.19 2.61
9.3 190 5.3
19.3 I» 7.5
.51
1.33
.72
.25
.042
2.45
4.85
4.9
Figure 43. Estimated Pressure Drop in Oil Distribution Lines 49 Continental Model 357-1 Engine.
71
FRONT
SECTION
CENTER
SECTION
TANK
U l
REAR I
SHOIONJ'
LOCATION EQUIVALENT L.SE SIZE
DIAMETER (INCH)
LENGTH (INCH)
FLOW {g.pm.)
AP AT 75° (p.s.i.)
LP AT 250° (p.s.i.)
1.2 STRUT .550 6 1.81 .048 .006
2-3 TWO TRANSFER TUBES
.375 5.3 1.81 .098 .013
4-5 CENTER SCAVENGE TUBE
.555 10.8 1.90 .087 .011
6-7 REAR SCAVENGE TUBE
.555 7 1.29 .039 .005
7-5 REAR FITTING CENTER FITTING
.555 14 1.29 .078 .010
S3 CENTER FITTING TO TANK
.680 13.5 3.19 .078 .010
3-8 PUMP INLET LINE
.680 10 5 .097 .012
Figure 44. Estimated Pressure Drop in Scavenge Lines - Continental Model 357-1 Engine.
72
The present lube system design dictates that the cooler be mounted outboard of the vertical engine centerline. Coolers may be mounted inboard of this centerline but would require modification to the lube system.
Studies indicate that cooler size and weight can vary consider- ably as a function of the cooling air pressure drop. An aluminum cooler design with 6 to 8 inches of water air pressure drop would re- quire about 40 square inches of area and weigh about 5 pounds includ- ing an automatic oil temperature regulating valve. The required air- flow would be about 2. 0 pounds per second. The ram drag on such a cooler would be approximately 40 pounds without any recovery, but with recovery this drag could be either totally eliminated or greatly reduced.
STARTING SYSTEM
The CAE Model 357-1 engine is designed for air starting. The axial compressor, Figure 45, is surrounded by an annular manifold (1) with a number of nozzles for air impingement on the blades of the axial compressor. The starting air is supplied to the engine through a flanged opening (2).
The axial compressor air starting characteristics are shown in Figure 46. This curve is based on actual starting tests conducted on a similar Continental engine. The recommended minimum light- off speed is 2000 r.p. m. Although successful starts have been ac- complished at lower speeds, the high turbine inlet temperatures ex- perienced would dictate that lower starting r. p. m. 's must be con- sidered only as an emergency measure.
The estimated windmill starting envelope is shown in Figure 47, It is planned to investigate various means of extending this en- velope during Phase II of the development. The target for the start- ing capability is to be able to make windmill starts at forward flight speeds up to 900 feet per second engine forward speed. This figure is greater than the flight speed experienced by the advancing engines during maximum helicopter flight speed at maximum rotor speed.
73
•myuwm
Fig. 45. Continental Model 357-1 Air Starting System.
74
SO
34 L •AXIAL STAGE - (30) 3/16 -in. NOZZLES 1 AAXIAL STAGE - (20) 3/16 -in. NOZZLES | ■ AXIAL STAGE - (30) 3/16 -in. NOZZLES,
1 ■ y
7
Y s 7
h- DEUVERY AIR TEMPERATURE
4 1
\A j f <;
30 ~t,0y IX
/ 1 fi s x1
— yjKjnainm j\ i tu t i e f /, s
^ r
r $ w
/ / 6 /
26 >E JS ER Y i ̂ R 50/ 4 V
| PRESSURE, p.t.i.a., j <? f^ ASUK£U AT TH*;
FMTP ANriP i n Y 1 ^ Yj
22 4( '/I/ f > / /
(^ r /
r
18 / / ^
30 ■ Z/' Ak f
^— / %
14 i J f J
* ',
in
s M
g Z D X
5 26 -
E
I
Q U W & CO
Ü Z OS 0 H O 2
0.25 0.50 0.75 1.00 1.25 AIRFLOW - POUNDS PER SECOND
1.50 1.75
Figure 46. Continental Model 357-1 Axial Compressor Starting Characteristics.
T1»
■^."H"1
.50
.45
.40 OS M P9 2 •» D Z X .30 O < S .25
.20
.15
P?^ Figure 47. Continental Model
357-1 Estimated Windmill starting Envelope.
5 10 15 20 25 30 35
ALTITUDE - FEET IN THOUSANDS
ANTI-iCING
Provisions have been made to provide an anti-iced air intake for the engine. The anti-icing system shown in Figure 48 is very similar to the system presently in use by Continental on a similar intake.
Compressor discharge air will be bled from the combustor housing (1) to an annular manifold (2) around the outside of the air intake. From the outside manifold air is piped down the five struts in the air intake to an annular manifold (3) and to the nose cone (4). The placing and number of bleed holes in the engine intake will be determined during subsequent engine testing, and is dependent upon the airflow required to provide sufficient heat to the parts requiring de-icing.
76
M^'1"«—
u
A
0 a g- A
O a n
i
W 9
B A
E A
I
?r >*•
« A
ELECTRICAL SYSTEM
A very simple electrical system, having a minimum number of components, is used on the Model 357-1 engine. A 24-volt DC electrical supply is required to the ignition coil and the starting fuel solenoid, which are used only during engine starting. In addition to these two items, the only other electrical connections on the engine are to the gas temperature-measuring thermocouples, the power sup- ply to the fuel control, and the tathometer instrumentation.
The complete electrical system. Figure 49, consists of an igniter coil, two igniter plugs, and an electrically operated primer fuel solenoid. The igniter leads, coil, and primer solenoid are shielded in accordance with Military specifications to prevent radio interference.
IGNITER .PLUGS
IGNITION COIL
AIRCRAFT FURNISHED
TACHQMETER
28 VOLT D.C. SUPPLY
—I'l' 1" TURBINE
^ EXHAUST W TEMPERATURE I
IUST I ATUREJ
g LOAD Figure 49. Schematic Diagram of Continental Model 357-1 Elec-
trical System.
Preliminary negotiations with Scintilla Division of Bendix Cor- poration have indicated the feasibility of using a conventional capaci- tance discharge spark splitter turbine ignition system in a 259g field.
78
The positioning of the vibrator mechanism, relative to the direction of the g field, is considered critical and a development and endurance testings program would have to be undertaken to guarantee the relia- bility of the electrical components in the high g field environment.
The engine is designed with five chromel-alumel thermocouples to measure the temperature of the exhaust gases. These thermo- couples are wired together to give a single temperature indication to the pilot.
79
h
SECTION FIVE. AEROTHERMODYNAMIC DESIGN
SUMMARY
The service-proven aerothermodynamic design of the J69-T-29 was retained virtually intact for the Model 357-1 tip turbojet: engine. AThile the mechanical design of the aerodynamic components was affec- ted by the operating environment peculiar to the application, the me- chanical design changes were made in such a manner as to not affect the basic aerodynamic performance of the component except as noted below.
Consideration was given to the fact that the J69-T-29 was a short life engine for drone applications and the Model 357-1 was to be a long-life engine for use on man-carrying aircraft. Compliance with this basic change in requirements dictated lower turbine inlet temper- atures than those experienced by the J69-T-29. Accordingly, as the mechanical design progressed, some minor aerodynamic changes were accomplished simultaneously in order to increase the mass flow and compressor pressure ratio of the Model 357-1. These changes re- duced the predicted turbine inlet temperature to acceptable levels for a long-life engine, and reduced the specific fuel consumption to a value that will permit more economical operation of the helicopter.
During this phase, tests were run on the J69-T-29 engine to establish the effects of inlet distortion on engine performance. The results of this testing will be combined with the results of the Phase II wind tunnel tests of the Hiller-de signed engine nacelle inlet to pre- dict engine performance in the vehicle under flight conditions.
A preliminary engine specification was prepared which re- flects the performance of the fully qualified engine. The prototypes used during the development conducted under Phase II of the program are expected to exhibit performance that agrees generally with this specification except in specific fuel consumption. At this stage of the development, we would expect the specific fuel consumption of the prototype to be between 1. 03 - 1. 06.
The primary differences between the prototype engines and the qualification engines are expected to be in the centrifugal
80
compressor and the combustor. The introduction of a compressor of the type shown in Figure 52 should raise the over a11 compressor effi- ciency by approximately 5 percent and the introduction of a combustor of 1 percent less pressure loss along with the reduction of cycle tem- perature at constant thrust that these component i.nprovements would permit, should enable the qualification engine to meet its guarantees.
DISCUSSION
PURPOSE
The prii ary purpose of the aerothermodynamic design work conducted under Phase I of "he tip turbojet development was to evalu- ate the effects of the operating environment on engine performance, and to recommend the component changes necessary to accommodate required mechanical changes without disturbing the aerodynamic per- formance of the components.
SCOPE
The aerodynamic effect of Coriolis forces, centrifugal forces, and varying inlet pressure recovery distortions were evaluated in making design recommendations. In order to counter the high cen- trifugal forces in the structural design, it was desirable to reduce engine length to the greatest extent possible, and a thorough evalua- tion of the effects of shortened flow paths on engine performance was made.
AERODYNAMIC DESIGN EVALUATION
General
The CAE Model 357-1 engine employs the straight-through aerodynamic flow path characteristic of the J69 engine family. Air enters the engine axially and passes through a single-stage transonic axial compressor, a single-stage centrifugal comprc-ssor with a
81
m?rw
radial and an axial diffuser, an annular combustor, a single-stage turbine, diffusing tailpipe, and plug-type jet nozzle.
The prototype engine cross section is presented in Figure 50. Differences from the established J69-T-29 design are most apparent in the jet nozzle and mechanical structure; the other aerodynamic components are similar to. and based upon, established CAE units.
Compressor Section
The axial compressor stage was scaled from a previously tested component selected to increase overall engine life by improv- ing compressor surge margin and increasing mass flow through the engine. This in turn will permit lower operating temperatures con- sistent with the requirements for lower specific fuel consumption and increased engine life for man-carrying applications. The scaled com- pressor was further modified to increase the number of blades to 26 and blade chord length was reduced as necessary to maintain the same aerodynamic performance while providing a mechanical design better suited to the operating environment.
The centrifugal compressor stage utilized is aerodynamically the same as that used in the J69-T-29 engine. This stage combined with the axial stage will produce a specific fuel consumption of about 1. 04 in the prototype engines to be utilized at the onset of Phase II of the program. Refinements in the centrifugal compressor and the ra- dial diffuser section will occur during tLe development testing, and at the completion of development, the compressor will perform essen- tially in agreement with the data presented in Figure 51. These data were obtained from a compressor whose mechanical design is shown in Figure 52. As this performance is achieved, the fuel consumption of the engine will correspond to that shown in Jie production engine specification; that is, a specific fuel consumption value of 0. 99 at the Military rating.
Turbine Section
The turbine stage is the existing J69-T-29 turbine. The ex- haust diffuser, tailcone, and jet nozzle have been redesigned for minimum length and weight.
82
€8
M
O
o
o o
5' a 9
2 o a A
w
W 9 00
9 (ft
o ►1 o n «
(ft n I— o 9
1493
1493
10 11 12 13 14 15 16 17 1 CORRECTED AIRFLOW - lb/sec.
8 19 20 21 22
Figure 51. Performance Comparison - Baseline and Two-Stage Diffuser.
84
WfCUEM
Figure 52. Two-Stage Radial Diffuser Configuration.
Thermodynamic Cycle
The thermodynamic cycle shown in Figure 53 represents the cycle required of the Model 357-1 in order to produce 1700 pounds of thrust at a specific fuel consumption value of 0. 99. To attain this cy- cle, the engine components will require rematching, with the areas and extent of the rematching to be based on the results of the develop- ment testing which will be accomplished prior to qualification of the engine for operational use.
Acceleration Characteristics
Figure 54 is the estimated compressor map for the prototype Model 357-1 engine. An estimated operating line and the maximum acceleration limit are also shown. Aerodynamically, this prototype engine is identical to the standard J69-T-29, except for incorporation of an improved flow axial compressor. Therefore, the Model 357-1 engine time constant was computed from prototype and J69-T-29 data adjusted for estimated characteristics. Figure 55 presents the time constant to be used for engine acceleration and response predictions.
85
EMTBOPT
OOMBIISTOC
AHM. q>M«g|WtWlOBCT jj»-
cemniGN. coMcnM Emcieia
OVfMU. c«rK$wi MtS»M cÄ;»»««» wownr im
ooMümtt cmascT = <n
I023LC WKJBCt = f«
WAT1C «» L«a www = i»» *•
•wraauct ton »«LW» y »"^ aMncnw ramiMMCf
Figure 53. Thermodynamic Cycle and Entropy Diagram.
"\ CCNTUrUCAL TUKHNK AXIAL COMPIIE9SOK
COMPHESSO«
ACCELERATION LIMIT
16 18 20 22 24 26 28 30
CORRECTED ENGINE AIRFLOW . POUNDS PER SECOND
Figure 54. Estimated Compressor Performance - Continental Model 357-1 Engine.
86
2.4
s 2 2.0
8 Q y U ^ ».* U • 0. >» V) i
O ^ 7 *■ «2
M
.8
.4
~ r " 'V \ \
1 1
1 k \
\ \
V \
V \
Si. — M
Figure 55. Estimated Engine Time Constant Versus Engine Speed - Continental Model 357-1 Engine.
10 12 14 16 18 20 22
ENGINE SPEED • r. p. m. IN THOUSANDS
EFFECT OF INLET DISTORTIONS
The J69-T-29 engine was operated on a sea level static test stand with various measurable inlet distortions. The distortions were accomplished by introducing varying density screens into the engine inlet. Data were gathered to establish the relationship between the engine performance parameters and various degrees of inlet distor- tion.
The data will be utilized during future test programs. At that tine. Hiller Aircraft will run inlet distortion tests on their en- gine nacelle during wind-tunnel evaluation of the complete nacelle design. Comparison of the two sets of data will permit a precise prediction of installed engine performance under flight conditions.
The following presents a summary of the inlet distortion test.
Object
To evaluate effects of inlet distortion on performance and surge on the CAE Model 357-1 engine.
Results
Performance was established for a standard inlet and two dis- tortion patterns at engine speed of 22,000 r. p. m., as shown in Table 8.
87
TABLE 8 EFFECT OF INLET DISTORTION ON ENGINE PERFORMANCE |
Standard Inlet
Screen Pattern "A" (Figure 62)
11.8 Max. -5.8 Avg. Press. Drop
Screen Pattern "B" ! (Figure 63)
19.5 Max. - 8.8 Avg. Press. Drop
Fn 1690 1540 1390
EOT 1220 1244 1270
j SFC 1.055 1.100 1.173
Wf
1770 1700 1630
wa 29.3 27.4 25.6 j
P2/Pl 5.70 5.35 5.05 I
Surge-free operation was demonstrated with the most severe distortion pattern. The engine was accelerated from idle (11, 000 r. p. m.) to maximum speed (22, 000 r. p. m.).
Instrumentation
Several different screens were obtained with varying density, Table 9, to develop the desired distortion. An aluminum ring was de- signed to be mounted between the inlet housing and the air inlet bell for support of the screens and necessary instrumentation.
TABLE 9 ] 1 TEST SCREENS
6x6 Mesh/Inch, . 035 Wire 50 x 45 Mesh/Inch, .010 Wire 50 x 50 Mesh/Inch, . 009 Wire \ 20 x 20 Mesh/Inch, .015 Wire
8700 Dutch Weave Filter Cloth, 80x700 Mesh, . 004 x . 003 Wire |
88
The inlet screen configuration is shown in Figure 56. The instrumentation, as observed from the front of the inlet bell, is illus- trated in Figure 57. The distortion equipment, as mounted on the engine, is shown in Figures 58 through 61.
Figure 56. Inlet Screen Configuration
•DENSE MEDIUM SECTION
DENSE
Discussion
This test was proposed to evaluate effects of inlet distortion on performance and surge margin of the Model 357-1 engine. The aerodynamic similarity between the J69-T-29 engine and the Model 357-1 was sufficient to permit use of the J69-T-29 to establish these effects: thus, J69-T-29 engine serial No. 069 was utilized for the test.
Performance was established for a standard inlet condition and two distortion patterns, as follows:
1. Screen Pattern "A," Figure 62, produced a maximum pressure drop of 11.8 percent and an inlet overall pressure drop of about 5. 8 percent.
2. Screen Pattern "B," Figure 63, produced a maximum pressure drop of 19. 5 percent and an inlet overall pressure drop of 8. 8 percent.
89
■Sf
No. 33 - P.
No.40 No. 34
No. 39
o. 36
No. 38
Figure 57. Instrumentation of Air Inlet ior Distortion Test.
90
MUT KLL P/N STUM 7«42MN0USW8
-B-
\
1 •- 1 1 1 1 1
J 1 1
EMMI
i 1 i
1 i 1 i w-. . . ■ _____—x. _ ___ _ * •
MSTNummmoN SCREEN LOCATMM
Figure 58. Schematic Diagram of Inlet Distortion Test Instrumentation.
Figure 59. Installation of Static Pressure Probes and Total Pressure Rakes on Adapter Ring - Model 357-1 Inlet Distortion Test.
91
Figure 60. Installation of Adapter Ring - Model 357-1 Inlet Distor- tion Test.
92
V
Sri-/J.:A^^-»M^: *
■^
Figure 61. View of Distortion Screen in Model 357-1 Engine Inlel.
93
PRESSURE GRADIENT
x 6 MESH/INCH. .03% WIRE
x 45 MESH/INCH, .010 WIRE
DISTORTION
19.5% MAXIMUM • 8.77% AVERAGE
II A II Figure 62. Model 357-1 Engine Inlet Distortion - Screen Pattern "A
94
6 x 6 MESH/INCH. .035 WIRE
PRESSURE GRADIENTS x 20 MESH/INCH,
.015 WIRE
DISTORTION
11. S^ MAXIMUM - 5.8% AVERAGE
Fig. 63. Model 337-1 Engine Inlet Distortion - Screen Pattern "B.1
95
The change of performance from standard inlet and Pattern "A" distortion at 22, 000 r. p, m. is as follows: 8. 5-percent decrease in thrust; 4. 5-percent increase in specific fuel consumption; 6-percent drop in compressor pressure ratio; 4-percent decrease in fuel flow; 6. 5-percent decrease in airflow; and a 250F increase in exhaust gas temperature.
The change in performance with Pattern "B" distortion at 22,000 r.p.m. produced a 17-percent decrease in thrust. 11.7-percent increase wn specific fuel consumption, 11. 5-percent decrease in pressure ratio, 7. 9-percent decrease in fuel flow; 13-percent decrease in airflow, and a 50oF increase in exhaust gas temperature.
The loss in total pressure across the screened section was measured by total pressure probes behind the screen. This loss was then divided by barometric pressure to obtain the percent pressure drop for every probe location. Pressure drop gradient lines, in per- cent, were plotted for Figures 56 and 57. The percent drop for the open, partially screened, and densely screened section is the average of the probe readings for each respective area. The overall average pressure drop is the sum of all the readings divided by the total num- ber of readings.
The change in the various parameters with change in inlet dis- tortion are presented in Figures 64, 65, and 66. The thrust, airflow, fuel flow, and pressure ratio decreases as distortion increases while the exhaust gas temperature increases as the distortion increases.
After obtaining steady-state points with the distortion screens, the engine was accelerated from idle (11, 000 r. p. m.) to maximum speed (22, 000 r.p.m.) with no observable surge. A second accelera- tion with the higher distortion screens, Pattern "B,M resulted in no surge during acceleration from idle to maximum r. p. m.
96
17 iO 21 30
O ul 06 a z D
7.
3 H < a: u a. S w H
< O H (A D
ss X Ul
16
IM a.
15 O
Is w
I a: 14 w
V) u a: Q. BS o
a
12
II
20
19
2 J Ul O u.
IB
17
16
15
29
28
o u
27
S Ul B5 Q Z D
Z
o 0.
26 g
25
24
17i
1 16
-s ü r-
i
^s , \, n [V THRUST
v] s. ^s V 1 i i
\ i "s si i i
. 1 > ̂ S . . N . i i
S t AIRFLOW "S^ 1 i
! V s
I 1 ^ SJ 15 j • >
Nv I K i 1
... x ! 1 \ i \ \
14 \ \
- N w
•*** FUEL FLOW i ^ 1 1 : "^ r4-L \
1 ! 1
13 Th ■^ \ 1 ^ hs \
L i
"1 —. ___ -
PVPI
v s \ f1
^m' i- ^ N ■v
12 **• ̂ ̂ i«*
1 1
1 1
SGT - h-h- -
"* ^- "
!
mmm
- _ M —
11 I 1 0123456789
AVERAGE INLET PRESSURE DROP - PERCENT
Figure 64. Parameter Chsnge Versus Inlet Distortion at Design Speed - Model 357-1 Engine.
97
s
18
16 U
X u 5 A 10
H 8 D
X b H H « 4
JNA in O86.S60
. 2 INLET PRESSURE DROP (%) STD. ENG. . NO DISTORTION Is
D 86. 560 19. 5% MAX - 8.8% AVERAGE A 86. 560 11.8% MAX - 5.8% AVERAGE y
S A fP W
^ ^
'#?
^
*" T L-W 1
rAT> 1 1
. REI 1 1
1 )UCED TO 1 1
STAN 1 1
DAR] >NAi i i
>. CONDITIONS 1,1. 1 i
10 U 12 13 U 15 16 17 18 19 20 21 22
ENGINE SPEED - r. p. m. IN THOUSANDS 23 24
Figure 65. Effect of Inlet Distortion on Thrust - Model 357-1 Engine.
20
n w 18
g z 16 S z 14
U
12
1
10
3 (x 8 J U
JNA in O 86.560
. 2 INLET PRESSURE DROP (%) STD. ENG. - NO DISTORTION i D 86. 560 19. 5% MAX - 8.8% AVERAC
A 86. 560 11.8% MAX - 5.8% AVERAC i i i i • • ■
;E IE J f
4 r &
L n ATA HASE DON FUE L 0 /
1 LOWER HEATING VALUE 1 OF 18. 400 Btu/lb. rf*^
x .cr"
&
P hi rDAi
«K |>taf
A RI :DUC EDT O ST> kNDAl ÄDNi \CA 30ND tT^o^ S
10 11 12 13 14 15 16 17 18 19 20 21
ENGINE SPEED - r. p. m. IN THOUSANDS
22 23 24
Figure 66. Effect of Inlet Distortion en Fuel Flow - Model 357-1 Engine.
98
Unclassified Security Cl«««ificatio«
DOCUMENT CONTROL DATA - R&D (Saeuritr cla—iHeatimi of ml», body ol mbmmel an* totfasfcif < vtmiian muml bm «ntaferf «A«n tfw orarvff report 19 clmmmiliod)
1. ONIGINATING ACTIVITY rCoiporai« aulheri
Hiller Aircraft Company, Inc. Continental Aviation and Engineering Corporation
12« memomr SCCUIITV c LASSIFICATION
Unclassified 2b CMOuP
>. flCr<MT TITLE
Heavy-Lift Tip Turbojet Rotor System. '»Engine Design". Volume XI
4- oeSCNiPTIVE NOTES (Typ» ol nport mnä Inelumirm 4»tmmi
%■ AUTHORfSj (Lmmt ttmmm. Hmt MOM. bitüml)
«. REPORT DATE
October 1965 7«. TOTAL MO. OP »*«C«
99 7k. NO. or RKPS
-1_ ■ •. CONTRACT OR «RANT NO.
DA 44-177-AMC-25(T) b. RROJCCT NO-
Task 1M121401D14412
ta. ONieiNATOR'» REPORT NUMBCRTS)
USAAVLABS Technical Report 64-68K
■ b. OTHCR RfRORT HO(S) (Any Ala import)
CAE Report No. 942 IP- AVA1LABILITY/LIMITATIOM NOTICES
Qualified requesters may obtain copies of this report from DDC. This report has been furnished to the Department of Commerce for sale to the public. 11. SUPPLEMENTARY NOTES 12- SPONSORING MILITARY ACTIVITY
US Army Aviation Materiel Laboratories Fort Eustis, Virginia
IS. ABSTRACT
Volume XI of Heavy-Lift Tip Turbojet Rotor System discusses the design of the Model 357-1 engine for rotor tip operational environment. The detail design meets requirements of MIL-E-5007B except for items peculiar to rotor tip operation. These items include operational attitides, effect of continuous gyro- scopic couples and high g field environment.
This volume includes an analytical design analysis of the static structure, rotating elements, bearings, accessory drives, lubrication system, and miscel- laneous engine parts.
DD /Ä 1473 Unclassified Security Clasiificaticn
Unclassified Security Classificatior
KEf WQROS
Tip Turbojet Engine Design
LINK A LINK a LINK C
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Unclassified " Security Classification
6548-S5
