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UNANNOUNCED "A.mL-TR-s"7
AD-A956 237
THE EFFFC OF RUNOL NUBER VARIATON OFFLAIL INITIATING FORCES ACTING ON A CREWMANDURLNG EMERGENCY ESCAPE
DTIC&ockheed Missiles & Space Company, Inc. ~THwumsille Research & Engineering Center S APR 2 11993 34"/ObrzndfordDrive, Hunsille, AL35807
Approvedforpublic release; distribution unlimited.
AIR FORCE AEROSPACE MEDICAL RESEARCH LABORATORYAEROSPACE MEDICAL DIVISIONAIR FORCE SYSTEMS COMMANDWRIGHT-PATTERSON AIR FORCE BASE, OHIO 93-08239. ,- q 4. I~l!lllll~lllli93 l-08239!ltil! ~
41
NOTICES
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TECHNICAL REVIEW AND APPROVAL
AFAMRL-TR-84-067
The voluntary informed consent of the subjects used in this research was obtained as required by AirForce Regulation 169-3.
This report has been reviewed by the Office of Public Affairs (PA) and is releasable to the Nz,'io. .1Technical Information Service (NTIS). At NTIS, it will be available to the general public, inclo(&ingforeign nations.
This technical report has been reviewed and is approved for publication.
FOR THE COMMANDER
I'41HENNING I,. VON GIERKE, Dr IngDirectorBiodynamics and Bioengineering DivisionAir Force Aerospace Medical Research Laboratory
UNCLASSIFIEDSECUAITY CLASSIFICATION OF THIS PAGE
REPORT DOCUMENTATION PAGEis. REPORT SECURITY CLASSIFICATION lb. RESTRICTIVE MARKINGS
UNCLASSIFIED2.L SECURITY CLASSIFICATION AUTHORITY 3. OISTRIBUTIONIAVAILABILITY OF REPORT
1. OECLASSIFICATIONIDOWNGRADINGSCHEDU-E Approved for public release; distributionunlimited.
4. PERFORMING ORGANIZATION REPORT NUMBERIS) 5. MONITORING ORGANIZATION REPORT NUMBER(S)
LMSC-HREC TR D867584 AFAMRL-TR-84-0676& NAME OF PERFORMING ORGANIZATION OFFICE SYMBOL 7.L NAME OF MONITORING ORGANIZATION
Lockheed Missiles & Space I.pp i Air Force Aerospace Medical Research LabCompany, Inc. Aerospace Medical Division, AFSC
fw. ADDRESS (City. State and 7P Code ) - 7lb. ADDRESS (City. State and ZIP Code)
Huntsville Research & Engineering Center4800 Bradford Drive Wright-Patterson Air Force Base, OH 45433-Huntsville AL 35807 6573
B.. NAME OF FUNDING/SPONSORING jib. OFFICE SYMBOL 9. PROCUREMENT INSTRUMENT IDENTIFICATION NUMBERORGANIZATION (i[ pplicbk)
____________ F3361 5-80-C-051 58c. ADDRESS (City. State and ZIP Code) 10. SOURCE OF FUNDING NOS.
PROGRAM PROJECT TASK WORK UNITELEMENT NO. NO. NO. NO.
11. TITLE 'Include Security Clasification) 62202F 7231 13 09SEE REVERSE
12. PERSONAL AUTHOR(S)
William F. Braddock13a. TYPE OF REPORT 13b. TIME COVERED 14. DATE OF REPORT (Yr.. Mo.. Day) 15. PAGE COUNT
Final-Report FROM 80 Jul TO 81 Dec August 1982 10016. SUPPLEMENTARY NOTATION
17. COSATI CODES 1B. SUBJECT ThRMS (Continue on reverse if necessary and identify by block number)
FIELD GROUP SUB. GR. Ejection Seat (ACES II) Flailing Limbs607 0Ni1 06I/ Aerodynamic Characteristics F-16 Aircraft
Crewman Limb Forces/Moments Crewman (Dummy) (con't)19. ABSTRACT (Continue on reverse if neceuary and identify by block number)
A 0.5-scale crewman/ejection seat model was tested to determine the effects of Reynoldsnumber variation on the aerodynamic forces acting on the crewman in and near the cockpitof an F-16 fighter during an emergency ejection from the aircraft. Two different cockpitconfigurations were investigated at freestream Mach numbers from 0.4 to 1.2 and dynamicpressures from 100 to 600 psf. The data were obtaii..d at various F-16 forebody-ejectionseat separation distances (0 to 24 in. model scale) for fuselage angles of attack andsideslip of zero.
20. DISTRI BUTION/AVAILABILITY OF ABSTRACT 21. ABSTRACT SECURITY CLASSIFICATION
UNCLASSIFIEO/UNLIMITED 0 SAME AS RPT. 0 DTIC USERS 0 UNCLASSIFIED22a. NAME OF RESPONSIBLE INDIVIDUAL 22b. TELEPHONE NUMBER 22c. OFFICE SYMBOL
(Include Area Code)LAWRENCE J. SPECKER 513 255-3122 AFAMRL/BBP
DD FORM 1473.83 APR EDITION OF 1 JAN 73 IS OBSOLETE. UNCL
UNCLASSIFIED
SECURITY CLASSIFICATION OF THIS PAGE
Block 11:THE EFFECTS OF REYNOLD,: NUMBER VARIATION OF FLAIL INITIATING FORCES ACTING ON ACREWMAN DURING EMERGENCY ESCAPE
Block 18:Ejection Sequence (Simulated)Reynolds Number Effects
PEEFACE
The wind tunnel test described herein was conducted by the
Lockheed-Huntsville Research & Engineering Center for the Air
Force Aerospace Medical Research Laboratory (AFAMNL) under Con-
tract F33615-80-C-0515. The Air Force Project Engineer for this
contract is Lawrence J. Specker, Biomechanical Protection Branch
of the Biodynamics and Bioengineering Division. The test was
conducted in support of workunit 7231-13-09, t"Definition of
Windblast Flow Fields."
ooss1on For
DTIC TABUnannounoed 0
Justification
By , --- --- --y- Distribution/
Availability Codes
Avail and/or
Dist Special
1 LA mIum imm
TABLE OF CONTENTS
Section Page
PREFACE .................................... 1
NOMENCLATURE ...................... ...... 4
1 INTRODUCTION ANDSUMMARY . ...........
2 TEST OBJECTIVE ........ .. . . . 10
3 MODEL DESCRIPTION . . .i. . . 11
3.1 Model .................... . . . 11
3.2 Related Hardware o .......... 16
4 INSTRUMENTATION . ..................... 26
4.1 Limb and Neck Gages ...... .................. 26
4.2 Crewman/Seat Total Forces and Moments .. ....... .26
4.3 Static Pressure Measurements .. . . 26
4.4 Flowfield Visualization .... .... ... . 27
4.5 Hydraulic Pressure .. o o............. 27
4.6 Position and Attitude Indicators ..... . 27
4.7 Calibration .. . . . . . . . . . . 27
5 TEST PROGRAM . . .. . 305.1 Configurations Tested . . . 30
5.2 Test Parameters . .. .. .. .. . . . . 30
5.3 Run Schedule . o o . o . . ...... * . . . a . . 32
6 DATA'REDUCTION AND PRESENTATION . o . o o o o . o o 33
6.1 Limb and Neck . o . . . . . .. . .. .. .. . . 33
6.2 Crewman/Seat . . . . . . . . . . . . . . . . . 40
6.3 Crewman Pressures ..... . . . .. . 40
6.4 Forebody . . . . . . . . . . . . . . . . . . . . . . . . 43
2
TABLE OF CONTENTS (Concluded)
Section Page
6 6.5 Data Tabulations ...... ..................... 43
6.6 Data Plots ......... ...................... ... 43
7 RESULTS ........... ........................... 44
8 CONCLUSIONS ............. ................. ....... .72
Appendix: Tabulated Test Data ...... .............. ... 73
REFERENCES ........... .......................... ... 99
3
NOMENC ATURE
ALPHC crewman angle of attack (positive nose up), deg
B.L. abbreviation for boundary layer
CA axial-force coefficient, body axis, (positivedownstream), FA/QS
CD drag coefficient
CFEYL,R crewman lower arm side force area parameter leftand right arms, (positive to crewman's right forboth arms), FEYL/Q, FEYR/Q
CFEZL,R crewman lower arm lift force area parameter, leftand right arms, (positive up), FEZL/Q, FEZR/Q, ft
2
CFHX crewman helmet drag force area parameter, (positivedownstream), FHX/Q, ft
2
CFHY crewman helmet side force area parameter, (positiveto crewman's right), FHY/Q, ft
2
CFIIYL,R crewman upper leg side force area parameter, leftand right legs, (positive to crewman's right forboth legs), FHYL/Q, FHYR/Q, ft
2
CFHZ crewman helmet lift force area parameter, (positiveup), FHZ/Q, ft
2
CFHZL,R crewman upper leg lift force area parameter, leftand right legs, (positive up), FHZLiQ, FHZR/Q, ft2
CFKXL,R crewman lower leg drag force area parameter, leftand right legs, (positive downstream), FKXL/Q, FKXR/Q,ft2
CFKYL,R crewman lower leg side force area parameter, leftand right legs, (positive to crewman's right forboth legs), FKYL/Q, FKYR/Q, ft
2
CFSXL,R crewman upper arm drag force area parameter, leftand right arms, (positive downstream), FSXL/Q,FSXR/Q, ft 2
4
NOMENCLATURE (Continued)
CFSYL,R crewman upper arm side force area parameter, leftand right arms, (positive to crewman's right forboth arms), FSYL/O, FSYR/Q, ft
2
CG center of gravity
CL lift coefficient
CMEYL,R crewman lower arm side moment volume parameter, leftand right arms, MEYL/12Q, MEYR/12Q, ft
3
CMEZL,R crewman lower arm lift moment volume parameter, left
and right arms, MEZL/12Q, MEZR/12Q, ft3
CMHX crewman head drag moment volume parameter, MHX/12Q, ft3
CMRY crewman head side moment volume parameter, MHY/12Q, ft3
CMHYL,R crewman upper leg side moment volume parameter, leftand right legs, MHYL/12Q, MHYR/12Q, ft
3
CMHZL,R crewman upper leg lift moment volume parameter, left andright legs, MHZL/12Q, MHZR/12Q, ft
3
CMKIL,R crewman lower leg drag moment volume parameter, left andright legs, MKXL/12Q, MKXR/12Q, ft
3
CMKYL,R crewman lower leg side moment volume parameter, left andright legs, MKYL/12Q, MKYR/12Q, ft
3
CML rolling moment coefficient, body axis, (positive clock-wise looking upstream) ML/QSd
CMM pitching moment coefficient, body axis, (positive noseup), MM/QSd
CMN yawing moment coefficient, body axis, (positive noseright), MN/QSd
CMSXL,R crewman upper arm drag moment volume parameter, leftand right arms, MSXL/12Q, MSXR/12Q, ft3
CMSYL,R crewman upper arm side moment volume parameter, left
and right arms, MSYL/12Q, MSYR/12Q, ft3
CN normal force coefficient, body axis, (positive up), FN/QS
CPA,CPSA static pressure coefficient, crewman abdomen, (PSA-P)/Q
5
NOMENCLATURE (Continued)
CPC, CPSC static pressure coefficient, crewman'check, (PSC-P)/Q
CPH, CPSH static pressure coefficient, crewman head, (PSH-P)/Q
CPLL, CPSLL static pressure coefficient, crewman left leg, (PSLL-P)/Q
CPRL, CPSRL static pressure coefficient, crewman right leg, (PSLR-P)/Q
CPSBR, CPSSB static pressure coefficient, seat back reference, (PSSBR-P)/Q
CY side force coefficient, body axis, (positive nose right),FY/QS
D diameter of sphere, in.
D diameter of arm, in.arm
d crewman reference length equivalent to the diameter ofa circle with area equal to S, 17.128 in.
EJPOS crewman/seat ejection position, in.
FA axial force, body axis, (positive downstream), lb
FEYL,R crewman lower arm side force, left and right, (positiveto crewman's right), lb
FEZL,R crewman lower arm lift force, left and right, (positiveup), lb
FFHC forebody hydraulic cylinder force, (positive up), lb
FHX crewman helmet drag force, (positive downstream), lb
FHY crewman helmet side force, (positive to crewman's
right), lb
FHYL,R crewman upper leg side force, left and right, (positiveto crewman's right), lb
FHZ crewman helmet lift force, (positive up), lb
FHZL,R crewman upper leg lift force, left and right, (positive
up), lb
FKXL,R crewman lower leg drag force, left and right, (positivedownstream), lb
6
NOMENCLATURE (Continued)
FKYL,R crewman lower leg side force, left and right, (positive
to crewman's right), lb
FN normal force, body axis, (positive up) lb
FSXL,R crewman upper arm drag force, left and right, (positivedownstream), lb
FSYL,R crewman upper arm side force, left and right, (positiveto crewman's right) lb
FY side force, body axis, (positive nose right), lb
L length from nose of forebody, ft
LT length of turbulent boundary layer, ft
MACH,M freestream Mach number
MEYL,R crewman lower arm side moment, left and right, in-lb
MEZL,R crewman lower arm lift moment, left and right, in-lb
MHX crewman head drag moment, in-lb
MHY crewman head side moment, in-lb
HHYL,R crewman upper leg side moment, left and right, in-lb
MHZL,R crewman upper leg lift moment, left and right, in-lb
MKXL,R crewman lower leg drag moment, left and right, in-lb
MHYL,R crewman lower leg side moment, left and right, in-lb
ML rolling moment, (positive clockwise looking upstream),in-lb
MM pitching moment, (positive nose up), in-lb
MN yawing moment, (positive nose right), in-lb
MRC moment reference center
MSXL,R crewman upper arm drag moment, left and right, in-lb
MSYL,R crewman upper arm side moment, left and right, in-lb
7
NOMENCLATURE (Concluded)
P freestream static pressue, psfa
PSA static pressure, crewman abdomen, psfa
PSC static pressure, crewman chest, .psfa
PSH static pressure, crewman head, psfa
PSLL static pressure, crewman left leg, psfa
PSLR,PSRL static pressure, crewman right leg, psfa
PSSBR static pressure, seat back reference, psfa
PT tunnel freestream total pressure, psfa
Q freestream dynamic pressure, psf
RExlO 6 freestream Reynolds number per foot x 1016
RN/PN data run number and test point
S model reference area equivalent to the projected frontalarea of the seat and man, 1.600 ft
2
SF full-scale model reference area equivalent to the pro-jected frontal area of the seat and man, 6.4 ft
2
TT tunnel total temperature, F
V freestream velocity, ft/sec
6 height of boundary layer or shear layer, in.
0 angular position around sphere at which separationoccurs, deg
angle between velocity vector and axis of symmetry, deg
8
SECTION 1
INTRODUCTION ABID SUMMARY
The Air Force Aerospace Medical Research Laboratory (AFAHRL), Wright-
Patterson Air Force Base, Ohio, sponsored an experimental aerodynamic test to
determine the effects of Reynolds number variation on flail initiating forces
acting on a crewman during emergency escape or after inadvertent canopy loss.
A one-half scale wind tunnel model of an ejecting crewman/seat was tested in
the presence of an F-16 forebody model. During the test, measurements were
made of the forces and moments acting on the crewman's arms, legs and head,
total forces and moments acting on the crewman/seat model, and static pres-
sures at six locations on the crewman/seat model.
Data were taken at Mach numbers from 0.4 through 1.2 and dynamic pres-
sures from 100 psf through 600 psf.
The test was conducted at the Arnold Engineering Development Center
(AEDL) in the 16T Propulsion Wind Tunnel. The project number was P41G-09.
The test was conducted during Septenber 1981. Detailed information on the
pretest planning and additional information on the test can be obtained from
References 1 and 2.
Lockheed concludes that there are Reynolds number related effects on the
crewman/seat aerodynamic coefficients and the crewman limb and head force area
and moment volume parameters. These effects are attributed to changes in the
local flow conditions on the crewman spherical (head/helmet) and cylindrical
(arms, legs) components. The forebody flow field outside the cockpit, the
cavity flow inside the cockpit, and the shear layer flow field have no dis-
cernible change with Reynolds number. In most cases when a Reynolds number
related variation does occur, data magnitudes are higher (more conservative)
at the lower Reynolds numbers. Thus conservative results can be obtained by
testing scale crewman/seat models at lower Reynolds numbers.
9
SECTION 2
TEST OBJECTIVE
The Air Force Aerospace Medical Research Laboratory has used wind tunnels
to conduct research on the prcblem of windblast protection since 1971. This
research has been focused upon the measurement of aerodynamic forces acting on
the human body during and after emergency egress from aircraft using scale
models, anthropometric dummies and volunteer human subjects. The data col-
lected has been used to develop principles and techniques of personnel pro-
tection in addition to the analytical techniques needed to predict the human
response to aerodynamic forces.
With the use of scale models in the wind tunnel came the problems asso-
ciated with attempting to create dynamically similar airflows. Dimensional
analysis has shown that the force coefficient for a body of given orientation
and shape is a function of the Reynolds number and the Mach number provided
that parameters such as surface roughness, stream turbulence and the presence
of other bodies in the vicinity are not neglected. In applying data from
model tests to full scale, these facts must be considered. The objective of
this effort was to investigate the effects of varying Reynolds number on the
aerodynamic force measurements made with the one-half scale wind tunnel models
of the F-16 forebody and crewman/seat. The range of Reynolds numbers for this
test was 1.4 x 106 through 6.6 x 106 based on the reference length of the
crewman.
10
SECTION 3
MODEL DESCRIPTION
3.1 MODEL
Crewman/Seat
The crewman/seat model (Figure 1) is one integral unit. The model seat,
fabricated from aluminum plate, not only simulates the full-scale seat, but
includes the structure for the crewman torso and the attach points for the
neck, arms and legs. The seat attaches to a balance adapter which fits over
a strain gage balance. This balance (discussed in Section 4.2) measured the
crewman/seat total forces and moments. A fiberglass reinforced shell attaches
to the seat framework to simulate the external contours of the torso. Instru-
mented beams attach to the frame through openings in the torso shell. These
beams are the structural components of the head and limbs. The contours of
the head and limbs are constructed of two pieces using fiberglass reinforced
plastic over a hardwood form. The two halves of the head bolt together and
the aft piece bolts directly to the top end of the neck beam.
All four of the limbs (left and right arms and left and right legs) are
constructed in two sections (upper and lower). The upper arms are bolted to
the torso frame at the shoulders. The upper arm exterior contour shells are
fabricated in the left and right halves and are bolted to the beams near the
elbow.
Except for the attach point at the elbow, the upper arm shell is free
of any contact with any other part of the model and is therefore free to flex
under aerodynamic load. The load seen by this shell is transferred to the
free end (elbow) of the upper arm beam and can bo measured using the strain
gages mounted on the beam. The lower arm was designed using the same concept
used in designing the upper arm beams. The fixed end of the lower arm beam
ii
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12
is attached to the upper arm beam with a taper joint at the elbow. Left and
right halves of the lower arm/hand shell are bolted together and attached to
the beam at the wrist. The shell is free of interference from any other parts
or internal structure and therefore transfers the aerodynamic loads "seen" by
it to the beam at the wrist. Strain gages mounted on the lower arm beam were
used to measure the loads acting on the shell. It should be noted that because
the lower arm beams are attached to the free end of the upper arm beams, the
gages on the upper beam measured the combined load of the whole arm. (Left and
right arms are mirror images of each other.)
The design concept of the legs is the same as for the arms. The upper
leg beams are bolt .t to the torso frame. The lower leg beams are attached to
the upper beams with a tapered joint at the knee. The upper and lower leg
shells are all in left and right halves and bolt to the free ends of the re-
spective beams. Strain gages on the lower leg beams measured the aerodynamic
loads on the lower leg/foot shell. Gages on the upper leg beam measured the
combined loads on the whole leg.
Forebody
The forebody model (Figure 2) is a half-scale representation of the nose
and cockpit section of an F-16 fighter aircraft. The main structure of this
model is made up of wood and aluminum frames and bulkheads. Attached to this
framework are wood splines and formers over which is laminated the external
shell of glass reinforced polyester plastic.
Supported Dy the internal framework of the forebody are two tracks, one
along each of the side walls of the cockpit. They are parallel to the path
along which a crewman would travel during ejection from the aircraft cockpit.
The forebody model is supported and positioned by slides and rollers which
contact the surfaces of these tracks (Figure 3). The slides and rollers are
part of the carriage assembly which attaches directly to the sting.
13
-Ilk
14
Cockpit Wall
Tracks
Slides Cockpit Wall
Rollersr
Tracks mounted tocockpit walls translatebetween slides
Sting Slides pivotin carriage
Carriage attachedto sting
Figure 3. Track and carriage.
15
Conceptually the crewman/seat (fixed in position relative to the carriage)
travels along the tracks, simulating an ejection from the cockpit. In actual
practice the forebody model is retracted (Figure 4), with the tracks translating
between the slides, exposing the crewman/seat model to increasing amounts of the
airflow.
Flow Diverters
During portions of the test, two flow diverters were attached to the fore-
body model. These diverters (shown in Figure 5) were in positions ahead of, and
to either side of the heads up display. They represent a design concept intended
to reduce the force of the flow hitting the crewman still in the cockpit after
canopy separation.
3.2 RELATED HARDWARE
Support Hardware
The models were supported by a sting through the bottom of the forebody,
which mounts to the test section floor (Figure 6). The crewman/seat model mounts
with the use of an adapter to the main force and moment balance (described in
Section 4.2). This balance mounts to the end of the floor mounted sting assembly.
In addition to supporting the crewman/seat model, the sting supports the fore-
body through the carriage-track interface. The carriage was assembled directly
around the outermost sting segment (Figure 3). When in place, the carriage is
an integral part of the sting assembly. At the four corners of the carriage are
pairs of rotating slides. Each pair of slides fits on opposite sides of one of
the forebody tracks (discussed in Section 3.1). Rollers, one adjacent to each
set of slides, position the carriage slides on the tracks. Translation of the
forebody with its tracks sliding between the carriage slides, is the method used
to simulate crewman/seat ejection from the cockpit.
16
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Hydraulics
The position of the forebody model relative to the carriage (and there-
fore the crewman/seat is changed by a remotely controlled hydraulic cylinder
(Figure 4). The fixed end of the cylinder attaches to the upper sting segment
pitch change ears. The piston rod end attaches to the forebody aft of the
canopy section. The cylinder is positioned parallel to the forebody track just
behind and below the left track.
Forebody Position Indicator
To measure the position of the forebody relative to the carriage, a rack
was installed in the forebody parallel to the tracks. A pinion was mounted on
the carriage in such a way as to mesh with the rack. A 10-turn potentiometer
(supplied by AEDC) was turned by the pinion. The potentiometer was calibrated
so that the forebody position could be determined during the test. The range
of forebody movement is approximately 24 inches. The zero position was with the
forebody completely extended, which simulates the crewman in the pre-ejection
position within the cockpit.
Forebody Seal
The forebody model is open on its top to simulate an open cockpit. The
model is open on its bottom to accommodate the support hardware. The sting,
balance and crewman/seat model, when assembled, must have clearance completely
through the forebody model. To prevent undesirable flow through this cavity,
it must be sealed around the sting. Because of the relative motion between
these parts, the seal cannot be rigid. A seal, originally designed at AEDC,
was made and fitted to the forebody model. The seal consists of two parts
(Figure 7). Each part (Figure 8) has a long 5-inch wide strip of fiberglass
cloth, impregnated with a sealout and with 0.062 x 0.75 x 5-inch brass strips
riveted to both sides. The brass strips are positioned crosswise to the fiber-
glass with approximately 1/8 inch between the strips. They provide the strength
to resist the airflow. One part of the seal closes the gap bound by the fore-
body tracks, the carriage and the top back part of the cockpit behind the pilot's
21
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head. The other part of the seal closes the ap bound by the tracks, the car-
riage and the bottom forward part of the cockpit below the pilot's feet. The
seal parts both slide in slotted tracks which are positioned above and parallel
to the carriage tracks. One end of each seal part is attached to the carriage.
The other end of each seal extends past the ends of the seal tracks, over a
pulley and is contained in a conduit. The thin, wide conduits run vertically
at both the forward and aft cockpit bulkheads. Tension is maintained on each
of the seal parts by sets of two spring loaded cable reels attached to the
bulkhead. The ends of the cables are connected to the free ends of the seal.
The seal was not air tight, but significantly restricted the flow through the
cockpit.
25
SECTION 4
INSTRUM NTATION
4.1 LIMB AND NECK GAGES
Crewman model limb and neck moments and neck lift force were measured
by 37 four-arm-active strain gage bridges. There are two bridge locations on
each of the nine beams. At each location, bending moments were measured in
each of two planes, 90 degrees apart, for a total of 36 moment bridges. The
one remaining briges measured lift force on the head (axial with respect to
the beam).
4.2 CREWMAN/SEAT TOTAL FORCES AND MOMENTS
The total forces and moments acting on the crewman/seat were measured
by the AEDC's Task Mk VII 2.5-inch balance, PWT No. 6-2.5-2.5-1.85 M-b. The
balance was mounted with the fixed end toward the tunnel floor and with the
free end approximately 38 degrees downstream. The top of the balance was on
the upstream side.
4.3 STATIC PRESSURE MEASUREMENTS
Static pressures were measured at six locations on the crewman/seat
model. They were on the:
1. Head between the eyes
2. Middle of the chest
3. Middle of the abdomen
4. Front of lower left leg between knee and foot
5. Front of lower right leg between knee and foot
6. Back of seat.
26
These locations are illustrated in Figure 9. Each orifice was connected by
flexible tubing to one of six individual pressure transducers mounted outside
the tunnel test section.
4.4 FLOWFIELD VISUALIZATION
Schlieren movies and schlieren still photographs were taken during the
test.
4.5 HYDRAULIC PRESSURE
The vertical position of the forebody model was changed and maintained
by a remotely controlled hydraulic cylinder (Section 3.2). The hydraulic
pressure required to do this can be used to evaluate the longitudinal loads
on the forebody model. Pressure transducers were placed in the supply lines
on both the push and pull sides of the hydraulic cylinder.
4.6 POSITION AND ATTITUDE INDICATORS
Two facility angular position indicators were mounted in the models to
determine model attitude. One indicator was mounted in the forebody to measure
pitch attitude. The other was mounted in the ejection seat model to measure
pitch attitude.
A 10-turn potentiometer was mounted inside the forebody for determina-
tion of the forebody translation position.
4.7 CALIBRATION
Calibration of all instrumentation was conducted at AEDC.
The crewman limb and neck beams were calibrated as bare balances, not
attached to the crewman model and without the fiberglass covered shells
(Section 3.1). The beams were then returned to Lockheed for assembly of the
crewman/seat model.
27
*1 ' I ~Pressure
II ~' (PSSOR)
Static Pressure(Connect to-Scanner Valve I
LEGEND
P511 a Pressure Static Head '
PSSBR a Pressure Static Scat Il Scanne fitBack ReferenceVav
PSC x Pressure Static ChestPSA a Pressure Static Abdo*"PSLR x Pressure Static Lea RigtPSLLR x Pressure Static Leg Left
Figure 9. Pressure orifice locadon.
28
Check loads were applied to both the main balance and the limb and neck
beams after the models were installed in the wind tunnel.
29
SECTION 5
TEST PROGRAM
The following sections describe the test program that was conducted.
Subjects include model configurations tested and the independent test param-
eters that were varied during the test.
5.1 CONFIGURATIONS TESTED
Two configurations were tested, "Base Model" and "Base Model plus Flow
Diverters." In both configurations, the crewman was seated in the ejection
seat with arms and legs in the non-flail position. The crewman/seat model
was positioned in the cockpit of the forebody model or at different positions
along the ejection path. The only difference between the two configurations
was the presence of flow diverters on the top of the instrument panel for
"Base Model plus Flow Diverters." The flow diverters are described in
Section 3.1.
5.2 TEST PARAMETERS
The parameters that were varied during the test were Mach number, dy-
namic pressure and ejection position. The combinations of Mach number and
dynamic pressure at which the models were tested are presented in Table 1.
The models were alwyad at a nominal attitude of zero.
At each combination of Mach number and dynamic pressure, the' model was
tested at five ejection positions. Ejection position is a linear measurement
along the direction of motion of an ejecting crewman. Zero is the position
of the crewman/seat prior to the start of the ejection sequence. The nominal
positions at which data were taken are 0, 6, 12, 18 and 24 inches model scale.
30
TABLE 1. TEST CONDITIONS
Mach Reynolds Q Total TotalNo. Number (psf) Temp., TT Pressure, PT
(RE x 106 /ft) (F) (psf)
0.4 1.18 100 80 9970.4 1.77 150 80 14960.4 2.35 200 80 19940.4 2.59 220* 80 2193
0.6 1.55 200 110 10120.6 2.32 300 110 15180.6 3.09 400 110 20240.6 3.87 500 110 25300.6 4.64 600 110 3036
0.8 1.23 200 110 6810.8 1.84 300 110 10210.8 2.46 400 110 13610.8 3.07 500 110 17010.8 3.69 600 110 2042
1.0 1.09 200 110 5411.0 1.63 300 110 8111.0 2.18 400 110 10821.0 2.72 500 110 13521.0 3.27 600 110 1622
1.2 1.00 200 110 4811.2 1.50 300 110 7221.2 2.00 400 110 9621.2 2.51 500 110 12031.2 3.01 600 110 1443
Maximum dynamic pressure obtainable at M = 0.4.
31
5.3 RUN SCHEDULE
A collation of test runs at common test conditions and for common con-
figurations is presented in Table 2.
TABLE 2. RUN NUMBER COLLATION
Q Mach Number
Configuration (psf) 0.4 0.6 0.8 1.0 1.2
Base Model 100 39150 38200 37,50 8,17,34 11 12 16220 36300 18 22 30 26400 19,46 23,47 31,48 27,49
500 20 24 32 28
600 21 25 33 29
Base Model with 200 60 59Flow Diverters 400 54 55 56 57,58
32
SECTION 6
DATA REDUCTION AND PRESENTATION
The following sections discuss data reduction and data presentation.
Sections 6.1 through 6.4 cover reduction of data obtained from limb and neck
beams, crewman/seat main balance, pressure transducers and forebody hydraulic
cylinder pressures. Sections 6.5 and 6.6 include data tabulations and plots.
6.1 LIMB AND NECK
Crewman strain gage force and moment data were reduced to resultant crew-
man limb forces and moments. The resultant forces and moments that were calcu-
lated are:
FSXL,R FHXMSXL,R MHXFSYL,R FHYMSYL,R MHY
FHZFEYL, RMEYL,RFEZL,RMEZL,R
FHYL,RMHYL,RFHZL,RMHZL,R
FKXL,RMKXL,RFKYL,RMKYL,R
Figures 10 through 14 show the resultant forces and moments calculated
and the axis systems and moment reference centers for each beam. Tables 3
and 4 list the data reduction equations.
33
FHX
MHY
FHZ FHiZ
FH Y$ FHX
FRONT VIEW
Figure JO. Neck beam axis system.
34
NOTE: RIGHT ARM SHOWN. LEMT ANDRIGHT ARMS HAVE SAME POSITIVE,DIRECTIONS.
L, R
FSYL,R -- 4 % FXMSYL,R
FRONT VIEW SIDE VIEW
Figure 11. Upper arm axis system.
35"
NOTE: LEFT AND RIGHT ARMS HAVE
SAME POSITIVE DIRECT'IONS.
FEYL ,R
TO~P VIEW MEYL,R
FEZL,R
SIDE VIEW MEZL,R
Figure 12. Lower arm axis system.
36
NOTE: IEFT LEG SHMN. LEffT .D RIGff LEGSHAVE S&1E POSITIVE DIRECTIONS.
FHYL,RI"
TOP VIEW
FIIZL,R
I
SIDE VIEI
Figure 13. Upper leg axis system.
37
NOTE: LEFT AM~ RI~if LEGS HAVE SMlEP~X'ITIVE DIRECTIcMS.
IWi L, R HKXL,R
FKYL,R - FKXL,R
0
0
0
00
FRONT VIEW ~ SIDE VIEW
Figure 14. Lower leg axis system.
38
TABLE 3. CREWMAN HEAD STRAIN GAGE DATA REDUCTION EQUATIONS
FHZ is measured directly by bridge
= Mil - MHi2KDi
Mli = MHi2 + (KXi x FHi)
where i = X,Y
and MHil = moment measured by lower (away fromtorso) bridge
MHi2 = moment measured by upper (nearest torso)bridge
KDi = distance between strain gages 1 and 2 (deter-mined during calibration)
KXi = distance from HRC to strain gage 2 (determinedduring calibration).
TABLE 4. CREWMAN LIMB STRAIN GAGE DATA REDUCTION EQUATIONS
Fij = Mijl - Mij2
KDij
Mij = Mij2 + (KXij x Fij)
where i = EY, EZ, HY, HZ, KX, KY, SX, SY
j = L, R (left or right side of crewman)
and
Mijl = moment measured by lower (away from torso)bridge
Mij2 = moment measured by upper (nearest torso) bridge
KDij = distance between strain gages 1 and 2 on a limb(determined during calibration)
KXij = distance from MRC to strain gage 2 (determinedduring calibration).
39
6.2 CREWHANISEAT
The crewman/seat force and moment data were reduced to coefficient form
in the body axis system. Force-area and moment-volume parameters, both model
(half) scale and full scale, were also calculated in the body axis system. The
axis system and the moment reference center are shown in Figure 15. The moment
reference center is located at the theoretical crewman/seat center of gravity.
Transfer distances from the balance center are shown in Figure 16. The force
and moment coefficients are based on a reference area, S, equal to the model
projected frontal area and a reference length, d, equal to the diameter of a
circle whose area is equivalent to the model projected frontal area. The values
of these dimensions are:
S = 1.600 ft2
d = 17.128 in.
6.3 CREWMAN PRESSURES
Crewman static pressure data were measured and presented in both engineer-
ing units of pounds per square foot absolute and as static pressure coefficients.
The variables are:
Static StaticPressure Pressure Coeff. Location
PSH CPH Head
PSC CPC Chest
PSA CPA Abdomen
PSLL CPLL Left Leg
PSRL CPRL Right Leg
PSSBR CPSBR Seat Back (Ref.)
Pressure orifice locations were discussed in Section 4.3.
40
94-
ozo
94~
CV-
C-i)
41J
FIN
II
Figure 16. Balance transfer distances.
42
6.4 FOREBODY
A calculation can be made of the force required to hold the forebody in
position (FFHC). This can be done using the pressures in the hydraulic cylinder
that moves and holds the forebody in place. These pressures were measured during
the test as described in Section 4.5.
6.5 DATA TABULATIONS
All the data of one type from all the different test points are listed
together on a series of pages in the appendix. Each page has as identification
the test number and the date it was produced. Each line of data for all types
has the following identification.
Run and test point number (RN/PN)
Mach number (MACH)
Dynamic pressure (Q)
Ejection position (EJPOS)
6.6 DATA PLOTS
Selected data have been plotted in the form of comparison . ts. These
plots are discussed in Section 7.
43
SECTION .7
RESULTS
The complete results of the test are presented in the tabulated listing
of the Appendix. These listings are summary in nature with data grouped accord-
ing to the instrumentation from which the data were obtained. Within each group,
the data obtained during each test run/point are displayed on a separate line.
The groupings are:
Crewman/seat total force pages 74 through 78
and moment coefficients
Lower leg force areas and pages 74 through 83
moment volumes
Upper leg force areas and pages 79 through 83moment volumes
Lower arm force areas pages 84 through 88
and moment volumes
Upper arm force areas pages 84 through 93
and moment volumes
Neck force areas and pages 89 through 93
moment volumes
Pressure coefficients pages 94 through 98
On the first night of testing fouling occurred between the crewman's
left leg and the forebody at EJPOS = 0. The total crewman/seat and left leg
data taken at EJPOS = 0 are therefore unusable and are marked out in the tables.
Another problem was found relating to pressure measurements. During the first
night of testing, the airstream pulled pressure tubes up through the floor and
consequently some tubes were pinched so that pressures could not be read by the
pressure transducers. Therefore those data are missing from the tables.
Determined from collected data were: crewman/seat forces and moments
in engineering units, force areas and moment volumes (both model scale and
44
full scale) and coefficients; crewman limb and neck forces and moments in
engineering units and in force areas and moment volumes; static pressures
and pressure coefficients and; balance outputs for main and crewman limb
and neck balances.
In addition to the complete tabulations presented in the Appendix, se-
lected data are graphically presented and discussed.
The objective of the wind tunnel test program was to evaluate the Reynolds
number influence on the aerodynamic force measurements made with the one-half
scale models of the ejecting crewman and F-16 forebody. Reynolds number is the
dimensionless ratio of inertial to viscous forces in fluid flow; it is used to
properly simulate these forces during scale model testing. The forces acting on
a model are a direct function of the inertial forces of the flow which are di-
rectly proportional to the product of the dynamic pressure of the flow (Q) and
the characteristic area of the model (S). The model forces are also influenced
by the viscous forces in the flow. However, they are not a direct function of
the viscous forces. The viscosity of the fluid affects the boundary layer (type,
transition point, separation point). The viscous force is directly proportional
to the velocity through the fluid (V), the characteristic length of the model (k)
and the absolute viscosity (p).
RN = 2 Inertial Force = p V kP V Z Viscous Force
In wind tunnel testing of scale models the characteristic length is gen-
erally smaller than the full-scale article and thus the unit Reynolds number
(Reynolds number per foot length (pV/p) has to be larger than anticipated flight
values in order to match full-scale Reynolds number values. The unit Reynolds
number is changed in the wind tunnel by operating the tunnel at a higher total
pressure which increases the density and the unit Reynolds number of the test
medium.
The problem in evaluating Reynolds number effects is that they are very
small relative to the inertial effects and neither the kinematic viscosity (p/p)
45
nor the model size can be changed easily during a wind tunnel test. Changes
in velocity also change the dynamic pressure. It is directly proportional
to the square of the velocity. Therefore, changes ini Reynolds number are
accompanied by changes in dynamic pressure.
To keep the viscous influence from being obscured by the larger inertial
airload on the model, the loads measured with the model need to be converted
to inertia force coefficient form (C = Force/Q S) or force area parameter
(CF = Force/Q). The influence of Reynolds number on the inertia forces can
then be identified by changes in the coefficient or force area parameter ob-
tained by testing at different Reynolds numbers.
The importance of the simulation of Reynolds number in general aero-
dynamic testing is shown in Figure 17 which presents the drag coefficient of
a sphere versus Reynolds number for several Mach numbers. The importance of
Reynolds number simulation is apparent in the drag coefficient variation with
Reynolds number. It is noted that the Reynolds number influence on the drag
force of spheres is not nearly as pronounced as it is in the drag coefficient.
Thus coefficient data or force area/moment volume data is used to evaluate
trends with Reynolds numbers.
The evaluation of the influence of Reynolds number on the crewman/seat
and component data (limbs) included an evaluation of the influence of the F-16
forebody and the local flow conditions on the crewman. A schematic of local
flow on the F-16 forebody and crewman/seat is presented in Figure 18.
Estimates of the boundary layer state on the forebody were determined
over the range of test Reynolds numbers. Using an average transition Reynolds
number of 106 the boundary layer should experience natural transition to
turbulent within the first foot of length. Thus, the boundary layer state
approaching the cockpit cavity should be a fully developed turbulent boundary
layer for all test conditions. At the cockpit cavity the boundary layer will
form a shear layer between the external flow and the cavity flow as shown in
46
4p 0
z00 .4 N
N co Io0 4.)
0
0
00
wl
Ci* U)
00Cl! 0)
CO.)P,
1) 130
C4j
47-
cqI
1L41
0
0
10, 0
rX4 0
.1-I
4)
00
'oo
48
Figure 18. The difference in the height of the shear layer for the low Reynolds
number condition has been exaggerated in the schematic to si. w relative trends
with changes in Reynolds number.
Local pressure coefficient data on the crewman centerline provide a means
for assessing any changes in the upstream flow conditions. A change in local
pressure coefficient would identify a change in the local cavity flow, shear
flow conditions or external flow conditions as the crewman/seat is traversed
through these flow environments. It was expected that if upstream flow condi-
tions changed due to Reynolds number variation these changes would show up in
changes in the local pressure coefficient for taps located in the shear flow
between the cockpit flow and the external flow. This is due to the sensitivity
of the shear flow velocity profile to any integrated change upstream of the
forebody nose.
The location of the pressure instrumentation was presented in Figure 9.
Selected body pressure coefficient data are presented in Figures 19 and 20.
These figures show that the magnitude of the pressure coefficient data is the
same over the range of Reynolds numbers tested. Each symbol on each curve
corresponds to the table of flow conditions on the figure. Thus each symbol
plotted (A, B, C, etc.) represents a different unit Reynolds number. If the
pressure coefficient data varied with Reynolds number the plot symbols would
be separated and each symbol would be identifiable as a different value of
pressure coefficient corresponding to different Reynolds numbers. Since the
symbols are essentially plotted on top of each other, separate symbols (A, B,
etc.) cannot be identified showing that the local pressure coefficient is
essentially identical for each unit Reynolds number tested.
The cavity pressure environment is noted in Figures 19 and 20 to have
a pressure coefficient value of approximately -0.20 as noted by the value of
CPSA, CPSC, CPSLL and CPSRL at the zero ejection position. The chest and
abdomen areas are in a cavity flow environment at zero ejection position
since the pressure coefficient for these locations at zero ejection position
is the same as the leg values. The head is in a shear flow environment at
49
4
Wd~
.4 S LI) I) 0 CO >
Ic.* U) 0 0 0 0
.4;
0 U)
CU 0~1 0 0 k05
004 r- Rol (%J
~~.C cn o n
44 C- C> C>ol-4C L.C> c
* L)Ix-
I-0\ in th If I I
a-s
CL
441
Co4
.j.
51
zero ejection position since the head pressure coefficient is higher than the
cavity and lower than the pressure coefficient at higher ejection positions.
The leg pressure coefficient data show the most extreme range (Cp =
-0.2 to +1.0) over the ejection positions (Figure 20). The leg pressure data
shows cavity flow conditions for ejection positions of 0, 6 and 12 inches.
Shear flow conditions exist at 18 inches and full stagnation external flow
conditions exist at 24 inches (C = 1.0). It is noted that there is no varia-
tion in the leg pressure coefficient data at an ejection position of 16 (shear
flow) or any other position. Thus the pressure coefficient data show that there
is no Reynolds number influence on the far field (F-16 forebody) or near field
(cavity, shear) flow environment near the model.
Figures 21 througi -I show selected examples along with supporting data
on spheres and cylinders, where the aerodynamic data changes with Reynolds
number. These plots, like the pressure coefficient data in Figures 19 and 20
use symbols (A, B, C etc.) to identify different tunnel flow conditions (Q and
unit Reynolds number).
Data plots that show a relatively large separation of the symbols identify
aerodynamic data that are substantially influenced by Reynolds number. Data plots
that show little or no separation of the symbols show little or no Reynolds number
influence.
Figure 21 shows the influence of Reynolds number on certain crewman/seat
data. The largest Reynolds number influence occurred on the crewman seat normal
force coefficient (CN) as shown by the separation of the plot symbols (A, B, C,
etc.). The axial force coefficient (CA) and side force coefficient (CY) show
little change with Reynolds number.
Figure 22 shows the influence of Reynolds number on the crewman head/
helmet lift data. The head/helmet lift force area parameter presented in the
figure shows a large Reynolds number influence by the large separation of the
52
f4-f
C0
- -- 0 r-4 I L- CA)
-4 r- r- M LO
0- UI
00- 4L Cl) 00 C)Chi
PJ 0 ) 0
4-
o4 11, *o*R
m ~ I-'>
* - - - -- - 0U(
r4
01A In u
$-4
.44
IA0
- aO. C4
.4 .4, - *~) 4!
2- 01
z.
n5
4-)
I7 CD uo n toC>00%
Ci) N> 0) C) 'ti>
C:) N: c) c) -
15 '.4
r 00 0 00 C)r-
U) N I) ~j- U')U'
OD11 r 0) £C) r- r4 N o
C,4 C-4
tn -n - - --n
1:4nX.#A qD \..Jn
54-
symbols corresponding to different unit Reynolds numbers. The head/helmet
lift force area parameter CFHZ was plotted versus Reynolds number (Figure 23)
of the head to evaluate specific Reynolds number trends. The model head/
helmet is nearly spherical with a diameter of approximately 5 inches. The
head Reynolds number is thus determined using the flow unit Reynolds number
(table on Figure 22 for Mach 0.6) times the head diameter in feet (0.416).
Figure 23 shows that there is a large change in the head/helmet force5
area parameter in the Reynolds number range between 6 to 9 x 10 . Figure 17
shows a corresponding large change in the drag of spheres at Mach 0.6 in the
same Reynolds number range. Thus spherical shapes experience a change in the
boundary layer separation characteristics in the Reynolds number range between
6 to 9 x 105 at Mach 0.6. Figure 23 shows that a similar phenomenon must exist
on the head/helmet in this Reynolds number range. The figure also shows that
extrapolation. of the helmet/head lift data at the lower Reynolds numbers to
higher Reynolds numbers would develop erroneous results. It is noted that a
head lift data repeatability problem existed during the test and the head lift
area parameter has a high level of uncertainty. The level of uncertainty,
however, is not felt to negate the Reynolds number trends and conclusions.
The head lift is envisioned as being due to differences in the flow sepa-
ration location over the lower and upper portions of the head as shown in Figure
24. Figure 24 shows the change in sphere lift coefficient versus the change in
the lower surface local boundary layer separation position. The model crewman's
neck corresponds to the angular location of the lower separation location that
creates large sphere lift coefficients. Figure 24 shows a Reynolds number trend
that corresponds to the trend at Mach 0.4 (Figure 22) rather than the trend at
Mach 0.6. The trend in Figure 24 corresponds to the lower Mach number.
The head/helmet axial force area parameter and moment volume parameter
shown in Figure 25 also shows some variation with Reynolds number. The largest
variation in head/helmet axial force occurs at an ejection position of 12 inches
and corresponds to the ejection position of maximum head/helmet lift (Figure 22).
55
0
rC'4
H
0 n04
- - z
ZD H
00 0
coa
0
00 0 U
W 0oo 0
-00 41
a) -iin
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a) )
U,;
Cf..
rZ4
Nl
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56
0
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00 0-
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0 $4
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CD 00) p
8 H -
H LH
ca 0 U
0H H0
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N 00
oc
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C-)4
00
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000-:3.
57
0: 00 1 Lf) CA
r- -r-- V) IfO
44.4V) 000)C)0 U
v4r- 4 -
w4 r 4.J
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03u
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14-
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58U
The upper arm axial and side force area and moment volumes are presented
in Figures 26 and 27 for Mach number of 0.4 and 0.6, respectively. Both the
axial force area (CFSXL) and the side force area (CFSYL) parameters show varia-
tion with Reynolds number by the spread of the curves. The side force area
parameter has the largest variation with Reynolds numbers. The side force area
parameter at Mach 0.6 (Figure 27) is presented versus Reynolds number based on
the upper arm diameter in Figure 28. The figure shows that the high upper arm
side force area parameter at the largest ejection position would be difficult to
extrapolate to high Reynolds number values based on lower Reynolds number trends.
The side force area parameter on the upper arm is due to the flow around
the upper torso arm combination as shown in Figure 29. Thus the local flow
angle approaching the upper arm is at an angle to the upper arm. The side
force on irregular cylinder shapes versus Reynolds number is also presented
in Figure 29. The figure shows that the critical Reynolds number is in the5range of 5 to 9 x 10 . It is noted in Figure 28 that the change in upper arm
5side force occurs in-the Reynolds number range from 5 to 9 x 10 . Thus it
appears that the upper arm side force may have similar Reynolds number trends
as an irregular cylinder.
The lower left arm lateral and lift aerodynamic data is presented in
Figures 30 and 31. These figures show smaller Reynolds number variations
than the upper arin data.
The upper leg vertical and lateral force and moment aerodynamic character-
istics are presented in Figures 32 and 33. Figure 32 shows a substantial varia-
tion in upper leg lift and moment data at Mach 0.4. Figure 33 shows that the
Reynolds number influence is evident at the higher Mach numbers although reduced
considerably. The lower leg aerodynamic characteristics showed almost no varia-
tion with Reynolds number except at the higher ejection positions as shown in
Figures 34 and 35. This is due to the lower leg being in the cavity flow field
of the cockpit except at the higher ejection positions. The influence of Reynolds
number is clearly discernible at the high ejection position however.
59
U;
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71L
SECTION 8
CONCLUSIONS
The wind tunnel test program and the resultant data show a substantial
Reynolds number influence on certain aerodynamic data. Analyses of the prob-
able forebody boundary layer conditions and pressure coefficient data show
that the F-16 forebody flow field outside the cockpit, the cavity flow inside
the cockpit, and the shear layer flow field have no discernible change with
Reynolds number for the range of unit Reynolds number tested.
The changes in the crewman/seat aerodynamic coefficient data and the
crewman limb force area and moment volume parameters due to unit Reynolds
number changes can be attributed to changes in the local flow conditions on
the crewman spherical (head/helmet) and cylindrical (arms, legs) components.
The Reynolds number influence appears to be similar to the local changes in
boundary layer flow and separation conditions that occur on flow over isolated
spheres and certain irregular cylinder configurations.
In most cases where a Reynolds number related variation does occur, data
magnitudes are higher (more conservative) at the lower Reynolds numbers than
at the higher numbers. For data where this is not the case, the variation due
to Reynolds number is either small (Figure 35) or the magnitude at all dif-
ferent Reynolds numbers at the specific Mach number is significantly less than
the magnitude at other Mach numbers (Figures 22a and 22b). This suggests that
for testing the 1/2 scale crewman/seat model at a single Reynolds number, less
than full scale Reynolds number, a value between 1.5 and 2.3 x 10 6/ft should
be selected. Further restriction of the M = 0.4 test conditions to RE = 1.77
x 10 6/ft or above would avoid the less conservative value for CFHZ at M
0.4 seen in Figure 22. Testing the 1/2 scale model at dynamic pressures
between 150 and 300 psf will produce Reynolds numbers which will give reason-
able conservative results throughout the 0.4 to 1.2 Mach range.
72
-App,-e nd ix
TABULATED TEST DATA
73
WLI
0
4x fr "0 1& I, en c, 0f fn to In I Q N i 0" CQ r I C"C, 'ID ' m 1 o O Q in. 0-1C'fP Q-Ifu N 00CQa n UVO-- 00CQ.in C
Iii Q9 000. 00 0 C D n 0000 000 0 t ."O 00 C Q0004 ~ CD 0.. *; *; * * * *; ~ * e be ,
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REFERENCES
1. Braddock, W. F., "Definition of Windblast Flow Fields Pretest Report,"IMSC-HREC TR D784481, Lockheed Missiles & Space Company, Huntsville,Ala., August 1981.
2. Guyton, F. C., "AFAMRL/Lockheed Ejection Set Final Tabulated Data,"Project No. P41G-09, Arnold Engineering Development Center, Arnold AFS,Tenn., September 1981.
3. Hoerner, S. F., Fluid-Dynamic Drag, published by the author, MidlandPark, N.J., 1958.
4. Davies, J. M., "The Aerodynamics of Golf Balls," J. of Appl. Phys., Vol.20, September 1949, pp. 821-828.
5. Polbamus, E. C., E. W. Geller, and K. J. Grunwald, "Pressure and ForceCharacteristics of Noncircular Cylinders as Affected by Reynolds Numberwith a Method Included for Determining the Potential Flow About ArbitraryShapes," NASA TR R-46, (1959).
99US GWJSRNMENTPAINTINGOFFICE tM- 559.-O6/10221
