tail rotor impressed pitch and sideslip angle. The reduced sideslip angle ...... AIRSPEED, knots. Figure. 65.-. Tail-rotor gearbox vibration versus calibrated.
NASA
Technical
Memorandum
85843
:_84"32381
NASA Rotor System Research Aircraft Flight-Test Data Report: Helicopter and Compound Configuration R.E. Erickson, R.M. Kufeld, J.L. Cross, R.W. Hodge, W.F. Ericson and R.D.G. Carter
0
August
1984
tU/ SA National Aeronautics Space Administration
and
NASA Technical
Memorandum
85843
NASA Rotor System Research Aircraft Flight-Test Data Report: Helicopter and Compound Configuration R. E. Erickson R. M. Kufeld J. L. Cross, Ames Research Center, Moffett R. W. Hodge W. F. Ericson R. D. G. Carter, Sikorsky
Aircraft
N/ A National Aeronautics and Space Administration Ames Research Center Moffett Field, California 94035
Company,
Field, California
Stratford,
Connecticut
TABLE
OF CONTENTS
Page i
SUMMARY.**,*,****.*,**,,********,*****,,**
i.
INTRODUCTION
AND
TEST
SUMMARY
......................
2
2.
HANDLING QUALITIES . .......................... Control-System Rigging ......................... Flight Envelope ............................ Takeoff and Landing Procedures ..................... Level-Flight Trim ........................... Static Stability ............................ Comparison of Predictions and Flight Data ............... Wing-Actuator Imbalance ........................
7 7 7 8 8 9 i0 i0
3.
VIBRATIONS
69 69 69 69 69
...............................
Engine Vibration ............................ Crew Comfort .............................. Airframe Vibration ........................... Frequency 4.
STABILATOR Stabilator Upper
.
Analysis
MAIN-
AND STABILIZER ........................ Stress and Loads ......................
Horizontal AND
. ..........................
Stabilizer
TAlL-ROTOR
6.
LOAD-CELL
SYSTEM
7.
CONTROL-SURFACE
8.
PERFORMANCE
LOADS
...................... AND
STRESSES
.................
............................ LOADS
DATA
AND
PYLON
TEMPERATURE
A-- FLIGHT-TEST
PLAN:
HELICOPTER
APPENDIX
B--FLIGHT-TEST
PLAN:
COMPOUND
APPENDIX
C-- RSRA
REPORTS:
FLIGHT
RSRA
...............
CONFIGURATION CONFIGURATION 740
.................................
93 iii
............................
APPENDIX
REFERENCES
83 83 84
144 158
........... ............
.................
174 196 227 240
iii
NASAROTOR SYSTEM RESEARCH AIRCRAFTFLIGHT-TESTDATAREPORT: HELICOPTER ANDCOMPOUND CONFIGURATION R. E. Erickson, R. M. Kufeld, J. L. Cross, R. W. Hodge,* W. F. Ericson* and R. D. G. Carter* AmesResearch Center SUMMARY This data report documents the flight-test activities of the Rotor System Research Aircraft (RSRA), NASA740, from June 30, 1981 to August 5, 1982. Tests were conducted in both the helicopter and compoundconfigurations. Helicopter vertical-drag test results are reported in NASAContractor Report 166399, December1981. Compoundtests reconfirmed the Sikorsky flight envelope except that main-rotor blade-bending loads reached endurance at a speed about i0 knots lower than previously. Wing incidence changes were madefrom 0° to i0 °.
*Sikorsky Aircraft
Company,Stratford,
CT 06601
i.
This System
report
The 18
presents
Research
three
were
high-speed
The
purpose
and to testing
Center.
See
In
included
and
these
reference
hr
17
were
in
the
June i0
to
to
of
A
and
obtained
August
1982,
time
the B
pilot
the
during
Rotor
which
time
In
addition,
accumulated.
flight
and
tests.
the
The
flight
training
developed aircraft
from
configuration.
flight
provide
for
by was
Sikorsky delivered
and
the
flight
reports
the
in
NASA
and
Aircraft to Ames
Army
during Research
report.
were
to
results
before
Sikorsky
objectives
attained,
evaluating
various
wing
flight
envelope
angles,
was
collective
posi-
scheduling.
following
i.
Airspeed
2.
Load
factor
3.
Side
slip
4.
Rotor
was
5.
Wing
angle
6.
Angle
of
7.
Collective
68%
8.
Flap
0 ° to
to
the 181
actual
knots
1.78
g
envelope
evaluated:
CAS
to
0.37
g
±15 °
speed
105%
bank
and
to
0 ° to
angle
1
the
SUMMARY
compound
1981
min
appendices
was
TEST
flight-test
conducted
in
respect
The
subject
from
and
for
test
with
rotor-speed
Tables
of 740)
period
tests
i
all
respectively.
2
The
97%
(trimmed)
i0 °
±50 °
are
(low
speed)
to
18%
25 °
flight
following
logs
for
remarks
the
are
helicopter
based
on
and
the
compound
results
flight
obtained
tests,
during
the
tests.
i.
Structures--The
speed
the
frame
stress
incurred
hr
was
and
fication ered.
to Table
2.
or
load
at
a rate
3
stresses
problems
were
encountered
low it the
stabilizer is
a
demonstrated
blade
considering the
airspeed
outboard
However,
Handling
control
encountering
maximum
main-rotor
configuration.
wing
tests
are
expanded
tions,
600
taxi
AND
evaluate the flight envelope at Wallops Island before the
general,
actually
the flown
of
summary
(RSRA-A/C
covers
flights
test plans (FTP) appendix C.
pilots their
a
Aircraft
report
test
INTRODUCTION
enough
to
was
evident
possible and
summary
permit
of
been
accomplished
any
control
margin
such
effect
the of
attachment
of within
fatigue
airframe demonstrated
2
pilot
at
testing
air-
that in
the
airframe
must
which
life
changes, be
seriously
was current of modiconsid-
damage.
the The
damage
planned
areas
CAS
unanticipated
configuration
the
concerns.
knots
No
flight
within
cumulative
181
fatigue
additional
adverse
qualities--Operation
has
was
limiting.
and
that
stabilator
log
were
with
workload
100% envelope was
rotary
and
fixed-
without acceptable
with
good been
control harmony. expected, but the
Problems with directional pilots experienced little
control during difficulty.
the
takeoff
roll
had
3. Airframe vibration-The cockpit environment was very acceptable throughout the envelope. The ll-Hz response reported during the Wallops Island tests was encountered at 150 knots IAS as expected and confirmed as a beat between the tailrotor blade first edgewise responseand rotor latter response will continue to be monitored
were
4. Engine vibration--Both generally acceptable, with
the the
T-58 and exception
gearbox measurement. Troubleshooting The measurement system was recalibrated resolved;
however,
these
parameters
speed. The filtered on telemetry.
TF-34 engines were monitored. of the No. i TF-34 vertical
indicated and it
will
level
a possible is assumed
continue
to
be
instrumentation that the problem
monitored.
of
the
Levels accessory problem. is
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i-i .
"""""" -_'*w
30 TIME,
Figure
-,-,. NO. NO. 2 1 ENGINE
...........
angle:
sec
performance.
AIRSPEED
,
,
40
50
LANDING,
7.5 ° WING RUN
46; FLIGHT
................ ' ..... "................
'oEE
_" 60
ANGLE:
80
HANDLING
740-2B-10;
QUALITIES
3/3/82
"............ i............
LONGITUDINAL
.... TERAL
40
.o
_" 60 20
8ol-....... ....._....................... .....................
_
STICK
.._-'_PEDAL POSITION
20 ............
_
40
ATTITUDE 180 ......
!
•
40
10 -
STICK
PITCH
"................
o-_
ROLL YAW
-10
O
.= 10
_
............
-4 10
ol
-180
...............
RATE
"o
"
"o
O
_..
E-lo
rr
--5 ,.J O rr -10
"o
_
ROLL
o-_'\..,--_/-"_/,-_-_--,-_-_,---_-'_'--J-_--_-v_i_j
_ -5 < >- -10
YAW i 10
0
i 20
_ 30 TIME,
Figure
18.-
Approach
......
and
go-around,
5 ° wing
28
_ 40
i 50
sec
angle:
handling
qualities.
i 60
APPROACH
AND
GO-AROUND, RUN
53;
5 ° WING
FLIGHT
ANGLE:
740-2B-7;
PERFORMANCE
2/2/82
80
o" z
Figure
19.-
6O
Approach
and
go-around,
29
5°wing
angle:
performance.
WING INCIDENCE
Iw = 10 °
EFFECT OF COLLECTIVE
STICK
SETTING,
%
40% 35% _:3"
25%
30%
N.U
1
z ¢n F 0.,,I 20" w > _"
50 UP =_ 40
I.g --
¢/3
_,_lV_.I-:0a. D_IO " 0 o 20' ._" _o z _=¢ _ _ O (J
(3
50 UP 40 30 DN. 20
N. UP
z___
10
¢3 .
:S
_ll
_
_L
_]
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-
l-J 0
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I
dl--
i.
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_u 0,.._ _
"_
s_ou_ OgL -11--
>.
_
-..
_o_
_Z'
! (D ii
\ $_OU_ OL t dl--
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_
a.O
_
s_,OU)l0 L L dl
I
LLI _.
i-u.
.I--I_(.9
•. I .' " I / :1.'_
_l!il
ii
///
I o
I
....#/ _O
_ I i i I/#" T
slOU>l 015 BIAII"13
_
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._
a < o .J
I
i,I
,
,
_
_
,
,
,iTl O
5!sd'S]l:IfISS31dd Id3(]NrIAO
67
INCIDENCE
CHANGE
RUN 27; FLIGHT
4
_:_
LEFT LOWER _/PRESSURE
oj"_xlil
c;ol 4
RT. TF34 (#2) EXH. FRM HORIZ
4
Q
3
1-
_
I
0 80
100
I
120
[
I
140
160
CALl BRATION
Figure
60.-
I
I
I
I
100
120
140
160
]
180
80
AI RSPEED,
TF-34
71
knots
vibration.
I
180
2B-18
O
LEFT
2B-21
a []
RIGHT HAND ENGINE LEFT HAND ENGINE
/
AFTER
_
LEFT
I
TRANSDUCER
2B
22
HAND
ENGINE
HAND
WALLOPS
ENGINE F-26
CHANGE
0
0 4
0
0
o O
0
[]
n
LU
>
0 2
I-1
-
tin
d_ %0 _o n n_n
0 2O
I 30
I 40
I 50
TF34
Figure
61.-
TF-34
accessory
gearbox
qZ_ O
I 60
O
I 70
- % NF
vertical
72
vibration
versus
fan
speed.
WING O 2B I-I 2B 2B Z_ 2B-
17 18 21 22
COLLECTIVE
10 ° 0° 7.5 ° 7.5 °
SOLID SYMBOLS
40% 40% 25% A/R = 100% NR
PILOT VERTICAL
.3
.2, ¢x
mo
o_°
n_
.J a.
o o
I
0
I
l
I
I
I
+I
m 0 m COPI LOT VERTICAL
a: .3z I.I.I .,J L_J
z_
(3
'_.2
a
E] O
o °°o°°
0
.1
0
60
Figure
'
62.-
I
I
80
100 CALIBRATED
Vertical
cockpit
I
I
l
120 140 160 AIRSPEED, knots
vibration
o.."
73
versus
,I
180
calibrated
airspeed.
WING O [] A
2B-17 10° 2B-18 0° 2B-21 7.5 ° 2B-22 7.5 ° SOLID SYMBOLS
LATERAL
.3
COLLECTIVE 40% 40% 25% 25% = 100% NR
PILOT
+I
z'.2
o
ill
"J .1 iii L) O 40 1
I
I
]
I
z VERTICAL
COPILOT
FLOOR
80
•
/1
d 1.7
150
1
I
]
I
160
170
180
190
2O0
CALIBRATED
Figure
74.-
1"7
I
Frequency
AIRSPEED,
analysis,
82
2B-19,
knots
cockpit
accessories.
4.
STABILATORAND S_ABILIZER
Stabilator
Stress and Loads
General-- The Wallops Island flight tests had shown that the stabilator attachment structure (fig. 75) had marginal strength. Methods were therefore developed to determine past fatigue damageand the mechanismof loading, improve the fatigue properties of the structure, and record and process future data as accurately as possible. To this end Sikorsky was instructed to perform the above tasks, the results of which are contained in reference 2. The main requirements may be summarized as follows: i. Removethe Teflon-lined bushings and replace them with phosphorus-bronze bushings machined to obtain maximumallowable interference fit and, thereby, maximum fatigue strength. The lugs were checked for fretting (none was evident) and linedreamed to remove any possible remnant fatigue damage. The latter procedure was also applied to other appropriate areas where holes were reamed and oversized fasteners used. 2. The stabilator and backup structure were statically calibrated to permit the calculation of derived loads and stresses. Computer programs were written to obtain these data utilizing 60/main rotor digitizing rate; the data were truncated and Goodman-corrected, and the damagewas calculated at the three most significant sample rates and added algebraically as follows: 2/M + (I/3)3/M + (2/5)5/M. The above method of processing the data was somewhatmore elaborate than normally used; however, because of the randomcomplex character of the yaw waveform this procedure was required to avoid unreasonable conservatism. The primary modes of the stabilator were defined from shake tests to be as follows: Rigid body (antisymmetric): Flapwise, 6.9 Hz Edgewise, 7.3 Hz (yaw about attachment) First
mode(symmetric): Flapwise, 18.4 Hz Edgewise, not defined
During ground runs and low-speed tests, the highest loads were a function of TF-34 thrust. The waveform had a beat character with a well-defined response at about i0 Hz which was probably caused by the wake shed from the engine afterbody. At higher speeds, the wake appears to have been diverted by the free-stream flow, and other forces increased, mainly 5 per main and randomairframe induced turbulence. To reduce the former loads, a takeoff technique was developed by the Ames pilots which involved release of the brakes, then a gradual increase in TF-34 thrust. Concern had been expressed regarding the poor response of the throttles and the effect on directional control; however, this did not cause any excessive pilot workload.
83
Results-- Figure 76 shows the accumulated damage of the lugs, dragbox, and tang compared with the predicted damage extracted from reference 2. It can be seen that the rate of damage was lower than predicted, primarily a result of the change in takeoff procedure and, to a lesser extent, of operating at wing angles of less than i0 °, the effect of which will be seen in the following figures. Figures 77-80 present the derived vibratory and steady loads and stresses in the lug, dragbox, and tang. Two points are of most interest; namely, the effect of wing angle and the relatively large degree by which the working endurance level (Ew)
was
exceeded.
With respect to wing angle--since the data were truncated and Goodman-corrected, which means that only tension loads were considered damaging--steady loads were significant in controlling the damaging vibratory loads and stresses. These data show that as the wing angle was reduced from i0 °, the steady icad decreased and, consequently, the truncated vibratory levels were reduced. In addition, there was a reduction in overall vibratory levels because of a decrease in vibratory forces as the wing angle was decreased. This is shown more clearly in figure 81. Regarding the second point, the degree by which the Ew were exceeded, inadequately represents the integrity of the structure because of the random character of the excitation. As stated previously, it was decided to monitor the cycle-by-cycle damage by th_ method described. Figures 81-84 present the measured stress data, which include the stresses from which the Wallops Island data were derived. These data were monitored to trend the Ames and Wallops Island data and showed good agreement. As stated previously, there was a reduction in truncated vibratory stresses with wing angle of less than i0 ° at speeds up to 170 knots CAS.
Upper
Horizontal
Stabilizer
The attachment stresses were monitored, with levels generally remaining below the Ew throughout the flight envelope; such excesses as did occur were at a low damage rate. It was apparent, however, that at a wing angle of 0 °, the accompanying change in the aircraft attitude relative to the main-rotor-tip path caused a significant increase in attachment vibratory stress because of main-rotor-wake impingement. As shown in figure 85, the levels peaked at 130 knots and gradually decreased to almost normal levels at high speed. Since operation at this wing angle is not currently anticipated, the planned program is not affected; however, in future programs this characteristic will have to be considered.
84
TAIL
CONE
DRAG
BULKHEAD
BOX
TANG
I STA, 654.6 I UP
! BOX i DRAG
i
I
./ FWD
Y /
LUGS
LOWER
Figure
75.-
Lower
stabilator
85
attachment
structure.
STABI LATOR
25
//
20
I/ o
I
i.u"
i/i/
,_15 lE
LUG-TRUNC DRAGBOX TANG
//// //
I-Z '"10 o nr"
/ / / /
/
/
/
_/" _
/
_
/
_ _
/
/...o
I/I 0 / 4
/
/
//
Ill O.
__.-oi
//
,.o_ _
6 8 NASA-AMES
i
76.-
Predicted
and
I
I
i
12 14 18 22 COMPOUND FLIGHT HOURS
SOLID Figure
/
I
26
SYM - PREDICTED
actual
86
stabilator
structure
damage.
WING a 2B-16 o 2B-17 2B - 18 o 2B-19
COLLECTIVE
10 ° 10 °
40% 40%
• •
10 °
25%
•
WING
COLLECTIVL.
5° 5° 0o 5° 7.5 °
40% 40% 25%
O 3200 © 2800 0
0
0
I:Z _ 2400
0
[3
o
C'_ +l O
• _ 2000
o O
O>
o--[]
Ew
',' 1200 >._= n n'C3 a --
•
OO
800
400
__
100
Figure
77.-
Derived lower airspeed:
t
110
l
_
i
120 130 CALIBRATED
stabilizer rotor effect of wing
87
t
140 150 AIRSPEED,
t
160 knots
i
170
i
180
lug load vibration versus angle and collective.
calibrated
WING o
2B-
[]
2B - 18
17
2B19 2B - 21
COLLECTIVE
10 °
WING
COLLECTIVE
40%
•
5°
40%
25%
• •
0° 5°
40% 25%
•
7.5 °
25%
0 •
•
•
1000
•
•
O
O->a 0
O, 2000
0
0 _ _-'r
©
3000 []
o --> >" 4000 _w k--
©
5000
6000
......... 100
110
120
130
140
CALIBRATED Figure
78.-
Derived
lower
brated
airspeed:
stabilizer effect
150
160
AiRSPEED,
average
steady
of
angle
wing
170
180
knots rotor and
lug
load
versus
cali-
collective.
WING
COLLECTIVE
[] •
2B-18 2B-18
10" 0°
40% 40%
>X
--
_"W
[]
A•
2000 10
t _en
1000 0C
C3 0 110
I
120
__
I
130
L
140
CALIBRATED Figure
79.-
Derived
lower
stabilizer angle
AIRSPEED,
dragbox and
I
150
I
170
J
180
knots
vibration
collective.
88
160
stress:
effect
of
wing
WING [] •
28 - 17
10 °
-6 40
I
I
60
80
(HIGH) 25% 40%
Main-rotor
total
STOP
I
100
I
120
CALIBRATED
87.-
A/R
--
DROOP
Figure
COLLECTIVE
I
160
AIRSPEED,
knots
flapping
96
I
140
-- down:
I
180
effect
I
200
of
collective.
EFFECT OF WING,
4
ANGLE WING
In 2B_k 2B-
19 19
COLLECTIVE
10 ° 5°
25% 25%
Z_
in "O
0
z Q.
< .J U= I
C9 Z
2
O
DROOP STOP OUT 6
1
40
Figure
88.-
60
Main-rotor
I
I
80 100 CALIBRATED total
I
1
I
120 140 160 AIRSPEED, knots
flapping
97
-- down:
I
180
effect
I
200
of wing
angle.
WING o A O
o o [] ,,1 6
COLLECTIVE
10 '_ 10 ° 10 ° 10 ° 5° 5° 5° 5o 7.5 ') 0°
40% 35% 30% 25% 40% 35% 30% 25% 25% 40%
EW
o
O . nCc_ 4 /1
n'O
/1
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Figure
89.-
I
0
80
Main-rotor
SER_72045
100 120 140 CALIBRATED AIRSPEED,
control
loads
98
I
160 knots
versus
180
calibrated
200
airspeed.
WING [] o o O 80
Ew
COLLECTIVE
10 ° 10 ° 5° 7.5 °
40% 25% 25% 25%
= _+285 Ib
J3 n
¢3 +_. 60 Za
$; 40
SER 72045
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loo
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Figure
90.-
Main-rotor
I
80
rotating
I
I
I
I
100 120 140 CALIBRATED AIRSPEED, and
stationary airspeed.
99
160 knots
scissors
180
loads
versus
calibrated
-
.m
4°%,
_6
== +I
SE R 72045
3
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60
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80
100
120
140
160
CALIBRATED
Figure
93.-
Main-rotor
blade
outboard
AIRSPEED,
station
102
I
180
___--J
200
knots
stress:
collective
setting
25%.
WING 8 UJ IX:
_-
--e---c}---o--
7
10 ° 7.5 ° 5°
O0 (JD._
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6
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LU ..I _J
0
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100
Figure
95.-
110
i
L
i
120 130 CALIBRATED
Collective
and
i
140 150 AIRSPEED,
wing
103
J
160 2 knots
incidence
i
J
170
180
airspeed
envelope.
25%.
7OOO
BR-7
© 2B-21, W.I. = 7.5 ° COL[_ = 25 ° '_170 KCAS
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er o0
5000 .....................
LU
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o 7OOO
BR-6
o
c_
er
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m
I-D D
080
16o
1_o
CALIBRATED
Figure
ii0.-
Effect
of
wing
calibrated
incidence
on
airspeed:
115
1_o AIRSPEED,
cumulative 40%
16o
18o
knots
wing
collective.
lift
load
cells
versus
COLLECTIVE o 40%
8O0
e40%, [] 35%
100%NR
I 35%, A 30% 0 25%
100%
NR
VIBRATORY
+I
150 uJr_
_ lOO ='_- so z 0 o
0
I
I
I
I
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J 140
i 160
) 180
STEADY
25O J_ m
¢ 160 knots CAS) 4800-ft density altitude, maximum 57- to 181-knot CASlevel flight airspeeds 7 knots maximumcrosswind componentfor takeoff and landing The envelope is further defined by the rotor speed, wing-flap angle, maneuvering, and V-N diagrams excerpted from reference 1 and presented in figures BI-B3. The demonstrated limits for adjustment of collective position are presented in figure B4 from data obtained from reference I. The envelope for safe ejection from the RSRAusing the EESis presented in figures B5 and B6. The evaluation pilot's station is most critical. FLIGHTSUPPORT FACILITIES The following
support facilities
are required.
I.
NASAAmesResearch Center, Moffett
Field,
2.
Chase/rescue helicopter
outside the field
for flight
3. Fixed-wing chase for flights Envelope maximumspeed is 181 knots)
flight
facility boundary
on which speeds exceed 120 knots CAS. (Note:
4.
Fire truck on station
from engine start
5.
Ground operations video tape coverage
6.
Airborne cinematic and still
to engine shutdown
Fhotographic coverage (when requested)
2O4
DATAACQUISITION,PROCESSING, ANDTELEMETRY Data Acquisition Measurementdata will be recorded on board by the RSRAPADSon two tapes. The PADSrecording capacity is 120 FMand 104 PCMmeasurements. The parameters to be recorded are listed in table B3. Note: table B3 is an initial list only and will be modified, summarily, by the test director, based on test requirements. Specifically, priority II and III parameters will be substituted for priority I tail parameters in accordance with reference 2. Data Processing Data will be processed as required by the Playback Schedule, issued i day before the flight, and the tape log (emery sheet). The FMdata will be filtered as required by the emery sheet before digitizing. Most FM data will require peak and hold processing as specified by the emery sheet, that is, l/T, l/M, and 5/M. PCMsampling rate is determined by the PADSsetup and is 80/sec. Oscillogram and PCMstrip charts will be required immediately after (i day) to check recorded data quality and to analyze waveform.
each flight
The EASEprogram will be used to process digitized data to provide DA (data analysis) task tabulations of data within 2 working days following the flight. The DA layout will be specified by the test director about i week in advance of the test. The EASEprogram will also be used to provide cycle count tabulations of parameters with vibratory concern levels. These concern levels will also be provided i week before the test. Postflight
frequency or harmonic data analysis will
be required,
upon request.
Telemetry (TM) Up to i0 FMparameters and the PCMwill be telemetered for real-time monitoring. The TMsetup and parameters to be monitored on oscillogram and Brush recorder output will be specified by the test director. Frequency analysis of TM data may be required to troubleshoot vibration problems. The Hewlett Packard 5423A analyzer with plotter will be required in standby condition to be switched upon request to any parameter on TM. The parameter to be telemetered will be specified by the test director based on test requirements. The aircraft
PADSTM shall be switchable to any FM track. Preflight
Action i.
Flight
2.
Data
require
request review
flight)
to
d FO
(previous
and Postflight When
and
Week
before
Prior
to
briefing
205
Data Sequence
hours a
or
By Test
at
whom
director
Test director, Smart neers, pilots
Book
engi-
Action Playback setup (DP)
,
Emery
o
schedule and TM to data processing
sheet
to DP
5.
Test
6.
PADS preflight and (before rollout)
card
Tape
.
When
required
to
review
A/C,
setup
for
test
tape
instrument
and
hours
By whom
-24
Test
-24
Instrument
-24
Test
-24
Instrument
engineer
and
technician
Instrument
engineer
and
technician
-3
1/4
Briefing
-2
9.
Man
-i
3/4
Test
-i
1/2
A/F
TM
Send
A/C
and
engineer
director
flight
8.
i0.
director
calibrations
to
TM
3/4 (or before)
day
Test director, pilots, TM A/C managers, instrument neers, support crews director,
TM
flight-test
team, engi-
team
engineer
tape
ii.
Pilots
to A/C
12.
Conduct
flight
-3/4
Pilots
0
Pilots,
TM
support 13.
Debrief
+1/2
Same
14.
Tape
+i
Instrument
15.
Flight
+3
Data
aide
16.
Pick
+24
Data
aide
17.
Scan stripouts crab sheets
+32
Flight-test assigned)
18.
Pick
+48
Data
19.
Review
DA
+54
20.
Update
Smart
+72
flight
_l_inus signs completion.
to DP log up
up
to DP
stripouts
DA
and
initial
output and
crab
sheets
Book
designate
hours
before
flight,
206
team,
test
director,
crews
as
(7) engineer
engineers
(Smart
Book
Flight test assigned)
engineers
(Smart
Book
Flight-test assigned)
engineers
(Smart
Book
plus
aide
signs
designate
hours
after
TABLEBI.Brake-on velocity,
BRAKECOOLING SCHEDULE knots
Changein temperature,
20 30 40 5O 60 70 8O 90 i00 Ii0 120
240 280 330 400 5OO 600 720 88O 1080 1240 1400
207
°F
TABLEB2.- COOLING TIME IN MINUTES Brake temperature, 200 250 300 350 400 450 500 55O 600 650 700 750 8OO 85O 900 950 i000 1050 ii00 1150 1200 1250 1300 1350
°F
To
200°F 0 13 26 38 47 56 62 68 75 81 87 93 I00 106 112 118 125 131 137 143 149
To
400°F 0
To
600°F 0
To
800°F 0
To
1000°F 0
6 12 19 25 31 36 41 45 49 53 57 60 63 66 70 73 76 79 82
8 12 16 20 23 25 28 30 32 35 37 39 41 44 46
5 7 i0 12 14 16
1 2
18 20 22 24 26
3 5 6 7 8
9 i0 ii 13 14
15 16 17 18 19
Caution 1400 1450 1500 1550 1600
91 94
48 5O 52 55
97
57
28 29 31 33 35
59 61 64
37 39 41
66 68
43 45
85 88 m
Danger 1650 1700 1750 1800 1850
i00 103 107 ii0 113
Notes: Cooling times are not linear. Tables are based on a gross weight of 27,000 ib and an engine thrust of i000 lb. "Caution" means to suspend all braking operations; allow brakes to cool a minimum of 1.5 hr. "Danger" means to avoid exposing personnel to the wheel because of danger of explosion.
208
TABLEB3.- PADSINSTRUMENTATION LIST FM Parameters MRpitch MR flap MRlag MRshaft bending MRdamper moment MR
blade
rear
station
6
MR MR
push rod blade rear
station
7
MR MR
rotating scissors stationary scissors
MR MR
right lateral contactor
Left landing Right landing
stationary
load
Longitudinal vibration, left wing tip Vertical vibration, tail-rotor gearbox Lateral vibration, tail-rotor gearbox Vertical cockpit load factor Lateral cockpit load factor Lower stabilizer lug stress Lower stabilizer lug stress Lower stabilizer lug stress Left-aileron control rod load Right-aileron control rod load TR blade normal bending right TR contactor
balance balance balance
MR balance MR balance Wing, left Wing, left Wing, left Wing, right
lift lift lift
lug stress lug stress drag box stress cell, forward cell, aft link load
cell, cell, cell,
X
X
MRFLAP
X
X
MRLAG MRSBLI MRDMPOI MRBR6 PRLREDUP MRBR7 MRROTSC MRSTASC MRRLSS
X X X X X X X X X
X X X X X X X X X
X X X
X X X X X X X X X X X X X
FLIGHT
70
-.0_
_ _
"q
t-
_w
40
-,___
__J DOWN O
30 I
20
90
I
I
I
0 LEVEL FLIGHT POINTS ACLIMBING AND DESCENDING
80
z"
G
I-ITURNING
FLIGHT
I
I
I
POINTS
POINTS
70 /_
0 UP t-
LEVEL
FLIGHT
ENVELOPE
60
O
50
LU
> (.) 1.1.1 _1 DOWN _1
40
0
30
-
"l '_'_'_"
A
f/ENVELOPE
"_"_
20 I
10
I
PARTIAL
I'
I
_u_
AND
POWER DESCENTS
I
ENVELOPE
70
FOR CLIMBS
I
I
FOR TURNS
_NWLO_
_._:__-_,
LEVEL
FLIGHT
\ _ DOWN40, '-' 30 ¢J 50
Figure
. i 70
B4.-
I 90
Collective
_ = 110 CALIBRATED
position
217
I_] i 130 AIRSPEED,
versus
150
I
I
170
190
knots
calibrated
airspeed.
MIN
ESCAPE
ALTITUDE
500
COMPOUND
MODE
EVALUATION
PILOT
400
300
