NASA CR -134609 SIDE INLET LOW-SPEED PERFORMANCE …

NASA CR -134609

BCAC D6-41521

727 AIRPLANE SIDE INLET LOW-SPEED PERFORMANCE

.CONFIRMATION MODEL TEST FOR REFANNED

JT8D ENGINES

(NASA-CR-134609) NHE 727 AIPLAN2 SIDE N74-2066

INBE LCW-SPEED PEdFCREACE CC FIMATION

ACDEL TEbZ FOE REFANNED JT8D ENGINES

(Boeinq Commercial Airplane Co., Seattle) Unclas

p HC $7.25 CSCL 01C G3/02 35579

by A. L. Schuehle

BOEING COMMERCIAL AIRPLANE COMPANY

A DIVISION OF

THE BOEING COMPANY

Prepared for

NATIONAL AERONAUTICS AND SPACE ADMINISTRA

NASA Lewis Research Center

Contract NAS3-17842

Reproduced by TEc NICAL

INFO oN SA c V\CEUS prg* e5D , IA. 22151

NOTICE

THIS DOCUMENT HAS BEEN REPRODUCED FROM THE

BEST COPY FURNISHED US BY THE SPONSORING

AGENCY. ALTHOUGH IT IS RECOGNIZED THAT CER-

TAIN PORTIONS ARE ILLEGIBLE, IT IS BEING RE-

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AS MUCH INFORMATION AS POSSIBLE.

1. Report No. 2. Government Accession No. 3. Recipient's Catalog No.CR-134609

4. Title and Subtitle 5. Report Date

727 Airplane Side Inlet Low-Speed Performance March 1974

Confirmation Model Test for Refanned JT8D Engines 6. Performing Organization Code

7. Author(s) 8. Performing Organization Report No.

A. L. Schuehle D6-41521

10. Work Unit No.

9. Performing Organization Name and Address

Boeing Commercial Airplane Company 11. Contract or Grant No.P.O. Box 3707Seattle, Washington 98124 NAS3-17842

13. Type of Report and Period Covered

12. Sponsoring Agency Name and Address Contractor Report

National Aeronautics and Space Administration 14. Sponsoring Agency CodeWashington, D.C. 20546

15. Supplementary Notes

Project Manager, A. A. MedeirosNASA Lewis Research Center, Cleveland, Ohio 44135

16. Abstract

This report presents the results of a low-speed wind tunnel test of a 0.3 scale model 727 airplaneside inlet for Pratt & Whitney Aircraft JT8D-100 engines. The test was conducted by the PropulsionTechnology Staff of the Boeing Commercial Airplane Company under authorization of NASA ContractNAS3-17842, "Phase II Program on Ground Test of Refanned JTSD Engines and Nacelles for the 727Airplane". The objectives of the test were to develop lines for a full-scale flightworthy inlet,to evaluate inlet total pressure recovery and steady-state total pressure distortion, and toobtain model-scale distortion data which can be used in the assessment of the compatibility ofthe inlet with the JT8D-100 series engines. A secondary objective was to obtain internal/external cowl static pressures for the determination of nacelle loads.

Two basic inlet models were tested at static, forward speed, angle-of-attack (inflow angle), andcross-wind conditions. One model was with and one without an acoustic ring. Two modifications tothe models were also tested, one with the ring closer to the inlet throat and one with a largerlip. Test measurements consisted of inlet surface static pressure, engine face total pressure,inlet airflow, tunnel total pressure, tunnel total temperature and tunnel velocity. Totalpressure traverses were taken directly behind the ring and strut. No dynamic measurements weretaken.

Test results indicate acceptable inlet total pressure recovery and total pressure distortion atstatic, forward speed, and angle-of-attack conditions. At cross-wind conditions, the two basicinlets had acceptable distortion at 11 knots cross wind, but the distortion became marginal whencompared to the Pratt & Whitney Aircraft limits at 20 knots cross wind.

17. Key Words (Suggested by Author(s)) 18. Distribution Statement

727 AirplaneSide Inlet Unclassified - UnlimitedInlet Pressure Recovery and DistortionRefanned JT8D EngineInlet Wind Tunnel Test

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages 22. Price'

Unclassified Unclassified 80

* For sale by the National Technical Information Service, Springfield, Virginia 22151

NASA-C-168 (Rev. 6-71)

FOREWORD

The low-speed wind tunnel test described in this report was

performed by the Propulsion Technology Staff of the Boeing

Commercial Airplane Company,A Division of The Boeing Company,

Seattle, Washington. The work, sponsored by NASA Lewis Research

Center and reported herein, was performed between October 1973

and February 1974.

This report has been reviewed and is approved by:

L. J. inslow, Group Engineer DatePropulsion Technology Staff

J. . F rell DateChief, taff TechnologyJT8D Refan Porgram

K. P. Rice DateProgram ManagerJT8D Refan Program

Preceding page blank

III

TABLE OF CONTENTSPage

1.0 SUMMARY -------------------------------------------- 1

2.0 INTRODUCTION --------------------------------------- 5

2.1 BACKGROUND ------------------------------------ 5

2.2 INLET DESIGN ---------------------------------- 6

2.2.1 Design Constraints and Goals ----------- 6

2.2.2 Design Procedure ----------------------- 7

3.0 MODEL AND TEST DESCRIPTION -------------------------11

3.1 MODEL DESCRIPTION AND MODEL INSTRUMENTATION---- 11

3.2 TEST FACILITY AND FACILITY INSTRUMENTATION-----12

3.3 TEST PROCEDURE AND TEST CONDITIONS ------------ 13

3.4 DATA REDUCTION AND PRESENTATION --------------- 15

4.0 RESULTS AND DISCUSSION ----------------------------- 17

4.1 SURFACE MACH NUMBER DISTRIBUTION -------------- 17

4.2 TOTAL PRESSURE RECOVERY ----------------------- 17

4.3 TOTAL PRESSURE DISTORTION -------------------- 19

4.3.1 Engine Face Pressure Recovery Maps ----- 19

4.3.2 Pratt & Whitney Aircraft Distortion-----20Criteria

4.4 COMPARISON WITH EXISTING 727 SIDE INLET ------- 23

4.4.1 Inlet Geometry and Airflow ------------- 23

4.4.2 Total Pressure Recovery ---------------- 24

4.4.3 Total Pressure Distortion -------------- 24

5.0 CONCLUSIONS ---------------------------------------- 25

6.0 FIGURES---------------------------------------------27

APPENDIX A - SYMBOLS ------------------------------------ 71

APPENDIX B - INLET COORDINATES -------------------------- 75

RFFFRENCES ------------------------------------------------ 79

Preceding page blankV

1.0 SUMMARY

A low-speed performance confirmation model test was conducted

for a Boeing 727 airplane side engine inlet designed for use

with refanned JT8D engines (designated the JT8D-100 series).

The test was conducted in the Boeing Low-Speed Wind Tunnel

located at North Boeing Field, Seattle, Washington.

The objectives of the test were:

* To develop lines for a full-scale flightworthy inlet, for

ground test, with performance (total pressure recovery and

distortion) comparable to the existing 727 side inlet.

* To evaluate inlet total pressure recovery and steady-state

total pressure distortion of the side engine inlet at JT8D-100

engine airflows at low-speed and cross-wind conditions.

* To obtain model-scale distortion data which can be used in the

assessment of the compatibility of the inlet with the

JT8D-100 series engines.

A secondary objective of the test was to obtain internal/

external cowl static pressures for the determination of nacelle

loads.

Two basic 0.3 scale models (AHI/ATH = 1.25), without acoustic

treatment, were tested at static, forward speed, angle-of-attack

(inflow angle), and cross-wind conditions. The two models were

without an acoustic ring (Configuration 1) and with an acoustic

ring (Configuration 2). Two inlets were designed and tested so

that full-scale ground test hardware fabrication could proceed

on the basis that either inlet Configuration 1 or 2 could be

ground tested. Two modifications to these basic models were

tested at a selected number of conditions. The modifications

consisted of using a larger lip (Configuration 1L,

AHI/ATH = 1.30) to investigate improved cross-wind performance;

and moving the acoustic ring leading edge closer to the throat

(Configuration 2R) to obtain more acoustic surface area.

Test measurements consisted of inlet internal and external sur-

face static pressure, engine face total pressure, inlet airflow,

tunnel total pressure, tunnel total temperature, and tunnel

velocity. Total pressure traverses were taken directly behind

the ring and strut. No dynamic measurements were taken.

Conclusions drawn from the test are:

* The cruise and angle-of-attack recovery and distortion of

the Configuration 1 inlet, accounting for acoustic treat-

ment, will be slightly better than that of the existing 727

inlet.

* The cruise recovery of the Configuration 2 inlet, excluding

acoustic treatment, will be slightly lower (0.003) than that

of the Configuration 1 inlet.

* The Configuration 1 and 2 inlets have an acceptable distortion

at 11 knots cross wind, but the distortion becomes marginal

when compared to the Pratt & Whitney Aircraft limit at

20 knots cross wind. The 30 knot cross-wind condition at

70 knots forward speed (cross wind/rolling takeoff) shows an

acceptable distortion level. The two inlets were designed

using a throat Mach number and lip contour that give an

equivalent or slightly better cross-wind performance than

the existing 727, and using normal operating procedures

cross-wind performance should not present a problem. For a

small cruise performance penalty, Configuration LL will

provide additional static cross-wind capability (up to

30 knots).

2

* The acoustic ring of Configuration 2 provides a baffling ef-

fect, when compared to Configuration 1, keeping the low

pressure distorted flow in the outer annulus. This results

in a "clean" core flow at all test conditions.

* The Configuration 2R inlet showed that the ring leading edge

may be moved further into the throat, thus providing more

acoustic treatment area, without a static, cross-wind or

angle-of-attack recovery or distortion penalty.

3

2.0 INTRODUCTION

2.1 BACKGROUND

The Pratt & Whitney Aircraft JT8D-100 engine is a derivative of

the basic JT8D turbofan engine modified to incorporate a new

larger diameter, single-stage fan with a bypass ratio of 2.0

and two supercharging low-pressure compressor stages. The

modification lowers jet noise, increases takeoff and cruise

thrust, and lowers specific fuel consumption. The use of a

JT8D-lOO series engine, also referred to as a refanned JT8D,

on the Boeing 727 airplane requires a larger side engine inlet

due to the increased engine diameter and increased engine air-

flow. A new airplane nacelle for JT8D-100 series engines is

currently being developed through a joint Boeing/NASA effort.

Two side inlet configurations have been designed for the

nacelle. The inlets were designed considering inlet acoustic

performance. The design and test results of the two inlets,

along with test results for two modifications to these inlets,

are the subjects of this report. The test evaluated the inlet

internal aerodynamic performance at low-speed and cross-wind

conditions using 0.3 scale models. The test was conducted

during November and December 1973, in the Boeing Low-Speed

Wind Tunnel, located at North Boeing Field, Seattle, Washington.

The objectives of the test were:

* To develop lines for a full-scale flightworthy inlet, for

ground test, with a performance (total pressure recovery and

distortion comparable to the existing 727 side inlet.

* To evaluate the inlet total pressure recovery and steady-

state total pressure distortion of the side engine inlet

at JT8D-1OO0 engine airflows at low-speed and cross-wind

conditions.

Preceding page blank5

* To obtain model-scale distortion data which can be used in the

the assessment of the compatibility of the inlet with the

JT8D-100 series engines.

A secondary objective of the test was to obtain internal/external

cowl static pressures for the determination of nacelle loads.

This test was performed under authorization of NASA Contract

NAS33-17842, "Phase II Program on Ground Test of Refanned JT8D

Engines and Nacelles for the 727 Airplane," to support the

development of a new 727 side engine inlet.

2.2 INLET DESIGN

2.2.1 Design Constraints and Goals

The following constraints were imposed on the side inlet design:

* The inlet diameter at the engine face (D2 ) shall be 50.1

inches.

* To avoid interference with galley door access the inlet

length to diameter ratio (L/D 2 ) shall not be greater than

0.8.

* Two inlets shall be designed, one with and one without an

acoustic splitter ring.

The inlet corrected airflow capability shall be as follows:

(1) 467 lb/sec at takeoff, sea level static condition,

std. day.

(2) 480 lb/sec at MCR, 0.8Mo , 30,000 ft., std. day.

(3) 501 lb/sec at MCT, 0.6M co, 35,000 ft., std. day.

The maximum JT8D-100 engine cold day airflow at both

sea level and 10,000 ft., -60OF ambient temperature, is

516 lb/sec. Applying a +3 percent production engine

airflow tolerance results in a 531.5 lb/sec maximum

airflow.

6

The inlet angle-of-attack (inflow angle) capability shall be

greater than or equal to the angles measured during the

Reference 1 wind tunnel flowfield test. The maximum angle

measured during the test was found to be a 17 degree down-

wash angle relative to the engine centerline during unstalled

wing operation.

The following design goals were set:

* A cross-wind capability equivalent to or better than the

existing 727 production side inlet.

Total pressure distortion and recovery levels, for the inlet

without an acoustic ring, equivalent to or better than the

existing 727 production inlet.

2.2.2 Design Procedure

A number of inlets, both with and without acoustic splitter

rings, have been designed and model tested during past studies

at the Boeing Company. The design procedure evolved from these

past studies was generally followed in this design. This

involves laying out a lip, diffuser, and centerbody contour

which accounts for the following:

* internal flow considerations

* acoustic considerations

* engine airflow schedule

e inlet angle-of-attack (inflow angle)

* inlet cross-wind capability

* external aerodynamic considerations

* manufacturing considerations

The empty inlet potential flow field is then computed using an

axially symmetric, compressible, potential flow computer analysis.

The contours are then modified if the potential flow results

indicate flow regions which may be improved. If the inlet is

to be used with an acoustic ring the empty inlet contours are

7

designed so as to accommodate the added blockage due to the ring.

The ring contour is then wrapped around the potential flow

streamline (particular streamline dictated by acoustic consid-

erations) with a new potential flow field being computed after

the ring insertion. The placing of the splitter around the

natural streamline should produce the least amount of flow

disturbance. Further refinements in the contours are then made

if flow field results indicate improvements can be made. No

account is made in the potential flow computations for the

struts which support the ring.

Based on the Reference 2 high speed nacelle drag test it was

determined that axially symmetric side inlets with a highlight

diameter of 52.2 inches would be satisfactory. A lip contraction

ratio (AHI/ATH) of 1.25, identical to the present 727, was

selected since the present 727 lip has not presented problems

during operation. Based on other Boeing studies a super ellipse( )2.2 + (y ) 2.2 = 1] was selected for the lip contour.

The highlight diameter and specified contraction ratio yield a

throat area (1712.81 inches2 ) which gives a one dimensional

cruise throat Mach number similar to the existing 727 with

JT8D-l or -7 engines. At a corrected airflow of 480 lb/sec.

the JT8D-100 inlet throat Mach number is 0.57. Based on manu-

facturing considerations it was decided to make all internal

surfaces axially symmetric. This could be done and still meet

the angle-of-attack and cross-wind objectives discussed in

Section 2.2.1.

The design procedure was applied to develop inlets with and

without an acoustic ring. Early in the design it was decided

that a common lip, diffuser, and nose dome would be used for

the two inlets. This common aspect makes it possible to decide

later in the development phase whether or not to use an acoustic

ring. It also makes it impossible to truly optimize both

8

configurations, though with the relatively thin acoustic ring

(0.7 inches full scale) the penalty is small.

The two inlet configurations are shown in Figure 1.

Configuration 1 is without and Configuration 2 with the acoustic

ring. Figures 2 and 3 show the Mach number distributions obtained

from the potential flow analysis. All analyses were done using

the 480 lb/sec cruise corrected airflow. The Configuration 2

ring leading edge location was selected based on the inter-

section of the near static (V, = 15 knots) and forward speed

(Moo = .8) streamlines as shown on Figure 2, and, on past Boeing

work. The ring trailing edge is one inch upstream of the engine

face. This provides ring removal clearance with a near maximum

of acoustic surface for the inlet length constraint. This also

allows a substantial distance (approximately 14 inches) between

the ring trailing edge and the first stage fan. The ring radial

location, determined from acoustic considerations, is on the

streamline which is equidistant from the cowl and nose dome at

the engine face.

For reference purposes the one dimensional throat and engine

face Mach numbers as a function of corrected airflow and the

one dimensional area distribution for the two inlets are shown

in Figure 4. It should also be pointed out that on the refan

engine the nacelle slopes 3 degrees out from the airplane body

buttock line and 3 degrees 22 minutes up from the airplane

body water line. This is approximately the same on the

existing 727 nacelle.

3.0 MODEL TEST DESCRIPTION

3.1 MODEL DESCRIPTION AND MODEL INSTRUMENTATION

The inlets tested were 0.2994 scale axially symmetric aluminum

models. All wall surfaces were hardwall, therefore without

acoustic material. In addition to the Configuration 1 and 2

inlets described in the previous section two modifications to

these basic inlets were also tested. An additional lip

(AHI/ATH = 1.30, Configuration IL) was tested at cross-wind

conditions to determine the potential for improved cross-wind

performance. A ring and centerbody assembly (Configuration 2R),

built for the NASA Phase I Refan Program, was tested to evaluate

extending the ring leading edge closer to the inlet throat. All

four inlet configurations, shown in Figure 5, used a common dif-

fuser. Configurations 1 and 2 internal contours are identical

to that currently being developed for the full-scale ground

test. The strut contour and strut circumferential spacing

have been changed slightly on the full-scale Configuration 2

inlet. A tabulation of the coordinates for all four models is

given in Appendix B.

The models were tested without fuselage simulation at static,

forward speed, and angle-of-attack conditions. The airplane

fuselage was simulated during static and cross-wind testing,

the windward side inlet being tested. Photos of the model in

the cross-wind and forward speed installation and a view into

the inlet are shown in Figures 6, 7, and 8 respectively.

Figure 9 shows a sketch of the inlet in the cross-wind and forward

speed installations.

A total of 57 static ports were utilized on the Configuration 1

inlet, 77 on Configuration 2. Configuration 1L and 2R utilized

the eight engine face static pressure ports only. Figure 10

shows the location of the static ports.

Preceding page blank 1 D

Engine face total-pressure measurements were made using an

existing fifteen-inch diameter, four-arm (16 probes per arm)

rotating rake. The closest total pressure probes to the cowl

and nose dome being 0.15 inches model scale. Measurements were

taken at angular increments of 15 degrees. For a portion of

the test conditions two total pressure traversing probes,

located 180 degrees apart at the engine face, were used to

obtain a detailed definition of the strut and ring wakes.

Figure 11 shows a sketch of the rotating rake and traversing

probes. Only steady state measurements were taken during the

test.

3.2 TEST FACILITY AND FACILITY INSTRUMENTATION

The test was conducted in the Boeing Low-Speed Wind Tunnel "B"

facility. The tunnel contains a 9 foot by 9 foot square test

section. An Allison 501-D13 gas turbine driving a variable

pitch propeller is used to draw atmospheric air through the

bellmouth, flow straighteners, and test section. The test section

airspeed could be varied from approximately 0 to 180 knots.

Engine airflow simulation was obtained by using a General Electric

J-47 turbojet engine as a pumping source. Air was drawn through

the inlet model, a venturi meter and into the engine. Venturi

airflow measurements were accurate to within +1 percent of the

measured value. A 19 percent blockage screen (0.41 inch mesh

with 0.041 inch diameter wire) was used for most test conditions.

The screen was located as near the fan face station as possible

(4-3/4 inches model scale from the engine face) in order to

simulate the upstream effects of the fan rotor on the flow. With

the screen installed the inlet choked at a simulated full scale

corrected air flow of 502 lb/sec. Without the screen airflows in

excess of 570 lb/sec. were possible.

Tunnel total and static pressure, tunnel total temperature and

venturi temperatures and pressures were recorded for each test

12

condition. This steady-state data along with the model steady

state data were recorded on the standard 9 x 9 Low Speed Wind

Tunnel data acquisition system. This system, a Hewlett-Packard

Dymec 2010D, is a trap and scan scannivalve system with output

on punched paper tape. The capability of monitoring on-line

engine RPM and a selected number of static pressures was available

for setting test conditions.

3.3 TEST PROCEDURE AND TEST CONDITIONS

After the model was assembled, leak checked, and the instrumen-

tation zero checked, the tunnel velocity and inlet airflows were

set and allowed to stabilize for at least 30 seconds. Data were

then recorded at various rake angles (at 15 degree increments

between 7.5 and 82.5 degrees). Model surface pressure measure-

ments were recorded at the first rake angle only.

Data were taken for static, crosswind and angle-of-attack

(0 to 22.5 degrees) conditions. Data were taken for a full

scale corrected airflow range of 180 to approximately 502 lb/sec

with the screens and up to 560 lb/sec without the screens.

Table 1 shows a summary of the test conditions at which the

various models were tested. As shown on Table I the inlet angle

of attack is actually a downwash due to the presence of the wing.

Because of facility limitations all angles were in the upwash

direction for the test model. Thus to go from the model test

configuration to the airplane configuration all angles-of-attack

results must be rotated by 180 degrees. The 76 knot, 22.5 degree

installation simulates a 30 knot cross-wind 70 knot forward speed

condition. The higher velocities at 22.5 degrees were run for

the determination of nacelle loads. The nacelle loads results

are reported in Reference 3.

13

TABLE I

CONFIGURATION NO. I CONFIGURATION NO. 2INLET ANGLE-OF-ATTACK INLET ANGLE-OF-ATTACK

RUN NUMBER -o*, 17.*, 22.s5 , . RUN NUMBER .-o*, 14.8", 17.5", 22.5, o*

B0, o, o', 9 0* 0-o, o', o*, o', 9o0

0 I 0 39 1342

II 4 10 14

1520 3 20 18

16

30 20 19

35 I45 35 38

8 70 1Q , 70 37

76 47 76 26

28 22180 43 4 " 100 Q) 31

*33 27

120 23CONFIGURATION NO. IL

INLET ANGLE-OF-ATTACK 150 49 32 24

RUN NUMBER .0' 34

0-o90 180 36 50 30 (S

0 S 40 5 1 3 5 2 5

CONFIGURATION NO. 2R

10 RUN NUMBER INLET ANGLE-OF-ATTACKSa-17.5*

, 0*

20 0, 90

I-bd 0 9

30 II

2 20 10

NOTES: 3 .

(D Cross wind (Bg 90 ) with 180 52

fuselage except as noted,.other conditions withoutfuselage.

) without screens also.

) Without fuselage. tWLEI) Static pressure data only. OWL

$ Without screens only. 3 22'

14

3.4 DATA REDUCTION AND PRESENTATION

During the test program data were reduced using a standard

Boeing data reduction computer program for inlet tests. Data

were reduced using a quick look and final reduction version of

the program.

Quick-look data were obtained by processing the punched paper

tape through the Boeing Mechanical Laboratories. SDS 92 computer.

The tabular output consisted of total and static pressure

measurements, surface Mach number distributions, surface pres-

sure coefficients, surface pressure ratios, the rotating rake

pressure array, inlet recovery, inlet airflow, and the steady-

state distortion parameter (PT Max - PT Min)/PT2.

Additional quick look (on line) data were obtained from the test

facilities own PDP8 computer. This consisted of tunnel condi-

tions, inlet airflow, and the rotating rake pressure array.

Final data were obtained by generating a magnetic tape from the

paper tape and processing it through the Boeing CDC 6600 computer.

The final data consisted of tabular information similar to that

obtained from the quick look data plus the distortion para-

meters defined by Pratt & Whitney Aircraft (see Section 4.3.2).

Engine face plots showing lines of constant total pressure

recovery were also generated on the CDC 6600 computer. All

final data are permanently stored on microfilm.

The total pressure recovery measurements (PT2 /PT) presented in

this document are computed on an area-average basis. The cowl

wall region is handled by taking the average of the wall static

pressure measurement and the closest total pressure probe

multiplied by the annular area segment between the two. Inter-

15

mediate regions are handled by multiplying the annular area

segment between any two probes by the average of their total

pressures. The nose dome wall region is computed by taking

a straight line fit from the last two total pressure probes to

the wall. The nose dome was handled in this manner because its

boundary layer is very thin, and of a high power law exponent,

and straight line fitting to the wall static pressure would

over estimate the loss in this region.

All airflow data presented in this report have been converted

from 0.2994 model scale to full scale values.

16

4.0 RESULTS AND DISCUSSION

4.1 SURFACE MACH NUMBER DISTRIBUTION

Surface Mach number distributions at zero and 17.5 degrees angle

of attack are shown in Figures 12 and 13 for Configurations 1 and

2 respectively. The analytical results plotted in the figures,

are in good agreement with the data for the region from near the

throat to the engine face. In the lip region the results differ

since the analysis was done for a Moo = .8 whereas the data is

for a VT = 180 knots. At 17.5 degrees inlet angle of attack lip

Mach numbers in excess of 1.3 were measured on the lower lip,

the internal surface Mach numbers downstream of the throat re-

maining approximately the same.

4.2 TOTAL PRESSURE RECOVERY

Configuration 1 total pressure recoveries at static, cross-wind,

forward speed, and angle-of-attack conditions are shown in

Figures 14 and 15. As shown in Figure 14 at cross-wind

velocities greater than 11 knots the recovery deteriorates

rapidly. Figure 15 shows the recovery at the static condition

with and without the fuselage. Without the fuselage a vortex

was visually observed at static conditions coming off the tunnel

side wall. This was also observed in the engine face measure-

ments and accounts for the lower recovery without the fuselage

at higher inlet airflows. At the angle-of-attack conditions,

of 22.5 and 17.5 degrees the inlet shows high recoveries. At

the 70 knot forward speed condition, shown on Figure 15, the

inlet recovery level is nearly as high as that measured at

180 knots.

Configuration 2 total pressure recoveries at static, cross-wind,

forward speed, and angle-of-attack conditions are shown in

Figures 16 and 17. The same type of statements can be made

17

concerning the Configuration 2 inlet performance as were made

for the Configuration 1 inlet, except, that due to the ring

surface area the recovery level is lower. There was no

evidence of ring leading edge separation at the static, forward

speed or angle-of-attack conditions. The inlet was tested with-

out the screens and without the fuselage at static and cross-

wind conditions. For the static condition the presence of the

screens made little difference in the recovery; but, at the cross-

wind condition of 30 knots the screens provided an improvement

in the recovery as shown in Figure 16. The presence of the

fuselage provided some improvement at the static condition as

shown in Figure 17 and a substantial improvement at the 20 and

30 knot cross-wind conditions as shown in Figure 16. The inlet

recovery did not vary substantially over the range of tunnel

velocities and angle-of-attack conditions tested as shown in

Figure 17. As with the Configuration 1 inlet the Configuration 2

pressure recovery at 70 knots forward speed is nearly as high

as that at 180 knots.

Total pressure traverses of the ring wake and strut wakes were

taken at the engine face for the Configuration 2 inlet.

Figure 18 shows a portion of the results. At the cruise air-

flow a minimum pressure ratio (PT/PTc) of 0.88 was measured

behind the ring and 0.958 behind the struts. By integrating

the ring and wake traverse measurements of Figure 18 (assuming

the wake does not change along the trailing edge) the loss due

to the strut and ring may be estimated. The cowl and nose dome

loss may be obtained from rotating rake measurements, along with

a second measurement of the ring loss. When this is done the

total pressure loss (1-recovery, where the recovery is as

discussed in Section 3.4) for each component of the Configuration 1

and 2 inletsat a corrected airflow of 481 lb/sec is as

follows:

18

Loss (1-Recovery)

Cowl 0.0050 } Configuration 1 and 2Nose Dome 0.0000 JRing 0.0030Struts 0.0003 JConfiguration 2

Total 0.0083

The static and cross-wind recovery measurements of the Configu-

ration 1L inlet are shown in Figure 19. Because of the high

contraction ratio (AHI/ATH = 1.30) the inlet showed high recovery

levels even at the 30 knot cross-wind condition.

The static, angle-of-attack and cross-wind recovery of the

Configuration 2R inlet is shown in Figure 20. The recovery

levels measured on the 2R model are comparable to the Configu-

ration 2 model indicating that the Configuration 2 acoustic ring

leading edge may be extended closer to the throat.

4.3 TOTAL PRESSURE DISTORTION

4.3.1 Engine Face Pressure Recovery Maps

Engine face pressure recovery maps show lines of constant steady

state total pressure recovery at the engine face station. Such

maps are useful as a visual aid. Although recovery maps were

made for all test conditions only a select few are shown. These

consist of:

Figure 21 Configuration 1 at 180 knots and ff= 17.50

Figure 22 Configuration 1 at 20 knots cross wind

Figure 23 Configuration 2 at 180 knots and d= 17.50

Figure 24 Configuration 2 at 20 knots cross wind

Figure 25 Configuration 1L at 30 knots cross wind

19

The Configuration 1 recovery map, Figure 21, shows a small low

pressure region downstream of the lower lip during angle-of-

attack operation. Figure 22 shows a large low pressure region

towards the windward side at a 20 knot cross-wind condition.

The Configuration 2 recovery map, Figure 23, shows a similar low

pressure region, as that of Configuration 1, at the lower lip

during angle-of-attack operation. Also the low pressure region

directly behind the acoustic ring can be seen in the figure. The

strut wakes do not show since the rotating rake was oriented so

as to avoid taking measurements directly behind the struts.

Figure 24 shows the Configuration 2 inlet at a 20 knot cross-

wind condition. As shown on the figure the ring provides a

baffling effect, keeping the distorted low pressure flow in the

outer annulus.

Figure 25 shows the Configuration LL inlet at a 30 knot cross-

wind condition. Even with the high cross flow velocity, and

high inlet airflow (481 lb/sec), the low pressure region remains

relatively small.

4.3.2 Pratt & Whitney Aircraft Distortion Criteria

Radial and circumferential total pressure distortion parameters

and limits have been defined by Pratt & Whitney Aircraft for the

JT8D-100 series engines. The distortion parameters and limits,

described in Reference 4, are for instantaneous total pressures.

The radial distortion limit of Reference 4 has since been updated

to the limit shown in this document. The parameters are defined

as:

* Radial Distortion =

PT Max Ring Avg - PT Local Ring Avg

T Max Ring Avg

where ring averages are taken over a full 3600

20

* Circumferential Distortion =

PT Ring Avg - PT Min Sector Avg

T Ring Avg

where ring averages are taken over a full 3600 and Min Sector

Avg is the lowest average total pressure at a given radius

for the sector of concern (either a 60 degree or 180 degree

sector).

During this test only steady-state total pressure measurements

were taken, consequently the radial and circumferential distortion

data presented are based on steady-state values only.

Distortion values were computed from the data based on the Pratt &

Whitney Aircraft definitions and are presented for the following

conditions:

Figures 26, 27 and 28 Configuration 1, 60 degree circumfer-

ential, 180 degree circumferential, and

radial at static, forward speed, and

angle-of-attack conditions.

Figures 29, 30 and 31 Configuration 1, 60 degree circum-

ferential, 180 degree circumferential,

and radial at cross-wind conditions.

Figures 32, 33 and 34 - Configuration 2, 60 degree circum-


and radial at static, forward speed,

and angle-of-attack conditions.

Figures 35, 36 and 37 - Configuration 2, 60 degrees circum-


and radial at cross-wind conditions.

21

Figures 38, 39 and 40 - Configuration IL, 60 degrees

circumferential, 180 degree circum-

ferential, and radial at static and

cross-wind conditions.

The Configuration 1 inlet distortion falls well below the limit

line at static, forward speed, and angle-of-attack conditions

as shown in Figures 26, 27, and 28. This is true even for the

30 knot cross-wind 70 knot forward speed (cross wind/rolling

takeoff) condition. The steady-state distortion is so low that

even if a dynamic component were measured and an instantaneous

distortion value computed it is expected that the distortion

would remain below the limit. This is also the case for the

11 knot cross-wind condition shown on Figures 29, 30 and 31.

At 20 knots cross wind the 60 degree sector distortion surpasses

the limit as shown on Figure 29. The 180 degree and radial

distortions fall below the limit line as shown on Figures 30

and 31.

The Configuration 2 inlet distortion, with the exception of the

region directly behind the ring (probes 7, 8 and 9), falls well

below the limit lines at static, forward speed, and angle-of-

attack conditions as shown on Figures 32, 33 and 34. This is

expected to be true even if the dynamic component were considered.

The ring wake may be excluded from the limits shown, and when the

limit is exceeded specific distortion patterns will be submitted

to Pratt & Whitney Aircraft for evaluation (Reference 4). The

circumferential distortion behind the ring shown on Figures 32

and 33 is more a measure of the asymmetry of the rotating rake

and the ring rather than a true distortion. For the same reason

the Figure 34 radial distortion behind the ring will be somewhat

larger than that shown. For example the traverse data of

Figure 18 (run 30.5) indicates a radial distortion directly behind

the ring of 12 percent. Whereas, the rotating rake data of

Figure 34 (run 36.5) indicates a radial distortion of 10.8 per-

cent directly behind the ring (probe 8).

22

The Configuration 2 inlet, unlike Configuration 1, falls below

the limit line at 20 knots cross wind as shown in Figures 35, 36

and 37 (run 15.3). The 60 degree sector distortion is near the

limit line and if the dynamic component were accounted for it may

surpass the limit. The 10 knot cross-wind data shown on the

figures (run 14.3) falls well below the limit and if the dynamic

component were considered it is expected that it would remain below

the limit. As with the pressure recovery an improvement was

obtained in the distortion at cross-wind conditions with the

fuselage as shown in the figures.

The Configuration IL inlet falls well below the distortion limit

even up to 30 knots cross wind as shown in Figures 38, 39 and 40.

Even with a dynamic component it is expected that it will fall

well below the limit.

4.4 COMPARISON WITH EXISTING 727 SIDE INLET

4.4.1 Inlet Geometry and Airflow

The JT8D-1OO0 inlet was designed to give internal performance

similar to the production inlet. Consequently many of the design

features are the same. Figure 41 compares the JT8D-lO0 inlet

with the production inlet on a JT8D-15 engine. Depending on

which basic JT8D engine is considered, the JT8D-1OO0 inlet

provided 40 to 48 percent more airflow and is consequently larger

in diameter. The lip contraction ratios of the two inlets and

cruise throat Mach number (when compared to the JT8D-1 and -7)

are the same. Because of galley door access interference problems

the JT8D-1O0 inlet was limited to 40 inches in length. This

is adequate for good internal performance but limits the surface

area available for acoustic treatment.

The production inlet has 40 of turning between the engine face and

highlight. The JT8D-100 is axially symmetric.

23

4.4.2 Total Pressure Recovery

A comparison between the total pressure recovery of the production

and the JT8D-100 inlets is shown on Figure 42. The production

data were taken in the same tunnel facility, for approximately

the same model scale (0.3 compared to 0.37), with comparable

engine face probe placement and data reduction. The data presented

for the production inlet is for a slightly modified model. The

modifications consisted of the cowl diffuser surface being

dished out somewhat, to accommodate acoustic rings in subsequent

testing, and a longer centerbody. As shown in Figure 42 the

JT8D-100 Configuration 1 inlet has a slightly higher recovery,

at zero and 17.5 degrees angle of attack at cruise airflows,

than the modified production model. The Configuration 2 recovery

is slightly lower (.003) than the Configuration 1 inlet at the

cruise airflow.

4.4.3 Total Pressure Distortion

A comparison of the total pressure distortion for the JT8D-100,

the modified production, and the full scale production inlet is

shown in Figure 43. The distortion parameter, shown in the

figure, is for the (average-minimum) /average total pressure ratio

which has been used in past inlet development testing. As shown

in the figure the Configuration 1 model distortion is lower than

that of the modified production inlet.

24

5.0 CONCLUSIONS

The cruise and angle-of-attack recovery and distortion of the

hardwall Configuration 1 inlet will be slightly better than

that of the existing 727 hardwall inlet. Since the new 727

inlet utilizes a maximum of peripheral treatment the same

statement will be true when acoustic treatment is included for

both inlets. The hardwall Configuration 2 cruise recovery will

be slightly lower (0.003) than that of the Configuration 1 inlet.

The Configuration 1 and 2 inlets have an acceptable distortion

at 11 knots cross wind, but the distortion becomes marginal when

compared to the Pratt & Whitney Aircraft limit at 20 knots cross

wind. The 30 knot cross-wind condition at 70 knots forward

speed (cross-wind/rolling takeoff) shows an acceptable distortion.

level. The two inlets were designed using a throat Mach number

and lip contour that give an equivalent or slightly better cross-

wind performance than the existing 727, and using normal operating

procedures cross-wind performance should not present a problem.

For a small cruise performance penalty, Configuration 1L will

provide additional static cross-wind capability (up to

30 knots).

The acoustic ring of Configuration 2 provides a baffling effect,

when compared to Configuration 1, keeping the low pressure

distorted flow in the outer annulus. This results in a "clean"

core flow at all test conditions.

The Configuration 2R inlet showed that the ring leading edge may

be moved further into the throat, thus providing more acoustic

treatment area, without a static, cross-wind, or angle-of-

attack recovery or distortion penalty.

25

6.0 FIGURES

Figure No. Title Page

1 CONFIGURATIONS 1 AND 2 JT8D-lOO SIDE INLET 30

2 CONFIGURATION 2 PREDICTED MACH NUMBER 31DISTRIBUTION

3 CONFIGURATIONS 1 AND 2. SURFACE MACH NUMBER 32DISTRIBUTION (Potential Flow Analysis)

4 CONFIGURATIONS 1 AND 2 AREA DISTRIBUTION AND 33MACH NUMBER

5 SIDE INLET TEST CONFIGURATIONS 34

6 CROSS-WIND INSTALLATION 35

7 CONFIGURATION 2 FORWARD SPEED INSTALLATION 35

8 INLET CONFIGURATION 2 35

9 WIND TUNNEL CROSS-WIND AND FORWARD SPEED 36INSTALLATIONS

10 CONFIGURATIONS 1 AND 2 STATIC PRESSURE 37PORT LOCATIONS

11 ROTATING RAKE CONFIGURATION 38

12 CONFIGURATION 1 SURFACE MACH NUMBER 39DISTRIBUTIONS

13 CONFIGURATION 2 SURFACE MACH NUMBER 40DISTRIBUTIONS

14 CONFIGURATION 1 PRESSURE RECOVERY VS 41AIRFLOW (Static & Cross-Wind Conditions)

15 CONFIGURATION 1 PRESSURE RECOVERY VS 42AIRFLOW (Static, Fwd Speed and Angle-of-Attack Conditions)

16 CONFIGURATION 2 PRESSURE RECOVERY VS 43AIRFLOW (Static and Cross-Wind Conditions)


27


17 CONFIGURATION 2 PRESSURE RECOVERY VS AIR- 44FLOW (Static, Fwd Speed and Angle-of-Attack Conditions)

18 CONFIGURATION 2 ENGINE FACE RING AND STRUT 45TOTAL PRESSURE TRAVERSE

19 CONFIGURATION IL PRESSURE RECOVERY VS 46AIRFLOW (Static & Cross-Wind Conditiions)

20 CONFIGURATION 2R PRESSURE RECOVERY VS AIR- 47FLOW (Static, Angle-of-Attack and Cross-Wind Conditions)

21 CONFIGURATION 1 ENGINE FACE PRESSURE 48RECOVERY MAP (180 Knots @ d = 17.50)

22 CONFIGURATION 1 ENGINE FACE PRESSURE 49RECOVERY MAP (20 Knots Cross Wind)

23 CONFIGURATION 2 ENGINE FACE PRESSURE 50RECOVERY MAP (180 Knots @ i= 17.50)

24 CONFIGURATION. 2 ENGINE FACE PRESSURE 51RECOVERY MAP (20 Knots Cross-Wind)

25 CONFIGURATION 1L ENGINE FACE PRESSURE 52RECOVERY MAP (30 Knots Cross-wind)

26 CONFIGURATION 1 CIRCUMFERENTIAL PRESSURE 53DISTORTION 600 SECTOR (Static, FwdSpeed and Angle-of-Attack Conditions)

27 CONFIGURATION 1 CIRCUMFERENTIAL PRESSURE 54DISTORTION 1800 SECTOR (Static, Fwd Speedand Angle-of-Attack Conditions)

28 CONFIGURATION 1 RADIAL PRESSURE DISTORTION 55(Static, Fwd Speed and Angle-of-AttackConditions)

29 CONFIGURATION 1 CIRCUMFERENTIAL PRESSURE 56DISTORTION 600 SECTOR (Cross-WindConditions)

30 CONFIGURATION 1 CIRCUMFERENTIAL PRESSURE 57DISTORTION 1800 SECTOR (Cross-WindConditions)

28


31 CONFIGURATION 1 RADIAL PRESSURE 58DISTORTION (Cross-Wind Conditions)



34 CONFIGURATION 2 RADIAL PRESSURE DISTORTION 61(Static, Fwd Speed and Angle-of-AttackConditions)

35 CONFIGURATION 2 CIRCUMFERENTIAL PRESSURE 62DISTORTION 600 SECTOR (Static and Cross-Wind Conditions)

36 CONFIGURATION 2 CIRCUMFERENTIAL PRESSURE 63DISTORTION 1800 SECTOR (Static andCross-Wind Conditions)

37 CONFIGURATION 2 RADIAL PRESSURE DISTORTION 64(Static and Cross-Wind Conditions)

38 CONFIGURATION IL CIRCUMFERENTIAL PRESSURE 65DISTORTION 60 SECTOR (Static and Cross-Wind Conditions)

39 CONFIGURATION 1 CIRCUMFERENTIAL PRESSURE 66DISTORTION 180 SECTOR (Static and Cross-Wind Conditions)

40 CONFIGURATION 1L RADIAL PRESSURE 67DISTORTION (Static and Cross-WindConditions)

41 DESIGN COMPARISON OF 727 PRODUCTION 68AND JT8D-1OO SIDE INLETS

42 PRESSURE RECOVERY COMPARISON 727-200 69PRODUCTION (Modified) AND JT8D-1OO0SIDE INLETS

43 TOTAL PRESSURE DISTORTION COMPARISON 70727-200 PRODUCTION AND JT8D-100 SIDEINLETS

29

W

:jI :7:, :P.I e ;::I:::, 2.5:1 :T::

-:-! ... suerelipe::i::;-: i: :iiii-r - i I .: :

_ :_i: :- r !! :!:I:.7- .: l:: : : :::- . :.: . :~ :I: :mde tes only)

.. . ... ...... ... V -1 :!ii':" : ::::I

....~~~.. .. .... . .. I:F ( ).::::::: !:. ::- .: :: :;: :..: 00 9 JI O

1- -W .17-- 7--

.: w ._' : -. '2-L- :- _: - : .:: ::::---

[.:: : i:L ii:

Lip ,1 iticit 2.5:1 '7: - - -::!:: Modifie N J 0009i

--- .. ...-;: :super :ellipse - ro .

'7- :A- It i:-iii -i: - -

-25-- Ajj /A" ; ... ... . ..... . :-

I:: -.:-: ..... ....-- .-:! . ::: -- : ::

. . . ... . _, - . .. . .. , . . .... .. I- _-- :: - --- 7 -1

w- !:i : I" " 7 7: I i- ! !- i ::,-::- :_ :U::"::'Iii7:: .~~~~I ;:: ' jii-lji . I:l:: ii i~l~i~i fi :.:l.LI: .- li:i

:A::

' ~ ~ ~ ~ . .... ... ... : 5 "--. --- --*.: -:,- - ' -7_ . . ,0 . H._lt-": ~~r~ '-f~ '' jiii : : :! .: " i :. _..1':!:::: .l''iii:1: ;: .. :1:h:1::j~INCHES: _ (FUL SC LE) -i i--:!ii':i:: ": - lfii

.--.: .... .. ...i _ i::i.... ... ....F:GURE I("AA OwC AN TD 0 E L

2 15ad NACA 63 ::f i:j:

H. F:: t _w 7 -ii r

:~::::1=r 7! MoWdified NACA ~~

a :~: irfoil. T. E. r

C3 i- ;il Confiurati on no. 2 ring -7 7

" 10 .. .. ... ...

I~i L --- .7r

7= -7 77 .: ; i~

aw ::10

X . CHS FULSCL

FIGURE ii I CONFIURATONS AND2 JT8-1OOSIDEINLE

727 JT8D-100 SIDE INLET CONFIGURATION NO. 2 Wa 4/6t2 =48O0 Ib/sec M. 0.8

0 .3

25.00

2 o.00 _ _t_ ____5 0s0 502 0 50t 05

X .. -55

< 15.00 ___U I ( f a o

5.00

. 5.00 60.00 .00 so.00 46.00 40.00 3500 30.0.00 20 00 15-00 10 00 5 . 0

X q- INCHES FULL SCALE

FIG 2 PREDICTED MACH NUMBER DISTRIBUTION (Configuration 2)

CAiiiiiiiiii iii-~~~iiiii~iiiI~ii

Milt 4111111V II t t I

it i lit W ! lifii iiil~i !~tiif j ii iiiiiiit 'W 4146 1141 ilt :.t i:11il~i

It U411 ING jl R CONFIGMATION NO. 2 7 0 FI 1-f 1:: 1 ~liii it

tH if 14,11 T3lll

.1. ~~Insi e sur ace :i -:l -T:

.1 . i t ; :J - m -pi

CJ :A :1

i it TIT u~: ;utside surface..;:.LAJ i 1114 w I it i i ii I i 04 1 r- I :: :-

,,%(52 W bse,::::~ , it~ i~rr;:i~ -i:i::T;; '

- 1 . I ; ... I : I t i

::1 ::: iri~~~~~i~' il! il i .... ... il: :l.ll:; ll ;. 'ii '.::::r:: .:: IIj I -_

i+I

...... ........ .

.6t -t t'ii . - :, i i t

CL3 Conifi t i o

-.C ij j ;: L':. I . ;?I

tit it 11;ilice

~::I

4w !tt .......,

it -Iri~it :1 -1

"Ji. i;:T i i, 1 1

lit l:l't .,.I i ;FIGUR 3 COFIGUATION I AN 2 SRFACEMACH UMBEDITRBUIO (otntalFlw nayss

32Ij i ~ ii i i~

itt l t Lll :: I:

727 JTBD-100 SIE INLETI: 5E. IN E ... . .. 7-

'i::i:..;

T-- 550 .-.r

M i '

T.-. 1.0

: " ,i - ' . : :: .2 .

I.. ui P2 "/-- --=

!:L;.:q . ..... ],~:I:: _1:: .-_ "!" .

"'N~~~~_1 . T... T ... I(... ...

2 00~

::i! ii . l k .

T, INCHES FULL .......

7:l / .. . .. .............. ..

I7:

;::: J. 00 R1N! d* ( .;.-: W 2

7t7:

FIGURE 4 CONFGURATIONS AD 2RAREA DISTIUINADMAHNME

tj d. LA33

7--: T .. . .. . .

it .71111 1, NCHES W FULL SCAI ' :~t

7 i i''' I::~n lr

T!:!.

FIUR 4CNFGUAIOS AD2 RE ISRIUTO AD AH UME

iiilii~ii ' I; : :: t: : :: 1 ; :I: : :::~r~tI33_

CONFIGURATION I CONFIGURATION 2

THROAT THROAT ENQ FACE(33.2) ENO FACE (33.2)

20

25I 25. 090 30 20 10 0 40 30 20 10 0

X I - INCHES (FULL SCALE) X t - INCHES (FULL SCALE)

CONFIGURATION IL CONFIGURATION 2R

THROAT ENG FACE THROAT(33.2) (33.2) ENO FACE

30% LIP 25% LIP

20 1.2 9.0 20

COMMON DIFFUSER I 0 COMMON DIFFUSER

0 30 20 040 30 20 10 0 40 30 20 10 0

X ( - INCHES (FULL SCALE) X (- INCHES (FULL SCALE)

Notes:o Throat diameter M.70 in. (full scale) - all models.o Engine face cowl diameter 50.10 in. (full scale) - all models.o Nose dome engine face diameter 16.00 in. (full scale) - all models.o Common nose dome on configurations I, 2 and IL.o 25% lip defined as AI/AT : 1.25.

FIGURE5 SIDE INLET TEST CONFIGURATIONS

34

FIGURE 7 CONFIGURATION 2 FORWARDSPEED INSTALLATION

FIGURE 6 CROSS-WINDINSTALLATION

j. reproducea atThis page e by a iff ereilt

back of the repod tpovderepro ar lrobetter tail

FIGURE 8 INLET CONFIGURATION 2

35

CROSS-WIND INSTALLATION

Tunnel CSTunnel t

Ceiling 4 - 8

Fuselag/

3C~

103 Tunnel 4C cn103" ! "'

Side Inlett 36.6"

0

056 0=00

+0

Floor / Configuration 2Foo / Side Inlet

Flow

TUNNEL CROSS SECTION PLAN VIEW(LOOKING DOWN TUNNEL)

FORWARD SPEED ANGLE-OF-ATTACK INSTALLATION

Ceiling

A / EL-Deg.'A'- In B'- In.

Configuration 2 00 0 54.0 33.0

S14.8 40.6 46.4

S 46" 4----5 17.5 38.0 49.0

22.5 33.0 54.0

Floor

TUNNEL CROSS SECTION ALL DIMENSIONS INCHES MODEL SCALE

(LOOKING DOWN TUNNEL)

FIGURE 9 WIND TUNNEL CROSS-WIND AND FORWARD SPEED INSTALLATIONS

36

30

LA-

20 -

_J

0 =

ANGULAR10 SYM POSITION-e

20 v 170"v 180'

-30

I I I I I.40 30 20 10 0 NOT SHOWM * x=

X - INCHES FULL SCALE PORTS e. 35,80! 125, 21,26,305,350'ON COWL

PORTS * 035 80'ON NOSE DOME

COWL AIG NOSE DOME

o 0 0 0 a 0.20 I

0. t • • xch

2.0 I4.05.0 x x x x

08.0 A z

12.0 2

15.0 [ 2 x x16.0 x20.0 x x x x31.0 I22.4 I 224.0 S A 2 x25.2 x'6. 0 I26. 2227.2 a 129.0 x 231.0 2

u 33.2 2

34.25 xw 35.24 I

S 36.24 x 237.2 338.15 239.05 . 23 9

. 8 x 2

4 09

V 39.73 2 A37.97 x x36.07 x

FIGURE 10 CONFIGURATIONS 1 AND 2 STATIC PRESSURE PORT LOCATIONS

37

ROTATION y Steady state traversing probe

.063 in.

ARM #4 (277.5') NoseDome

- ARM #2 (97.50)

RCowl

19% blockage screen

Steady state traversing probe TYPICAL ARM TO SCALE

ARM #3 (187.5) 1806RI6 --

PROBE LOCATIONS* R -

FULL SCALE MODEL SCALE Ri -

RADIUS INCHES RADIUS INCHESRi3 -

RI 8.500 2.545

R 2 9.018 2.70

R 3 10.521 3.15 R12

R4 12.525 3.75

R5 14.529 4.35 RI - Rco. 25.05"

R 6 15.665 4.69 (Full ScaR7 16. 166 4.84 Rio - = 7.50"

R9 -- (R8 16.466 4.93 8 - (Model Scal

R9 16.767 5.02 7

R l0 17.268 5 17 R6R II 18.270 5.47 R5

R 12 19.940 5.97

R 13 21.610 6.47

R Iq 23.046 6.90 Ru -

R 15 24.015 7.19

R 16 24.549 7.35

R3 -

o All arms 900 aparto All four arms identical

o Probe size .062 inches max. R2 -

probe size O.D. (model scale) RI -o Probe tips are located in the C "

engine face RNose = 8.00" (Full SalejDome =2.3952" odel ale

FIGURE 11 ROTATING RAKE CONFIGURATION

38

I: *l . .:Ii i;;; '

'iii ii!i! It RN I I lOTSl:-1 b/sec ',:li::il

;I:- , : . 't'l_,-it .*li i i ii l ii . , .,: I1 ... ... . "

o.: poent ial fl w allyl 2 . :::

,:i;: i:. 4€,, )~t•j lit -" .Sm . se! ., : •

..... :3 43- _l.3, 8 ... !: 0 .- :-.- .I

if; :.c.)';ijij i: : I :: I: / :1 : 1 ' ; " I t 'it ........ . . ... .,5

Jii lit

'~~~ ..4:,:!

uJ::1;:!. :::: -:'! : :: '7 7 li "I' "

Z:

,I:: , 'i: ..7! ... ... ..

:... ... t.- :"

1. ,:,;€ -i t l l ,i i l i l Ii t l ;i:l t -f tl-,

7.:

-- L0t_ Lt] 1. I ,L[ i : ; ~:tl]L _ I_ -I I J 1 .L .1

:4 7i4 :

:: jj :: ...t..ftitl1i.iii ..!!{!!.: ',i : : I ;: : i •'

7 -7- 17 -. :2c

i! ta:

7 .

7 ..........

t--: ---- -----i-

1-77-T-= , 'r 7

it iI-: t

FIGURE~ ~ ~ ~ ~ ~~ ~I 12 COFGRTO 1 SUFCEM NME DSRBUIN

t.: t t39

'7 .8 1. .. ...

T7. : : I: I.

it .;-1 8 ... ... r'0 Externa Inte ai :

'-'30 i :Ii!; i ij4 1

x ES (ULL SI E)

... ... h it itt-..

FIGRE 2 CNFGURTIO ISURACEMAH NMBE DSTRBUTON

:iilii:/::iii:: ::::;:::::::::~.39

: : ::;;+ ,ItNelt~S ": i , ::IFI:

:: r , +,il,+::i~ !+iil .... .. :ii i +I ; e :: ::::';:! t i : t+:1': +. :-++-

:;:'-- " +- ' '! : : ": '+ ' !. ... . ...

+++I.-+ ! , + +..:; !l tu~1

tr NO 4;:'U G ~~ :~il

i t ! I, .. K 1.,

tt.4 1=1 U:: _H

,:::: " i '~ "l '-Poe tilflw n ly i--11.

i W t t!i , 1_0 L181fit: i 7

I. T14 W- ..11/ W4 41! ", Poenialo aals

S 1: i l l ljl j : 1: Ir l .: [i - i:

"::+t; '.1:ii~~ili :; '' I:

I....I 1 I :h i Ii .. ... H 8

; .I :t ' '

: -+z- - : I: F";

:, i= ": :i : : :1: I::j/ iii i

T7 14 -1: :I:

.. . . .. .....

-7,11 ... ... I 4

FIGURE !-7D

74

7:. .. Lr "I:: P .. ;

I: ,t-T :;I: :1::i I":I.:

T.;.. ~I-" ::1!::~:::1:: ll.' 7

::T

; : : :: : : :: :I . : : : ::iiil'iilii -L~~iii i: : :I':~Tl~r:i~7I ' - -:T 7Truji : l : : I:::; : : I

::::~~::'iiiiii :~il: : i-t-7

tiij~~ b: : :

"A 7:, '7.:

7, ~ ~ ~ : ' :~iii :p::;~3 i~ili rf ace Sur :; 7i:: :

14 it : 31

-c INCHES FULL ALEJ: :f-, j id: L: I: :: : L 4i'p

m,.J til ITT", I:d di 1111 "Id f I I F, 11 t41 mill RiW41 ii", I Hit !FIUE1 OFGRTIN2SRAEMC NME ITIUIN

40ii

100 ..

. ----99- ..

0

200 25 300 350 400 4 50 500 550 600CORRECTED ARFLOW Wa2 LB / SEC

(FULL SCALE)

FIGURE 14 CONFIGURATION 1 PRESSURE RECOVERY VS AIRFLOW (Static & Cross-Wind Conditions)

100 .. .... ..F0

0

u

LJ

w

w

" "

fr-

CORREC TED AlIRFLOW Wa4'O2/ 't2 " LB I SEC(FULL SCALE)

FIGURE 15 CONFIGURATION 1 PRESSURE RECOVERY VS AIRFLOW (Static, Fwd Speedand Angle-of-Attack Conditions)

EL

. . . . . . . . .....

200 250 300 350-U~C~- ~?~r 400 450 500 550 GM

CORRECTVED Al RFLOW W 2,1 St 2 LB I E(FULL SCALE)


100

> II

L0Z . .n .... C T A. .

Cr

uL

94 --- --

200 250 300 350 400 4 50 500 550 600

CORRECTED AI RFLOW Wa 'rt2/ t2 " LB / SEC(FULL SCALE)

FIGURE 16 CONFIGURATION 2 PRESSURE RECOVERY VS AIRFLOW (Static and Cross-Wind Conditions)

-F

100

w>0U

-J

200 250 300 350 400 450 500 550 M

CORREC TED Al RFLOW Wa t 2/ t2 - LB / SEC(FULL SCALE)


, ,: /I:! i~~~t -W 4 ,t .. ..A- tSt ti nA' i

•-: ;. :::: .. .. ii I '::::

.4-.9-

o. F36.- 3 3

-t o.96 t I j j

......- : .. . . :- 1 :i: :" t : ,

-r, iti: :- I ~-..ii. .1. • : , " :- : ' '' e. '.I 35.2 ---- . .

C. 36. 3 " I 36.

!::RADIUS

336.

LI

ft . Kt i

16~t 17 _~

4 i 7 7'' ! AI L. . vi bi- --.I.' .. .: " ' I RUN I K l bsec

,* T I .;: ! '3 6 .3 . - - . V136.4 0'....0, 81. Cru {

i: I:: : iI ::CL i -7f"]- "t-- + ....... 3.5. , -502 - "

"~:. I:: ': : Lt

Sn; D . . 35.4. i7.. 808

.7 -t SetonCC4Ll olIi.

36.36.. ... 3 6 . ... .......

94 -;Jl 7

I- -

I

go 5 62 ; 64 I' .

u i 77a e

60 : 9 298 300 i 302 304 ,

_;.1. 1111. t' , 6

.ANGULAR POSITION DEGREES

: .... .. '; ; "i L-: :' : ''': '" " ' '" + i .. i' :ii :1 -

: ,RADIUS 20 INCHES FULL SCALEJ:t "

::1 jii i j i~jjf~f!jfIIi l~i; I- - I I A°i[.7+, : : .., 8 . .' " ., . . r -- r . i. . F[ :!m !m N i .''4 ' i" r . . !- !:i . : . .. . : .

FIGURE 18 CONFIGURATION 2 ENGINE FACE RING AND STRUT

TOTAL PRESSURE TRAVERSE

45

100

wt

U-- - -- --- _ _ _ ___._ _ _ _ _ __-'_-.' :

wf

-- ----

--+-- ---_El _ --m-lEE -E

0 ___

. . , . ... ... b- -- -- - - - , -, -- - -. . . . . .. .... ... . ... .......

LI 97

Lli

----- - - ----- -- -

-:~~~_ , _1__ Ti ..- '.-- -- ------

I-~

200 250 300 350 400 450 500 550 600

CORREC TED Al RFLOW Wa~/~2 LB/ISEC(FULL SCALE)

FIGURE 19 CONFIGURATION 1L PRESSURE RECOVERY VS AIRFLOW (Static & Cross-Wind Conditiions)

.. . I ,.. - " ! -' I l.. . :-I-- l ' ... . : ... .... : - = r - -- - --+d -- A=---- : - ---_:: , . == - = -i: ====-=-==: _ ,... . +: ii . -- i t L C' - i- .. . ,- -'i -', i:::A + ; -----rT

• :, !,, ! ,, " ' - - - i: i ,i '. -.- '- " , i =:= - ---. ... .... ..... .h ....

CORE TED AI RFO Wa"t, t - LB / SEC94 --(FULL SCALE)FIGRE19COFIURTIN L RESUE ECVEY S IRLOW- (Satc& rssWndCndtios

100

@

0 W~t 9

1

I, == -

..... . .. . . .. . .

N

(FLLSCLE

L

w05

FIGUE 20CONFIGURATION 2R PRESSURE RECOVERY VS AIRFLOW (Static, Angle-of-AttackadCross-Wind Conditions)

-'ai

(FLLSCLE

V) UE2 OFGRTO 2PRSUERCVR SARFOW(ttc nl-fAtc

Ln rs-id odtos

Yt7 REFAN, SIDE INLET WOCEL TEST

CONFIG. 1, ALPHA s17.50EC. FwD. VELOCITY

TUNNEL VELOCITY z= 18 KNOTS

P IP

T .PTOD .I .1 6 .38 .990 7 .900

3 .0900 8 .860

4 .90O 9 .040

9 .940 0 .800

TEST NO. 2390 TEST CATE 12/12/73 CALC. CATE 01/16/74

RUN NO. 46 RECOVERY .9944 PRI RECOVERY .9995

COND. NO. 4.0000 wa/u2/6t2 468.904 LB/SEC FAN RECOVERY .9919

FIGURE 21 CONFIGURATION 1 ENGINE FACE PRESSURERECOVERY MAP (180 Knots @ 5 = 17.50)

48

7IZ REFAN, SIDE INLET MODEL TEST

VTs tO KTS , CONFIGURATION NO. I , 90 DEG. X-WIND

/ P .9= 0 P / PT

S .99S 6 .920

a .09o 7 .900

3 .90 8 .680

4 .940 9 .640

S .940 .600

TunnelFlow

TEST NO. 2390 TEST DATE 11/15/73 CALC. DATE 12/06/73


COD. _NO. 5.0000 Wa V' 2/6t2 485.492 LB/SEC FAN RECOVERY .9623

FIGURE 22 CONFIGURATION 1 ENGINE FACE PRESSURERECOVERY MAP (20 Knots Cross Wind)

49

711 REAN, SIDE INLET MODEL TEST

COrts. 2, ALPHA 217.50DE. FrD. VELOCITY

TUNNEL VELOCITY' 180 KNOTS

T oo I T180

N NO. R .0$ 9 2

8 90 ? 900

a 900 * M0

4 9OO 9 e840

0 0940 0 .00

SEE FIGURE 18 FOR WAKE DEFINITION

POO

TEST NO. 23L90 TEST DATE 12/ 5/?3 CALC. DATE 01/09/74


COND. NO. 3.0000 Wa!t-216t2 466.275 LB/SEC FAN RECOVERY .9879

FIGURE 23 CONFIGURATION 2 ENGINE FACE PRESSURERECOVERY MAP (180 Knots @ = 17.50)

50

?zr REFAN. SIDE INLET MODEL TEST

VT a 20 KNOTSi CONFIG. 2 , 90 DEG. CROSSWINC

P IPT e0 PT /PTT ToD + T TDa

i . .92

a .06.t .' .

.90

4 .960 9 .840

S .4 .800

TunnelFlow


RUN NO. 1s RECOVERY .9690 PRI RECOVERY .999?

COND. NO. 3.0000 waFt2 /6t 2 472.283 LB/SEC FAN RECOVERY .9701

FIGURE 24 CONFIGURATION 2 ENGINE FACE PRESSURERECOVERY MAP (20 Knots Cross-Wind)

51

1?7 REFAN, SIDE INLET MODEL TEST

VT a 30 KNOTS, CONFIGURATION ILe 30 PERCENT LIP, 90 DEG. X-WIND

p 8:0 PT IPToT IPT ----- +

3 ,II . .90

3 .e0 a .660

4 .60 9 .840S .940 0 .600

TunnelFlow



COND. NO. 3.0000 w~ 1i26t2 480.784 LB/SEC FAN RECOVERY .9878

FIGURE 25 CONFIGURATION 1L ENGINE FACE PRESSURERECOVERY MAP (30 Knots Cross-wind)

52

727 JT8D-100 SIDE INLET

tiI t I i

0.14 I.

0.12

O 00 0

0,10 H I

T I fill

0

2 4 5 6 a

RADIUS INCHES(MODEL SCALE)FIGURE 26 CONFIGURATION 1 CIRCUMFERENTIAL PRESSURE

DISTORTION 600 SECTOR (Static, Fwd

Speed and Angle-of-Attack Conditions)

53

727 JTBD-100 SIDE INLET

titlii-

I I iflit

0.14 iit: .

ItI

0.12

i1 i ,

i! Illi lil i! i ii iI! Jll !

H 44I i i

i P- •- i { 1

RADUS '-" ICHE (MDE SCALE)iii iit ''

and Agle-f-Atack ondiions

5; o

C602l lill iii 11 IT: rlii "i T

1; ii~ii i iiif r!'

It I Ill' i

liti1 iI ;i ::I ii0 if '

(32 3 4 5 ii

RADIS ICHES(MOEL SALEFIUR 2 CNIGRAIO CRCMFRETALPRSSR

DISTORTION 180 SECTOR (Staic, Fwd Speeand Anle-ofAttac Condtions

54 i i!: I i

727 JT8D 100 SIDE INLET

0.14

0.12

0

RAUIUS.- IH MO LS E

00

R S ICE (

FIGURE 28 CONFIGURATION 1 RADIAL PRESSURE DISTORTION

0

1727 JTBD-100 SIDE INLET.. .........W ill I I I;i

til tii l -

i i il lii ili iii i -

H:it

i:! ii i!;.LU. ,;;..

O,1 i ii lili -i ill-iii i

IINI

'o i i i0.12 ij ~iiii i

-4 1 .iii~i~i,, i ,I ii i ."; i .

z .... ... .

RADIUS - INCHES(MODEL SCALE)FIGURE 29 CONFIGURATIONI:: i li CICMFRNTA PRESSUREt

DISTORTION ~~i 600 SETO (osWn Codtins)

50010tt I6

11 q I I I I !hE 11.. ..

H :: jl

I i I

i it

9 3 4 5 GRAIU ICES(MDL CAE

FIUE 9COFGUAIO CRUMEENIL RSSR

DITOTIN 00SCTR CrssWndCodiios

56 ' 'j


It AL .I . .

...... ......... T1.11

WW I i il

q it -ft4+ :H

104104 I tKbi'H t fI H 111M 'i if I i 11 I0.14

..... I l

0.10f +1 . i i..... ~ ~ ~ ... .. .. .....

. . . . .1 i 11" I

D0 .0

IX-:

0i'

a.- 0,06 ..... .... ..

A .. .. .... N.......... ...... ... f

' ll'' i 111 ii 'i1 tlll il-- il if

" i

0 ,0 4.... ... ...

FIGURE 30 CONFIGURATION 1 CIRCUMFE ENTIAL PRESSUREDISTORTION 1800 SECTOR (Cross-Wind Conditions)

:57

J.iSI . . . . . . . . . fif

01 !

off if

2 3 4 5 6 7i'RAIU INCHES (MODEL SCALE) ;

F.IGURE 30 CONFIGURATION 1 ESRDISORION180 SCTR Cos -WndConi i o s

c I! iii iii 57

727 JT8D 100 SIDE INLET

014

0.12

09 0.10

a a

o

n 3 4 5 8 7

RADIUS~ INCHES(MODEL SCALE)FIGURE 31 CONFIGURATION 1 RADIAL PRESSURE

DISTORTION (Cross-Wind Conditions)

58


0.14

0.12

0.12

I

2 A

RADIUS INCHES(MODEL SCALE)

if

and Angle-of-Attack Conditions)

1I

0.04 i67

1727 JTBD-1O0 SIDE INLET|

i i '

t:i I! ::::o., ::iii ii i i i i1• i: T .6 iis

IT I ... :...

iii.; .1tij ! t

0.10 iIii i!;

....

i !:,;4: 4i II

11.1 L H. ifi 11 j~ ii ii

' " 0i lil' ii;

e)~~~., .h 1. .. .... .i... . ... .

RADUS INHE(MDE SCALE)i iii ii ii

iii i i:i ti: i

ifi

:i:I I i t

DIST TO 180 "0CTO (Sai, +w Seed

060

0 7;

t If I it it1! it I

2 3 4 5 G 7::

and Angle :of Attai k Conditions)

60;i ::I::i::

-727 JT8D 100 SIDE INLETJ

0.14

0.12

o. Hio il 1 V [ I i

106o0

0.04

0.02

RADIUS--INCHES (MODEL SCALE)FIGURE 34 CONFIGURATION 2 RADIAL PRESSURE DISTORTION(Static, Fwd Speed and Angle-of-Attack Conditions)

61

LLo

a ~~~I~ III~~~i~JMIW. -II~IIrK /[

61

7Z27 J7GDCO SSF O~ :I

N EE1j7t...............

?l{.i 1t .I Y screens

i:i

I~ Illt;i:1 i i ;i i tII1 i LI I ill

.14 4.j Th1 4

0A440 lfii- I=it i1o iii

0.0I I I t ' i l ,-ii~ iIiii ' '; I 1iiiifl i

hd .. .

4 4 Jfi~ Iitt!

1 !d I0

i t

1t

I IT ; "

2 I

7 iI

FIGURE 35 CONFIGURATO r tRCUMFIRENTIIL H,#RI

Wind Conditios

62

I 1 H 1 - ll -"! li

0 I I ,::, li t:~

,1T1

FIGURE 35 CONFIGRATION 2 CIRCUMFRENTIAL PRESSUR

Wid onitios

62~ij 1i E Iii1 . ~ i: iii ' i i :Ij ii

727 JT8D-100 SIDE NLET

screens

0.14

0.12

si o

0

I-

2 35 7

RADIUS 0 INCHES(MODEL SCALE)

FIGURE 36 CONFIGURATION 2 CIRCUMFERENTIAL PRESSUREDISTORTXON 1800 SECTOR (Static and Cross-Wind Conditions)

63

FIUR 3 ONIGRAIN 0CICMFRETILPRSSRDIIOTXN10SETR(Sai ndCos-idCndtos

i6

727 JT8D 100 SIDE INLET!

screens

it

0.12

Sii

. ..T 1 1

IilOl ift! i

0.02

RADIUS- INCHES (MODEL SCALE)

FIGURE 37 CONFIGURATION 2 RADIAL PRESSURE DISTORTION

(Static and Cross-Wind Conditions)

64


0.14 ..

0.12 T1hI II

0610 4j

,I- +

0 1 ..I I

RADIUS INCHES(MODEL SCALE)

DISTORTION 600 SECTOR (Static and Cross-ind Conditions)

65

[I

Q 04


.. .... ....

0.14

iti0.10

. l i

0602 I

FIGURE 39 CONFIGURATION IL CIRCUMFERENTIAL PRESSUREDISTORTION 1800 SECTOR (Static and Cross-Wind Conditions)

OC1 Ii :xl i ii l tlI ffiI 11Illiiiii Iii l i

RADIUS~ ~~ i i INHE (MDE CAFIUR 39 COFGRTO LCRUFRN IAL RESUR

DITRIN18 o SCO (Sai ands Cos-inCodtns

6Gllii

727 JTBD 100 SIDE INLET

I 2:2If 4 i

fill ~. ... . .

0.14... .i t $k li 6 it l o :pr rt i iii i I llli: ::

... ... ... s : y .2 i :: ... .. ..: .: : .: .: ........ ......

, !...1.. 1..... .... ... :

.............0.10 .. ..

Q !T 7: . .I... . ..

... ... .........

... .... ...i ii i lil l il~i i . . .... ... .... .. . ...

0.04

7i :::- 7 -r- 't<:

i

FIGURE ~ ~ ~ ~: 40 CNIUATO&LRDAL PRESUR DISORIO

rx :4t: )A :i~ ::':: ' '

w e ... .... ... 4 : i ::i

....:i:. .. ...

::L .. :: .... .. I ... ... : I:

: : : : :: >: :::: ::0. 2... .... ..

(St ti and. Cros W..i nd Cond.. ... i ) 0.047

777 . .. .... .. ..... ... ...: : ::I:: :.: I:I: ::

.......... ....r i0.02 f7ii i IIliII II

.... .. .... . :O ~i 0 :: ::: "" .;i.... ... .... i:.

. .. . . . . . .. . .

4 3 4 5 7: '::C3 1i : RADIUS- INCHES (MODEL SCALE):I:I: ::

FIGUE 40 CONFGURAION L RDIALPRESURE ISTOTIOaoe ~i(Static and Cross-Wind Condit'i1.:....,..;::ons) :

I:: :,:::i67

727-200 PRODUCTION(TRUE VIEW)

727 JT8D-100CONFIGURATION I

* The production inlet is identical 727-200 727-200 *on all JT8D series engines. JT8D-I00 (PRODUCTION) JT8D-15,a.a cruise corrected airflow, a 2/I t 2 - Ib/sec 480 326

M.=:0.8, 30,000 feet, std. day

Highlight diameter - inches 52.21 42.305

Highlight area - sq. in. 2141.07 1405.64

Lip loading, WaV/t 2 A HI - lb/sec-ft 2 32.28 33.40

Highlight Mach number 0.420 0.438

Throat diameter - inches 46.70 37.84

Throat area - sq. in. 1712.87 1124.59Throat Mach number 0.572 0.604Contraction ratio (highlight area / throat area) 1.25 1.25

Lip contour 2.5:1 super ellipse 2.111:1 ellipse

Distance between inlet e and fuseolage - inches 44.5 (approximate) 39.01

Engine face diameter - inches 50.10 40.50Nose dome diameter - inches 16.00 12.030

Engine face area ~ sq. in. 1770.29 1174.59

Engine face area / throat area 1.034 1.044Engine face Mach number 0.543 0.563Nose dome length - inches 25.20 15.038.Degrees of cant 0.0 4.0

Length / engine face diameter 0.8 1.0

FIGURE 41 DESIGN COMPARISON OF 727 PRODUCTION AND JT8D-100 SIDE INLETS

68

I -. .- . ........

727-200 P DUCTION DDIFIED :- -: 727 JTD- 100 SIDE INLET '.. SIE INLET .0.37 SCALE: CONFIGURATION 0.30 SCALE :iii. . . . .. :..... 17 . .: + .

" 0Z V-KNOTS SYM V KNOTS : .

p+: .. . . - . . ' -7 = 7 -. 77 ...... F I-o aE 4 0 _ou...... 17 18 0 7 .5 - 18 0" Configuration 2

-.1 7.5 180_ _+_ -- +_..0....urtion 2.... + '-

.. est prob (full i i.. -.

7-f-L - from the cowl wall-all data. -. ii ::i :-.S- 'i Al data without acoustic treatment :

.4.ips

L-.-9 6-

:+ ... ..

01) 350 00 0 0

I t I 6 t2

... 4 .CORR CTED AIRFLOW, L B/SEB . .

(FULL SCALE)

FIGURE 42 PRESSURE RECOVERY COMPARISON 727-200 PRODUCTION (Modified) AND JT8D-100SIDE INLETS

: - --

-..4 . 727-200 PRODUCTION (MODIFIED) 727-200 PRODUCTION 727 JT8D-100 SIDE INLET -SIDE INLET 0.37 SCALE SIDE INLET- FULL SCALE! CONFIGURATION I - 0.30 SCALE

S SYM IV KNOTS SYM ALTITUDE MACH NO. SYM -KNOTS: - h i

S o 180 9o b.3o 0 o 180

.. C> 10,250 0.34

- - ± L : --. 9,900 -- 0.21

C . 21,300 - 0.37 i : : ::I::::

. . . -- . . 20,200 0.41I 31,000 0.47

SA 30,700 0.851

S 10o : - -- - O 30,800 0.88 ---.. ..

. .. . . L 41,600 0.65

S- 9 .i 41000 0.80 8. .-- - _ _ e- .. .......7 .. ..... .. . -r .Note: ?2 -

8 .*Closest probe inch (full scale)from the cowl wall - all data

t - .Model data without acoustic treatmen i7

- * Full scale data with 12 inchesL ... ..of treatment ....

0 :.

--

7 7250 300 350 ... 50 500

2 ----- • ---

v CORRECTED AI :-2.t2 . LB/SEC:..-.. . .- --, .- . 3. . . . .

FULL SCALE)

S- T200TA i 250 RE.300 DISTORTI35ON COMPARISON 727-2050 00 PRODUCTION AND JT 100 SIDE INLETS

--.-CORRECTED AIRFLOW, W 82 /6t2 LB/SEC .. : - - .-- ..... . . . ------ , : I

FIGURE 43 TOTAL PRESSURE DISTORTION COMPARISON 727-200 PRODUCTION AND JT8D-100 SIDE INLETS

APPENDIX A

SYMBOLS

71

APPENDIX A

SYMBOLS

a Half the major axis of an ellipse

AHI Highlight area (or Hilite)

ATH Throat area

BBL Body Buttock Line

BWL Body Water Line

b Half the minor axis of an ellipse

D2 Engine face diameter

Kt Knots

L Inlet length

L.E. Leading edge

MCR Maximum Cruise

MCT Maximum Continuous Thrust.

Moo Free stream Mach number

MTH Throat Mach number (one dimensional)

M2 Engine face Mach number (one dimensional)

PToo Freestream (tunnel) total pressure

PT Engine face total pressure measurement

PT Min Minimum engine face total pressure measured

PT Max Maximum engine face total pressure measured

P Area average (see Section 3.4) of engine faceT2 total pressure measurements

PT2/PToo Total pressure recovery

PT Ring Avg, Average engine face total pressure at a

PT Local Ring given radius

Avg 73


Max Ring Avg The maximum average engine face total pressureT Max Ring vg computed at the given radii (Maximum PT Ring Avg )

P Average engine face total pressure at a givenT Min Sector Avg radius in the minimum total pressure sector

(for 60 or 180 degree sectors)

R Inlet Radius

TT2 Total temperature at engine face (average)

Vt Tunnel velocity

V, Freestream velocity

T.E. Trailing edge

Wa Actual airflow

Wa 2/ 6t 2 Engine face corrected airflow

X Inlet centerline distance, or coordinate inellipse equation

Y Coordinate in ellipse equation

Inlet angle-of-attack (inflow angle) relativeto engine and inlet

SInlet yaw angle relative to inlet

0 Angular position around engine face

9 t2 TT2 / 5 18 . 7 when TT2 is in degrees Rankine

6 t2 PT2/ 14 .69 when PT2 is in pounds per square inch

74

APPENDIX B

INLET COORDINATES

75

APPENDIX B-INLET COORDINATES

ALL MODELS AXIALLY SYMMETRICALL DIMENSIONS INCHES FULL SCALE

COWL CONFIGURATIONS 1,2, and 2R NOSE DOME

(CONFIGURATION IL from X - 33.2 to-94 C)NFIGURATIONS 1.2. and IL

X4 R COMMENT x R COMMENT x R COMMENT

6.631 31. 000 37.200 23.766 25.200 0.00 - Nose

14. 092 30. 945 Contour not 36. 700 23. 652 (The lip contour 25. 000 1. 006

17.784 30. 658 critical in 35.700 23. 489 t .etremely 24.000 2.439

24. 476 30.123 thi region 34. 700 23. 394 critical) 23. 000 3.269

27.822 29.745 33.700 23.354 22.000 3.902

31.168 29. 272 33. 200 23. 350 - Throat 20. 000 4.867 (3.15:1 Ellipse)

33. 398 28. 890 33. 000 Z3. 3505 18. 000 5. 599

35. 629 28.421 Estereal 32. 000 23. 364 16. 000 6.181

37.302 27.962 Co l 31.000 23.398 12.000 7.035

28.417 27.553 30.000 23.449 8.000 7.586

39.254 27.137 28. 000 23.600 4.000 7.89839.S2 2.95 NASA fefan 9

39. 532 26. 90 Nacelle Cro 26. 000 23. 806 2.000 7.975

39. 755 26. 76 Contour - Ftl 360J 24.000 24.046 0. 00 8. 000 . Face, Slope = 0

39. 867 26. 631 22. 000 24. 295 Internal 0. to -9 8.000

39.9785 26.469 20. 000 24. 532 Cowl

40.034 26.3579 16.000 24.922

40.090 26.1061 - Hilite 14.000 25.059

40.000 25.558 12.000 25.150

39.900 25.340 10.000 25.190

39.800 25.182 - 8.000 25.17939. 600 24. 943 Lp - Super Ellip'e 6.000 25.13739.400 24.758 2.2 2.2 i4.000 25. 090

39.200 24.606 ) +/'\ =11 3.000 25.07238.700 24.308 l \b / 2.000 25.060

38. 200 24.084 1.000 25.052

37.700 23.907 0.000 25. 050- Eng. Face, slope = 0

o... , 2,.o5o RING-CONFIGURATION 2

NOSE

x R Slope

27. 179 14. 835 (±) 50. 10 L. E. Radiue . 090"' 1

27.200 14.770 With Center X 27. 11

STRUTS -CONFIGURATION 2 27.168 14.709 (39.9 14.

Strut x Inside Radius Outside Radius

/c y/ _ ,/ /c (NACA 0009) 7. 179 - 14.835

.00 .00 .300 .04501 Airfoil 27. 168 14.709 -

.0125 .0142 .400 .04352 27. 050o 14.64 1.937

.025 .019642 .00 .04352 7 26.800 14.579 15.044 Modified NACA

.025 .01961 .500 .03971 C 14.547 15.120

.050 .02666 .600 .03423 26.500 14.547 15.0 632 AO 526. 150 14. 536 15. I1O

.075 .03150 .700 .02748 25.800 14. 547 15. 243

.100 .03512 .800 .01967 Rig 25.500 14.569 15.273

.150 .04009 .900 .01086 25.000 14.614 15.318

.200 .04303 .950 .00605 24.500 14.658 15.362

.250 .04456 L00 .00095 24.000 14.703 15.407

Note: C = 10.3 inches, perpendicular to leading edge 23.000 14.792 15.496 tr

L. E. Radius = . 0089XC 22. 000 14. 881 15.585

Strutes 120° intemle 18. 000 15.237 15. 941

16.000 15.415 16.119

14.000 15.593 16. 29

12.000 15.738 16.439

10.002 15.858 16.559

8.000 15. 94 16. 646

6.000 16.025 16.700

5.000 16.076 16.696

4.000 16.140 16. 670 Modified NACA

3.000 16.222 16. 622 0009 Airfoil

2.500 16. 271 16. 589

2.050 16.319 16.554

1.500 16. 385 16. 504 -1.000 16.4507 16.4507

Preceding page blank i77

RING-CONFIGURATION 2R

Nose

i Slope

31.176 14.691 () 47.75 0 E. radis- .1090-31. 200 14.626 0 a31. 171 14.561 () 42. 25o

= 14. 630

X Inside I Outside R

31. 176 - 14.691

31.171 14.561 STRTS-CONFIGURATION 2R

31.050 14.492 14.783 STRUTS-CONFIGURATION 230.800 14.416 14.8 I30.500 14.377 14.941 i.a Strut

30. 150 14.356 14.995 63 2A015 a/c y/c x/c y/c NACA 0009

29.800 14.350 15.032 Airfoil Airfoil

29.500 14.355 15.03729.500 14.35578 15.037 .0 0.0 .30 .03001S. 12.0125 .00947 .40 .02902 .

28.500 14.413 15. 125..025 .01307 .50 .02647

28. 002 14.454 15. 151 .025 .01307 ;s0 .026472.00 Z 14.537 1.230 .050 .01777 .60 .02282

27.000 14.537 16.230

26.000 14.620 15.312 .0 ..0210 .0 .01832Rin

24.000 14.792 15.496 .100 .02341 .0 .0131222.000 14.984 15.691 .150 .02673 .90 .0074

20.000 15. 182 15. 886 .200 .02869 .95 .00403

18.000 15.368 16.076 .250 .02971 1.00 .0063

16. 000 15. 540 16.254 Note: C = 12.6", Perpendicular to L. E.

14.000 IS.706 16. 395 L. E. Radius = .0040 . C

12. 000 15. 895 16.42 Modified NACA 0009 Struts 6 1200 intervals

11.000 16.016 16.420 Airfoil

10.500 16.083 16.408

10.002 16.153 16.386

9.500 !6.228 16.352

9.000 16.306 16.30J

LIP AND EXTERNAL COWLCONFIGURATION IL

NOSE DOME Comment

CONFIGURATION 2R S.385 3 1.5119.076 31. 175 Not Critical

XI R Comment 25. 768 30. 640 in This Relton

29.200 0.0 - Nose 29.114 30.262

29.000 .935 32.460 29.789

28.000 2.268 34.691 29.407

27.000 3.046 36.921 28.939

26.000 3.641 38.594 28.479

24.000 4.557 3.65:1 39.710 28.070 External Cowl22.000 5.260 Ellipse 40.546 27.654

20.000 5.829 40. 825 27.467

16.000 6.692 41.048 27.273

12.000 7.293 41.159 27. 149

8.000 7.694 41.271 26.986

4.000 7.925 41.3267 . 26.875

Z.000 7.981 41.3827 26.6231 Hilite

0.000 8.000 --Eng. ace, Slope 0 41.300 26.044

0. to -9 8.000 41.000 25.47440.800 25.241 (Lip Super Ellipee40.500 24.971 2.2 2.240.000 24.633 - .39.400 24.330 -. b) =38.700 24.06238.200 23.91037.700 23.78336.700 23.59035.700 23.46234.700 23.38633.200 23.350 - Throat

78

REFERENCES

1. Easterbrook, W. G., Roberts, W. H., The Boeing CommercialAirplane Company, "Low-Speed Wind Tunnel Flow Field Resultsfor JT8D Engine on the Boeing 727-200", NASA CR Document,to be released.

2. Easterbrook, W. G., Carlson, R. B., The Boeing CommercialAirplane Company, "Cruise Drag Results from High SpeedWind Tunnel tests of NASA Refan JT8D Nacelles on The Boeing727-200", NASA CR-134546, February 1974.

3. Bailey, R. W., Vadset, H. J., The Boeing Commercial AirplaneCompany," 727/JT8D Refan Side Nacelle Air Loads", NASACR-134547, March 1974.

4. Pratt & Whitney Aircraft, "Phase I Engine Definition andCharacteristics of the JT8D-100 Turbofan Engine",P&WA TM-4713, April 13, 1973.

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