exper imental quiet engine program: aerodynamic performance of fan a

108 Pages·2009·3.33 MB·English

https://ntrs.nasa.gov/search.jsp?R=19720019045 2018-11-18T23:15:01+00:00Z
,t
; '
I ~..fI1§,; , I . I", ··I ;I. , ''
,{11'.', ; ','' / ' ' " ' i."}'~'~,· 'i.. ,;'
I "''I?;,~-. ;, , i
.X .. I
I.I I I,I .i. .. NA. SA CR-120858
, ' , ' ,.,' ' .1L :)
J
. I.
. ..
, ,1 , , ,! :;' ;
,;{,,
., -... --"
. .
_ _ . .,,
I
I ; ; I "`' ··
r- I | ,, .
EXPER IMENTAL QUIET ENGINE PROGRAM:
AERODYNAMIC PERFORMANCE OF FAN A
If
(NASA-CR-120858) EXPERIMENTAL QUIET ENGINE N72-26695
PROGRA. AERODYNAMIC PERFORMAN.CE OF FAN A
-R.:G. ,,Giffin, et :al (General ;Electric :Co. )...
.May 1971 ,112 p. CSCL 21B Unclas
G3/28 33583 it
J
I I .
I, , I ' r ,
by . I ,..,
I
R.G. Gif,fin, D.E. Parker, and L.W. Dunbar '.
,l(
'.
I ;
GENERAL ELECTRIC COMPANY
. I l
')', ;9.', ! .:
t
;"' ' ' :.I' . .1. C
;, ... .
. .,, .
I·:
1,, 1 prepared for *,\
NATIONAL AERONAUTICS' AND, SPACE ADMINISTRATION
I
.; ,,,, l
Reproduced by
NATIONAL TECHNICAL
INFORMATION SERVICE
U S Department of Commerce
Springfield VA 22151
NASA-Le'wis Research Center ,
. ; Contract NAS 3-12430 : ,
Arthur A.,Medeiros; Project Manager
, .' I': ii
* I ,, ,
//, . . ,; ' ' 6',IQ-.MV
I
1. Report No. 2. Government Accession No. 3. Recipient's Catalog No.
NASA CR-120858
4. Title and Subtitle 5. Report Date
May 1971
Experimental Quiet Engine Program Aerodynamic
Performance of Fan A 6. Performing Organization Code
7. Author(s) 8. Performing Organization Report No.
Giffin, R.G., Parker, D.E. and Dunbar, L.W.
10. Work Unit No.
9. Performing Organization Name and Address
General Electric Company
11. Contract or Grant No.
Aircraft Gas Turbine Division
Cincinnati, Ohio 45215 NAS3-12430
13. Type of Report and Period Covered
12. Sponsoring Agency Name and Address Contractor Report
National Aeronautics and Space Administration
WWaasshhiinnggttoonn,, DD..CC.. 2200554466 14. Sponsoring Agency Code
15. Supplementary Notes
Project Manager, Arthur A. Medeiros,
V/STOL & Noise Division
NASA Lewis Research Center
Cleveland, Ohio
16. Abstract
This report presents the aerodynamic component test results of Fan A, one of two
high-bypass-ratio, 1160 feet per second (353.6 m/sec) single-stage fans, which was
designed and tested as part of the NASA Experimental Quiet Engine Program. This fan
was designed to deliver a bypass pressure ratio of 1.50 with an adiabatic efficiency
of 86.5% at a total fan flow of 950 lb/sec (430.9 kg/sec). It was tested with and
without inlet flow distortion. A bypass total-pressure ratio of 1.52 and an adiabatic
efficiency of 88.3% at a total fan flow of 962 lb/sec (436.4 kg/sec) were actually
achieved. An operating margin of 12.4% was demonstrated at design speed.
17. Key.Words (Suggested by Author(s)) 18. Distribution Statement
Fan, Compressor, Inlet Flow Distortion, Unclassified - Unlimited
Aerodynamic Performance
19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages 22. Price
Unclassified Unclassified 112
* For sale by the National Technical Information Service, Springfield, Virginia 22151
PRIECEDIJNG PAGE BLANK NOT FILMED
TABLE OF CONTENTS
Page
1.0 SUMMARY 1
2.0 INTRODUCTION 2
3.0 DESCRIPTION OF TEST FAN 3
4.0 TEST SET-UP AND PROCEDURE 7
4.1 Test Facility 7
4.2 Overall Performance Instrumentation 7
4.3 Bypass Ratio Schedule 9
5.0 RESULTS AND DISCUSSION 11
5.1 Presentation of Overall Results 11
5.2 Undistorted Flow Performance 11
5.2.1 Fan Bypass Region 11
5.2.2 Fan Core Performance 14
5.2.3 Bypass Ratio Excursion 20
5.2.4 Bypass OGV Schedule 41
5.2.5 Traverse Data 44
5.3 Distorted Flow Performance 44
6.0 SUMMARY OF RESULTS 63
APPENDIX I - INSTRUMENTATION 65
APPENDIX II - COMPUTATION OF OVERALL FAN PERFORMANCE 71
APPENDIX III - AEROMECHANICAL CHARACTERISTICS OF THE NASA
QUIET ENGINE FAN "A" FSFT VEHICLE 77
LIST OF SYMBOLS 107
iii
LIST OF ILLUSTRATIONS
Figure Page
1 Fan A Component Test Vehicle Layout. 5
2 Full-Scale Compressor and Fan Test Facility. 8
3 Clean Inlet Flow Test Bypass Ratio Schedule. 10
4 Fan A Performance Characteristics in the Bypass Region. 12
5 Fan A Performance Characteristics in the Core Region. 13
6 Radial Efficiency Profile at Discharge. 16
7 Radial Profiles. 17
8 Radial Profiles. 18
9 Radial Profiles. 19
10 Variation of Element Properties with Throttling, Design
Stream Function = 0.0606 (Bypass OD). 21
11 Variation of Element Properties with Throttling, Design
Stream Function = 0.303. 22
12 Variation of Element Properties with Throttling, Design
Stream Function = 0.5454. 23
13 Variation of Element Properties with Throttling, Design
Stream Function = 0.7878 (Bypass ID). 24
14 Variation of Element Properties with Throttling, Design
Stream Function = 0.8637 (Core OD). 25
15 Variation of Element Properties with Throttling, Design
Stream Function = 0.8947. 26
16 Variation of Element Properties with Throttling, Design
Stream Function = 0.92. 27
17 Variation of Element Properties with Throttling, Design
Stream Function = 0.9546. 28
18 Variation of Element Properties with Throttling, Design
Stream Function = 0.985 (Core ID). 29
19 Bypass Ratio Excursion, Bypass Performance. 30
20 Bypass Ratio Excursion, Core Performance. 31
21 Variation of Element Properties with Bypass Ratio, Design
Stream Function = 0.0606 (Bypass OD). 32
22 Variation of Element Properties with Bypass Ratio, Design
Stream Function = 0.303. 33
23 Variation of Element Properties with Bypass Ratio, Design
Stream Function = 0.5454. 34
24 Variation of Element Properties with Bypass Ratio, Design
Stream Function = 0.7878 (Bypass ID). 35
iv
LIST OF ILLUSTRATIONS (Continued)
Figure Page
25 Variation of Element Properties with Bypass Ratio, Design
Stream Function = 0.8637 (Core OD). 36
26 Variation of Element Properties with Bypass Ratio, Design
Stream Function = 0.894. 37
27 Variation of Element Properties with Bypass Ratio, Design
Stream Function = 0.9243. 38
28 Variation of Element Properties with Bypass Ratio, Design
Stream Function = 0.9546. 39
29 Variation of Element Properties with Bypass Ratio, Design
Stream Function = 0.9849 (Core ID). 40
30 Effect of Bypass Ratio on Rotor Hub Total Pressure Profile. 42
31 Effect of Bypass Ratio on Splitter Wall Static Pressure. 43
32 Traverse Radial Profiles, Radius Vs. Total Pressure Ratio
(Bypass). 45
33 Traverse Radial Profiles, Radius Vs. Corrected Total.
Temperature Rise (Bypass). 46
34 Traverse Radial Profiles, Radius Vs. Adiabatic Efficiency
(Bypass). 47
35 Traverse Radial Profiles, Radius Vs. Absolute Air Angle
(Bypass). 48
36 Traverse Radial Profiles, Radius Vs. Relative Air Angle
(Bypass). 49
(I
37 Traverse Radial Profiles, Radius Vs. Total Pressure Ratio
(Core). 50
38 Traverse Radial Profiles, Radius Vs. Corrected Total
Temperature Rise (Core). 51
39 Traverse Radial Profiles, Radius Vs. Adiabatic Efficiency
(Core). 52
40 Traverse Radial Profiles, Radius Vs. Absolute Air Angle
(Core). 53
41 Inlet Flow Distortion Level. 55
42 Tip Radial Distortion Inlet Pressure Profile. 56
43 Crosswind Distortion Pattern. 57
44 One/Rev Distortion Pattern. 58
45 Inlet Pressure Recovery Contour, Crosswind Distortion. 59
46 Fan A Performance Map, Tip Radial Distortion. 60
47 Fan A Performance Map, Crosswind Distortion. 61
v
LIST OF ILLUSTRATIONS (Concluded)
Figure Page
48 Fan A Performance Map, One/Rev Distortion. 62
49 Fan A Test Vehicle Instrumentation Location. 67
50 Inlet Boundary Layer Description. 72
51 Typical Inlet Duct Pressure Profiles. 73
52 Campbell Diagram, Rotor Blades. 83
53 Rotor Blade Strain Gage Locations. 85
54 Goodman Diagram for Rotor Blade Scope Limit Calculations. 86
55 Campbell Diagram and Scope Limits, Bypass OGV. 88
56 Campbell Diagram and Scope Limits, Core OGV. 89
57 OGV Strain Gage Locations. 90
58 Goodman Diagram for OGV Scope Limit Calculations. 91
59 Rotor Blade Vibration, Clean Inlet. 93
60 Rotor Blade Vibration, Effect of Inlet Distortion. 94
61 Rotor Blade Vibration, Effect of Bypass Ratio. 95
62 Fan Bypass OGV Vibration, Clean Inlet. 96
63 Fan Bypass OGV Vibration, Effect of Inlet Distortion. 97
64 Fan Bypass OGV Vibration, Effect of Bypass OGV. 98
65 Fan Core OGV Vibration, Clean Inlet. 99
66 Inlet Distortion Magnitudes. 100
vi
1.0 SUMMARY
This report presents the aerodynamic component test results of Fan A,
one of two high-bypass-ratio, 1160 feet per second (353.6 m/sec) single-
stage fans, which was designed and tested as part of the NASA Experimental
Quiet Engine Program. The fan was tested with an undistorted inlet flow
and with circumferential, tip radial, and crosswind distortions. The
hub tip radius ratio was 0.465 at rotor inlet. A flow splitter, immediately
downstream of the rotor, separates the fan bypass flow from the fan core
flow. An axial distance equal to two rotor chords was employed between the
rotor and bypass portion outlet guide vanes (OGV). The vane-number-to-
blade-number ratio was 2.3. No inlet guide vanes were employed.
The fan was designed to deliver a bypass total-pressure ratio of 1.50
at a total fan flow of 950 lb/sec (430.9 kg/sec). The design bypass adiabatic
efficiency was 86.5%. A bypass portion total-pressure ratio of 1.52 and
an adiabatic efficiency of 88.3% at a total flow of 962 lb/sec (436.4 kg/sec)
were actually achieved. The fan core region was designed to develop a
total-pressure ratio of 1.32 at a flow of 144.0 lb/sec (65.3 kg/sec). A
fan core portion pressure ratio of 1.356 was actually achieved at its design
flow. At this condition, a fan core adiabatic efficiency of 83.1% was
measured.
The operational limit line was determined up to 100% corrected speed.
Rotating stall was the operational limit at all speeds except 100%
where high bypass OGV stress precluded further increases in back pressure.
At 100% corrected speed, an operating margin of 12.4% was achieved relative
to the design operating line at altitude-cruise conditions and at 90%
corrected speed the operating margin was 10.8% relative to the design
operating line at sea-level-static conditions.
The fan was tested with one-per-rev circumferential, tip radial, and
crosswind distortion screens installed. At 90% corrected speed, the screens
were found to produce approximately 15% distortion (maximum minus minimum
pressure divided by maximum pressure). The fan demonstrated minimal
operating margin loss to one-per-rev circumferential and crosswind distortions.
The tip radial distortion resulted in a substantial operating margin degrada-
tion.
2.0 INTRODUCTION
In the last 20 years the rapid growth of the commercial aviation
industry has demanded large increases in aircraft size and flight
frequency. This, coupled with the increased public awareness of the
noise pollution problem, has prompted the initiation of an Experimental*
Quiet Engine Program by the NASA-Lewis Research Center (Contract NAS3-12430).
The major objectives of this program are: the demonstration of noise
levels produced by turbofan engines which are designed for low noise out-
put and confirmation that predicted noise reduction can be achieved;
demonstration of the technology and design innovations which will reduce
the production and radiation of noise in high-bypass turbofan engines;
and acquisition of experimental acoustic and aerodynamic data for
turbofan engines, which are designed for low noise output, to give a
basis for correction of acoustic theory and experiment and to give a
better understanding of the noise production mechanisms in fans, com-
pressors, turbines, and nozzles.
Observations of past trends indicate that as tip speed increases,
at constant aerodynamic loading, fan broadband noise increases. Also,
at constant tip speed, a reduction in aerodynamic loading is observed
to decrease fan pure tone noise. Hence, for given mission requirements,
the minimum fan noise configuration requires consideration of the weighted
sum of the two types of noise. In order to confirm acoustic noise pre-
dictions and to acquire acoustic and aerodynamic data for typical high-
bypass fans two low-speed fans, designated Fan A and Fan B, and one low-
aerodynamic-loading fan, designated Fan C, were designed.
The aerodynamic performance of the three fans has been determined
in the General Electric Large Fan Test Facility in Lynn, Massachusetts.
Upon completion of the aerodynamic testing, each fan, in turn, is to be
tested in the NASA-Lewis Acoustic Test Facility.
Fan A, one of the low-speed fans, and the low-aerodynamic-loading
Fan C are to be mated with a TF39/CF6 engine core. The complete full-scale
high-bypass fan engines will then be tested at the General Electric Remote
Test Site at Peebles, Ohio, to determine their overall aero/acoustic
performance as well as thrust and specific fuel consumption.
This report presents the aerodynamic component test results on
Fan A with uniform inlet flow and with tip radial, crosswind, and
circumferential distortion of the inlet flow. Details of the design
of Fan A, and the other fans to be evaluated in this test series, are
given in Reference 1. The aerodynamic component test results of Fan B
are given in Reference 2.
2
3.0 DESCRIPTION OF TEST FAN
The aerodynamic design point for this low-speed fan was selected at
the altitude cruise condition, a flight Mach number of 0.82 at an altitude
of 35,000 feet (10.67 km). At this condition, the corrected tip speed is
1160 ft/sec (353.6 m/sec) with an average fan bypass total pressure ratio
of 1.50 and an average fan core total pressure ratio of 1.32. The design
corrected fan flow of 950 lb/sec (430.9 kg/sec), with a hub tip radius ratio
of 0.465 and a specific flow of 41.3 lb/sec/ft2 (197.0 kg/sec/m2) of annulus
area, results in a tip diameter of 73.354 inches (1.863 m). The design flow
for the fan core, corrected to fan inlet, was 144.0 lb/sec (.65.3 kg/sec)
resulting in a bypass ratio of 5.60.
Figure 1 shows a meridional view of the Fan A component test
vehicle. Immediately downstream of the rotor, a flow splitter is located
which separates the fan bypass flow from the fan core flow. The axial
distance between the rotor and bypass outlet guide vane (OGV), expressed
in number of rotor chords, is a minimum of 2.0. The axial distance between
the rotor and fan core OGV, expressed in number of rotor chords, is 1.25.
No inlet guide vanes were incorporated in the configuration. The moderate-
aspect-ratio (2.32) tip-shrouded rotor contained 40 blades. The rotor tip
and hub solidities were 1.45 and 2.50, respectively. Ninety vanes were
incorporated for both the outer and inner OGV's for a vane-number-to-blade-
number ratio of 2.3.
The design rotor tip relative Mach number is 1.20. The rotor blade
employs a profile shape that is specifically tailored to prevent excessive
shock losses on the suction surface and still be compatible from a throat
area and energy addition standpoint. The blade meanline shape and point
of maximum thickness vary radially. The blade is similar to a double-
circular arc profile in the hub region. Profile shapes at other radii
are generally similar in appearance to the NASA multiple-circular arc
profiles.
The profile shapes for the bypass and core OGV's, which operate at
moderate conditions of inlet Mach number and diffusion factor, were designed
with a modified NASA 65 series thickness distribution on a circular arc
meanline.
Tabulated below are some of the pertinent Mach numbers and diffusion
factors for the rotors and outlet guide vanes:
Outlet Guide Vanes
Rotor Bypass Core
Inlet Mach OD 1.20 0.71 0.67
Number ID 0.75 0.65 0.71
Diffusion OD 0.350 0.442 0.423
Factor ID 0.368 0.412 0.383
Complete design details are presented in Reference 1.
3
If
La
.4
0
....
c-4)
c.)
C)
0
Er
'o
L a)
bb
-r
. t4
rT-
E--
C:
¢
4q
7
a,
'' .
,'- I
LA,
u::
LL-