1971025733
NASA TECHNICAL NASATM X-62,083 MEMORANDUM
O3 00 O
< z
A FLIGHT EVALUATION OF A VTOL JET TRANSPORT UNDER VISUAL AND SIMULATED INSTRUMENT CONDITIONS
Curt A. Holzhauser, Samuel A. Morello, Robert C. Innis, and James M. Fatten
Ames Reseach Center Moffett Field, Calif. 94035
Langley Research Center Hampton, Va. 23365 and
(ACCESSI R) _._ 1971025733-002
A FLIG}IT EVALUATION OF A VTOL JET TRANSPORT UNDER VISUAL AND SIMULATED INSTRUMENT CONDITIONS
By
Curt A. Holzhauser*, Samuel A. Morello**, Robert C. Innis *, and James M. Patton**
.,."2,
*Ames Research Center **Langley Research Center
_. '_
ii " ]97]025733-003
TABLE OF CONTENTS
SUMMARY Page INTRODUCTION NOTATION ...... i
DESCRIPTION OF AIRPLANE AND EQUIPMENT ...... 5 TEST PROCEDURES AND CONDITIONS ...... ii RESULTS AND DISCUSSION ...... 14
Performance and Basic Operating Procedures ...... ],5 Low speed operational envelope ...... 15 Vertical takeoff and transition ...... 17 Approach and vertical landing ...... 18 Handling qualities ...... 20 Conversion ...... 21 ILS acquisition and tracking, longitudinal flight path control ...... 21 Final transition and vertical landing, longitudinal and height control ...... 27 ILS tracking, lateral-directional flight path control . 29 Final transition and vertical landing, lateral control. 31 Trim considerations in transition and hover ...... 36
Terminal area operation ...... 37 Cruise letdown to pro-approach configuration ..... 37 Conversion ...... 37 ILS Acquisition ...... 38 ILS Tracking ...... 40 Final transition and vertical landin_ ...... 42 Low speed translation ...... 42 Environmental effects ae transits,on _peeds ...... 43 Environmental effect8 at very low speeds ...... 44 Comparison of approaches ...... 46 CONCLUDING REMARKS ...... 48
APPENDIX A.- Computation of data pertaining to flight path deviations ...... 51 APPENDIX B.- Miscellaneous engine and control relations .... 56 REFERENCES ...... , ...... 57 .i _ TABLE I.- ...... 59 1971025733-004
SUMMARY
A flight investigation was performed with the Dornier DO 31 VTOL transport to evallate the performance, handling, and operating characteristics that are considered to be important when operating a commercial VTOL transport in the terminal area. The DO 31, a
20,000-kilogram transport, has a mixed jet-propulsion system; i.e., there are main engines with nozzles that deflect from a cruise to a hover position_ and vertical lift engines that operate below 170 knots.
Inarethisincorporated.VTOL mode, pitchThe mainandandrollliftattitudeenginesandareyawusedrateto s control• "_ ð
VTOL forces and moments.
The tests concentrated on the transition, approach, and vertical landing. The mixed Jet-propulsion system provided a large usable performance envelope that enabled simulated IFR approaches to be made
.." 2, on 7° and 12 ° glideslopes. In these approaches management of thrust magnitude and direction was a primary problem, and some form of integrating the controls will be necessary. The handling qualities evaluation pointed out the need for additional research to define flight-path criteria.
The aircraft had satisfactory control and stabillty in hover out of ground effect. The recirculation effects in a vertical landing were large below 15 meters. _
l 1971025733-005
INTRODUCTION
Commei_cial V/STOL aircraft offer the possibility of overcoming many of the shortcomings of present short-haul air travel. The low- speed characteristics make it possible to operate from small airfields which can be conveniently located near the centers of population.
Additionally, these characteristics should red_ air .andground maneuver time, and improve reliability under adverse weather conditions
(refs. 1 to 5). Although considerable research and development have been done over the past decade in the United States and abroad in studying the performance, handling quallties, and operating character- istics of different types of V/STOL aircraft (refs. 6 to 14), it has been difficult to realistically assess the potential of commercial
V/STOL transport aircraft and to define the desired characteristics, particularly fo'_'IFR conditions. Each aircraft tested had limitations either due to size, stability and control, inherent characteristics, or inability to represent IFR flight; consequently, there is a continuing need to update available informat].on by terminal=area tests with V/STOL aircraft that permit a better simulation of terminal-area operation.
A flight evaluation ,_asmade with the Dornier DO 31 Jet VTOL transport because this aircraft has several features that offered a better assessment of the terminal-area operation than other research aircraft tested. First, it is sufficiently large (20,000 kg) to
i
i 1971025733-006
represent a first generation transport. Second, it has a mixed propulsion system (main fan-jets with vectoring nozzles plus lift-jet engines) that provides a very broad performance envelope. Third, it has an advanced control and stabilization system that can reduce pilot workload. Fourth, the controls and displays are duplicated so that IFR operation can be simulated. The NASA flight tests were primarily concentrated on the transition, approach, and vertical landing phases of operation since these are generally considered to be the most demanding phases in terms of aircraft performance and handling qualities.
The tests were conducted on 7° and 12 ° glideslopes with some simt_ated
IFR operation.
The tests were conducted by NASA personnel from Ames and Langley
Research Centers in cooperation with the Dornier Company,
Bundesministerium fur_#issenshaft und Forschung (BWF), Bundesministerium fur Verteidigung (BMVg), and the Deutsche Forschungs und Versuchsanstalt f_r Luft und Raumfahrt (DFVLR). 1971025733-007
NOTATION
Ax Longitudinal acceleration of center of gravity as measured
1 dVx by an accelerometer, sin @ + --_ g g dt' dvx ax Longitudinal acceleration of aircraft, d-_'m/s2
Ay Lateral acceleration of center of gravity as measured by
an accelerometer, sin _ + _ dry g dt' g ay Lateral acceleration of aircraft, _ dr' m/s 2
A z Normal acceleration of center of gravity as measured by an
1 dVz accelerometer_cos @ + --g"--dt'g
dVz m/s2 a z Normal acceleration of aircraft, d--T'
CD Drag coefficient, including propulsive thrust
CLlg Lift coefficient in steady-state flight, including propulsive thrust
CL_ Power-off lift curve slope, per deg FCUL Position of fuel control unit for left lift engines, deg
FCUR Position of fuel control unit for right lift engines, deg
g Acceleration of gravity, 9.81 m/s 2
h Height above the runway, m 1971025733-008
- 2-
lyy Moments of inertia, Kg-m 2
Izz Ixx I
L Rolling moment, newton - m m mass, kg
N1 Speed of lift engine furthest forward in left pod, rpm
N 5 Speed of lift engine furthest forward in right pod, rpm
N_ Main engine fan speed, measured for left engine in percent
of maximum speed, percent q_ Free stream d_namic pressure, Kg/m 2
R/C Rate of climb, m/s
R/S Rate of sink, m/s s Horizontal distance, m or km
S Wing area, m 2 t Time, s
T Thrust, newtons
TA Ambient air temperature, °C
T1 Average temperature at main engine inlet, °C
V True airspeed, knots or m/s
VC Calibrated airspeed, V_, knots
Vy Velocities in body axes, m/s
VZ _"/
L 1971025733-009
- 3 -
W Weight, ne_._ons
X Longitudinal displacement, m
Y Lateral displacement, m a Uncorrected angle of attack measured at nose boom, deg
6 Uncorrected angle of sideslip measured at nose boom, deg y Flight path angle (climb, positive), deg
TG Glideslope angle at centerline of ILS (descent, negative),
deg
_A Left aileron deflection (trailing-edge down, positive), deg
_E Elevator deflection (traillng-edge down, positive), deg
_F Flap deflection (trailing-edge do_m, positive), deg
6Lp Lateral stick deflection (right, positive), deg
_MP Longitudinal stick deflection (aft, positive), deg
_NP Rudder pedal deflection (right pedal forward, positive), mm
_PN Pitch nozzle deflection (nose-up pitching moment, positive),
deg
6R Rudder deflection (trailing-edge left, positive), deg
eG Glideslope error (above, positive), dots or deg
eL Localizer error (to right, positive), dots or deg
_M Pitch stabilization actuator position (nose-up pitching
moment, positive), deg
_L Roll stabilization actuator position (right rolling moment,
positive), deg 1971025733-010
- 4 -
_N Yaw st_)ilization actuator position (right yawing moment,
positive), deg
_VTOL VTOL roll rate damper actuator position (right roll rate,
positive), deg
e Pitch attitude (nose up, positive), deg
eTRIM Pitch trim position (nose up, positive), deg
Pitch rate (nose up, posztl.e),• 'v deg/s
Density ratio
GFCU Lift engine thrust lever position, deg
_M Main engine nozzle lever position, deg
GNL Nozzle deflection for left llft engines (aft,'positive),
deg
_NR Nozzle deflection for right lift engines (forward positive),
deg
_T Main engine thrust lever, measured for left engine, deg ..
Bank angle (right wing down, positive), deg
$ Roll rate (right wing down, positive), deg/s
Angular acceleration in roll (right wing down, positive),
rad/s 2
Heading angle, from measured clockwise, true north, deg
Yaw rate (nose right, positive),deg/s 1971025733-011
-5-
DESCPIPTION OF AIRPLANE AND EQUIPMENT
The DO 31-E3 is a high-wing, mixed-propulsion, Jet V/STOL transport
with two main engines (vectored lift-cruise) and eight lift engines.
The aircraft was designed and constructed by the Dornier Company for a
V/STOL research program which was initiated in 1962 and sponsored by
the German Federal Ministry of Defense. Figure l(a) is a photograph of
the airplane in the VTOL mode. A three-vlew drawing is given in
Figur,_:l(b) with additional details in Table 1. The first flight of
the aircraft was in 1967, and it was followed by 24 hotu_s of flight
tests primarily to define the operational envelope and document the
performance. The subsequent NASA program for ll flight hours
primarily evaluated and documented handling qualities in the W/STOL
mode and simulated,IFR operation. The normal operating mass of the
airclaft was about 19,500 kilograms (43,000 pounds).
Propulsion
The E-3 aircraft is equipped with two Rolls Royce (Bristol % Division) Pegasus 5-2 turbofan engines and eight Rolls Royce RB-16_-4D
lift-Jet engines. The two Pegasus 5-2 engines are mounted under the
wing and hav,s nozzles to vector the thrust. Each engine has four
nozzles _hat vector the total thrust forcc from a lO° thrusting to a
120 ° braking position. This engine and nozzle arrangement is
_ essentially the same as used on the Hawker-S_ddeley P.I127 aircraft.
_ • Each Pegasus 5-2 engine is_rated at 69,000 newtons (15,500 pounds force) 1971025733-012
- 6 -
of uninstalled sea-level static thrust. At each wing tip are four
RB-162-4D lift engines housed in pods. Each RB-162-4D engine is rated
at 19,6C0 newtons (4,400 pounds force) of uninstalled thrust at sea
level.
Flight Controls
Figure 2 is a schematic of the separate control functions, and
figure 3 is a photograph showing the cockpit layout. The flight
attitude controls are a stick and rudder pedals. In cruise the
controls are linked to conventional ailerons, elevator, and rudder.
In hover, rolling moments are produced by differential thrust
between left and right sets of lift engines. The thrust is commanded
by the fuel control units (FCU) which are linked in each pod to the
stabilization system and the stick. The pitching moments are produced
_y reaction controls located at the aft end of the fuselage; high
pressure air is supplied from each main engine through separate ducts :'"'
and nozzles. Yawing moments are created from fore and aft movements
of nozzles on the tail pipes of the lift engines. As with roll control,
the pitching and yawing motions can be controlled by either the pilot ......
or stabilization system. In the transition the moments are produced
by a combination of the hover and conventional controls because the j
latter move at all speeds. In addition to the stick and rudder pedals,
one set of main engine throttles, two main engine nozzle control levers
_ (one for each pilot), and one lift engine thrust lever are used for
.1 flight control and are located in the center console. 1971025733-013
-7-
The DO 33 flight control system is powered by dual hydraulic
actuators with control rods_ summing bars, etc. The relation between
the control forces and deflections are given in figure 4. The stick
deflection in millimeters is given on the secondary scale of the
deflection in degrees. It is of interest to note that this transport
has a stick rather than a wheel, and also that the lateral control
motion is obtained by movement about a pivot near the center of the
stick (figure 3(a)). The pilots found the use of a stick acceptable,
and in fact was preferable to a wheel for a transport VTOL because
there was less obstruction to the view_ and it was more natural with
the one-hand method of control required during transition and hover.
The forces and deflections were quite satisfactory.
Figures 5(a) and (b) relate the throttle and engine characteristics.
The lift engines are started together with the lever at 17 ° FCU; after
about l0 seconds a stable subidle is achieved and the individual
: warning lights go out. Then the lever is advanced to 30° , and
lO seconds later a stable flight idle is attained as indicated by
another set of lights going off. The forces and deflections of the
engine levers were satisfactory except for the fact that it was
unsatisfactory to have the height control in the VTOL mode split
between the main and lift engines; one control combining the two
functions would be desirable. The main engine nozzle control was
also satisfactory, but the deflection indicator on the panel had to
_ be monitored and a better display was warranted.
"ri 1971025733-014
- 8 -
Stabilization System
The aircraft is equipped with a full authority single channel attitude command control system for the pitch and roll axes and rate command for the yaw axis. Figure 6 presents the block diagrams of the
stabilization and control system of each axis in the VTOL mode.
Figure 7 is a schematic of the stabilization EJstem for each axis. The
pitch and roll attitude stabilization system compares the commanded
attitude from the control stick signal to the actual aircraft attitude
derived from the attitude gyro signal. In the yaw axis, rate is
compared rather than attitude. These _rror signals are then used
through the servo-motors to drive the aircraft to the commanded
steady-state conditions shown in figure 8. The control signals are
introduced additively through a mechanical linkage; thus, in the event
of a stabilization system failure, the control reverts to a direct
mechanical control immediately. If the control is deflected beyond ....._. / the position of the limit switch (figure 8), the stabilization system
is disengaged to provide safety in the event of a "hard over" failure.
The pilots noted a reluctance to use large control deflections
because of fear of disengaging the stabilization system. They
considered _his method of disengaging the system unsatisfactory. A
roll damper is incorporated to improve the lateral controllability
throughout all flight modes. In the VTOL mode, trim is provided in
pitch only. The pilot can control the pitch trim in two manners; _ -';
one, by a trim switch ol_ his stick, and two, by a preselect switch ! 1971025733-015
-9-
located on the instr_.nnentpanel (figure 3(a)). For the latter case,
he can dial in the desired pitch a_±_ at an_ _.m_, ana_.by pressing
a button on the sti_;k the attitude will change to the preselected
value at the rate of 3° per second, The preselect trim system was a
desirable festure, but the panel mounted switch was somewhat awkward
to use. in the conventional mode, trim is provided for each control.
Co_kplt Instrumentation and Displays
Figure 3(b) illustrates the arrangement of the cockpit instruments
and displays fo_- the evaluation pilot. Glideslope and localizer error
information was displayed on the attitude director indicator (ADI); no
steering information was used. True airspeed was obtained from the
"Fluglog" (a free-turning, self-alining propeller utilizing the
anemometer principle with optical pickups to sense rpm) developed by
Dornier and mounted on the end of the nose boom. The face of the
_ standard production airspeed indicator was changed to display 5-knot
_ increments of true airspeed. Angle of attack and sideslip were taken
_,_ from the deflection of the fluglog and displayed to the pilot.
Data Acquisition The airborne equipment was capable of registering 208 different
data channels simultaneously. One portion of the data was stored in
analog form on magnetic tape onboard the aircraft, and at the same
time transmitted _o a ground station onto magnetic tape. The remaining
data were sampled azd then stored in digital form on a tape recorder in
! ! 1971025733-016
- i0 -
the aircraft. Safety of flight information was telemetered to a ground
station to be recorded and monitored during the flights. Computers
were used to reduce the data to engineering units on plots and tabulated
printouts for data analysis.
A ground-based radar operated by DFVLR was used to obtain the
position of the aircraft during the approach and landing phase of the
_Sights. The measurements were printed out at 1-second intervals and
were time correlated with airborne data.
Guidance
_idance for the instrument approaches was provided by ILS
equipment based at the airfield and operated by DFVLR. This system
provided a wide range of glidepath angles and sensitivities.
Figure £ illustrates the profiles and sensitivities used during this
investigation. The semi-beamwidth provided a full-scale deflection
of three dots on the ADI. No other approach or navigational aids /
were available. The glideslope transmitter was located adjacent to
the VTOL landing area, see figure 9.
4
f 1971025733-017
- ll -
TEST PROCNDURES AND CONDITIONS
Test Location
All of the tests were made at the Dornier Flugplatz in the outskirts
of Munich, Germany. The field elevation was 600 meters and the
temperature ranged between 0° and 12 ° C. The flights were made under
visual flight rules (VFR) over a range of wind conditions. The wind
speeds, measured near the ground with an anemomete_ ranged from
0 to l0 meters per second and included headwind, crosswind, and
tailwind. For all flights, the Dornier pilot was in command in the
left seat.
Hover-Rig Tests
A hover rig simulating the VTOL mode of the DO 31 was used for
pilot checkout and training, This rig, shown on the pedestal in
figure 10, was also flown in free flight over a range of speeds up
_" to 40 knots (forward and sidewards) and a range of altitudes up to
_ 100 meters above the runway. The tests were limited to a 5-minute g, _r7, duration by the fuel available and the continuous lift engine
_::. rum_ing time. The rig was similar to the DO 31 in terms of the VTOL
"_i,J, propulsion, control, and stabilization systems, and it had similar _T response. It had three lift engines in each pod rather than four, ¢ and the mass and inertia were lower. When mounted on the pedastml
in its raised, operating position, the hover rig had restricted
i angular movement and no vertical movement. It was very useful
because the pilots could evaluate the angular response in hover _
f2 e_ l .,_# '1 _-L 1971025733-018
- 12
with and without stabilization and also the response due to shutting
down a lift engine or main engine. After a few pedestal runs, each
pilot made several free-flight tests. These free-flight tests
included tests with the stabilization system turned off, but did not
include engine failures.
DO 31-E3 Tests
All flight tests were within the operational envelope established
by the manufacturer. Engine failures were not simulated or performed.
All tests were made with the stabilization system engaged, except for
limited tests with the yaw rate stabilization disengaged during the
approach.
Most of the tests were started with a conventional takeoff at a
mass of about 21,500 kg (47,000 pounds), and consisted of three to
five approaches, terminating at 70 meters of altitude either in a
waveoff at 50 knots or a hover and a vertical landing at a mass of /_
about 18,500 kg (hl,000 pounds). These procedures were used
primarily to maximize research time with the limited lift engine
time and fuel. The lift engine time was limited to 5 minutes per
start by the simple oll system, and each flight was about 20 minutes
total. Several vertical takeoffs were also made. One flight was
devoted to a climb to 3,000 meters to document and evaluate the use of the main engine thrust deflection for rapid descent and deceleration
from cruise altitude and speed. The total of ii flight hours was _'"
e@_allF divided between the Langley and Ames pilots. This £1ight time
permitted 90 approaches to be made of which 40 simulated IFR operation. .ff 1971025733-019
- 13 -
The glideslope angles and sensitivities were varied during the
program. Tests were made primarily with a-7 ° glideslope with a beam
width of _2 °. Limited tests were made with-7 ° ±l ° and with-12 ° ±2 ° .
The variations in flight profile tested are tabulated below:
Glideslope, Intercept Altitude, Lift engine starting deg meters (feet) condition
-7 600 (2,000) Libel flight
u7 450 (1,500) Level flight
7 300 (1,000) Level flight
-- 7 600 (2,000) On glideslope
- 7 450 (1,500) On base leg in turning flight
-- 12 600 (2,000) Level flight
-- 12 450 (1,500) Level flight
--12 900 (3,000) On glideslope
During the majority of these tests, the location of the aircraft was
_ recorded and correlated with onboard measurements.
: STOL tests were not performed in these NASA tests primarily
because of the high risk due to the nature of the runway. The entire
runway, except for the VTOL landing area (figure 9), was surfaced
with high friction asphalt that eroded rapidly by hot gases emitted
_ by the lift engines and also by the main engines when they were
_ deflected downward. This erosion was of concern not only because it
could damage the runway, but also because of potential damage to the
DO 31 engines. 1971025733-020
- 14 -
RESULTS AND DISCUSSION
To fully realize the commercial potential of VTOL transports, it will be required to operate routinely to low visibility minimums under Instrument
FliBht Rules (IFR) and to minimize ground and air maneuver time, fuel, airspace, and noise. These requirements clearly indicate the landing approach to be the most critical flight condition for these aircraft. Further, previous NASA re- search has shown the approach phase to be the most demanding in terms of pilot workload and has indicated that a number of unresolved questions exist. There- fore, the NASA flight tests ol the DO 31 concentrated on evaluating the per- formance, handling and operating characteristics of this mixed Jet-propulsion concept with an advanced stabilization system in simulated IFR approaches.
It must be recognized that the flight test time of the program was limited, and included the time required to familiarize the 2 NASA pilots. Consequently, the operating procedures and patterns used were primarily those developed by
Dornier personnel, and the documentation and evaluation of handling qualities was limited. For these reasons the pilot comments are given in an adjective _ and commentary form rather than in a quantitative form.
The section entitled "Performance and Test Procedures" contains static climb and descent characteristics supplied by Dornier to describe the operating envelope, and also time histories that illustrate a typical vertical takeoff and transition to conventional flight and a transition to VTOL configuration, approach, and vertical landing. The 'IHandli_g Qualitles" section contains the measured characteristics to support the NASA pilots' evaluation of work- load in approach and landing. The last se_ion "Terminal Area" presents pri- marily the results of complete approaches. There are also some re- sults of translating near hover, and a simplified comparison of different
approach and landing techniques in terms of airspace and time used. 1971025733-021
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Performance and Basic Operating Procedures
L0w speed operational envelope.' The operational envelope for the
DO 31 is illustrated in figure ii as climb and descent for unaccelerated
flight versus airspeed. Included in these figures are lines of i0 and 20
degrees climbing and descending flight paths in unaccelerated flight. Since
the majority of the tests were in accelerating or decelerating flight, it
should be recognized that these lines also approximate 0.17 and C.35 acceler-
! dV - CD . First_ ation and deceleration in level flight; i.e. _ tad +g (_) = _ig examining the conventional mode (lift engines inoperative), figure ii (a),
it can be seen thst with the two main engines operating at high power set-
ting, an extremely large range of flight paths can be obtained by deflec-
ting the nozzles of the main engines from I0 ° (thrusting) to 120° (braking).
Buffet occurs with nozzle deflections greater than 85 ° at the higher air-
speed; however, more than sufficient descent performance is provided for
a rapid let down° Although 20 to 30 knot reductions in stalling speeds
were achieved over the power off value, the operational speed could not -_
be reduced because of the minimum control speed requirement, VMC ' The
VMC was defined with the nozzles at i0°, and it was limited by the directional
control of the conventional rudder. For this conditio_^a rate of climb of 5 meters per second (i000 feet per minute) was obtained.
In the VTOL mode (lift engines operating), a very lar_r range of oper-
ation can be obtained (figure ll(b)). The curves shown are for all engines
running and are illustrative of the configurations used during NASA tests.
The range of lift engine throttle settings shown provide sufficient range
_. for lateral control, for a margin above the flight idle setting, and to 1971025733-022
- 16 -
compensate for engine failure. The curve with_Fc U = 40 ° , N F = 80%,_M = 120 °
is about the maximum descent capability of the DO 31. The main engine speed,
NF, was not reduced below 80 percent so that sufficient bleed air could
be provided for pitch control. Steady descent rates below 50 knot airspeed
were not defined, but it can be presumed that instantaneous values greater
than i0 meters per second are attainable. In hover, the maximum descent
rate is dictated by the landing gear touchdown design speed of 4 meters
per second. The wave off case is with a typical approach power setting
and the nozzle deflection reduced from 120° to 70° . The takeoff case is
with reducing nozzle deflection as speed increases, (artificially presented
at 0 acceleration for comparitive purposes). Obviously, different perfor-
mance curves can be established for different power settings, nozzle de-
flection, and angle of attack, dependent on the desired feature to be op-
timized. The maximum airspeed at which the llft engines have been operated
is 170 knots; normally, the maxi_,m airspeed was 160 knots to have a me.rgln _
w and to avoid instability of the £ixed-gain stabilization system. Although
_- it is not shown in the figure, a positive climb can be achieved over the
entire speed range with one lift engine inoperative; and over a large part
of the range with a main engine inoperative.
The thrust-weight ratios available in hover, out of ground e£fect
are given in figure 12 for an aizcraft mass of 18,500 kilogram _nd a field
_. elevation of 600 meters. Three curves are shown: the upper curve is with
all lift engines operating at maximum continuous thrust and both main engines
at 2-1/2 minute ratiugt the middle and lower curves show the e£fect of a
lift engine failure and amain engiu_ failure, respectively; both curves
I 1971025733-023
- 17 -
are with lateral symmetry maintained. It should be noted that the vertical
force can be further increased by utilizzng a main engine emergency thruDt
which increases the thrust-weight ratio by as much as 0.05 (dependent on
the bleed and control used). These curves illustrate the magnitude o =
thrust-weight ratio that is installed to compensate for an engine failure
and that might bc available to develop normal acceleration for maneuvering
near hover out of ground effect. It is seen that the effect of a lift en-
gine failure is smail compared to a main engine failure; however, even in
the latter case the aircraft can be ballanced and a thrust-welght ratio
in excess of one can be developed.
The proximity of the ground _h < 15 meters) was estimated to reduce
the vertical force by about I0 percent; this reduction was caused by re .....
circulation and re-ingestion of gases into the main e:gines.
Vertical takeoff and transition: The takeoff performance and proce-
dures are illustrated in figure 13. Once the llft engines have been started _
it is necessary to proceed rapidly with the takeoff for two reasons;
i) the idle thze_t is so high that the aircraft is very light on the gear,
and 2) it J s desirable to move away from the large hot gas cloud developing.
A nozzle deflection of 75° is used for takeoff to minimize the reclreulation
effects. Even though the main engine nozzles and the lift engines _re now
both deflected 15° aft of the vertic&l, there is practically no grc_nd roll
in the takeoff because of the large thrust*to-weight ratio applied to take-
_ off. The result is a steep, high acceleration takeo£f and transition with
an average acceleration of more than 0.28's. In Just over 20 seconds suf-
ficient airspeed is attained to shut off the lift engines. A steep climbout 1971025733-024
- 18-
can be continued because of the high thrust-weight ratio of the main engines
i_I_ 0.6).
Approach and vertical landing: The procedures and flight paths used
for two different approaches to a vertical landing are illustrated in figure
14. The three primary phases of the approach for this VTOL are:
I) The conversion from conventional mode to the VTOL mode;
2) The initial transition where the aircraft is decelerated from
about 150 knots to 50 knots during which time a precision approach
is made;
3) The final transition from 50 knots to hover at 60 meters followed
by a vertical landing.
In the following discussion of handling qualities the longitudinal aspects
of these three phases will be discussed first; _L__L[L_will be followed by the
lateral-directional characteristics.
In order to complete these three phases of the approach with repeatable
precision and with a reasonably low pilot workload, it was desirable to fly i the aircraft with prescribed discrete operations that the pilot could per-
form at selected locations on the path. In addition, the recommended pro-
cedure was to track the glide slope at a negative angle of attack (near
=: zero aerodynamic lift); the engines were then near a hover setting so that
a hover could be attained with no change in main engine thro_tle setting ,o
and _nly a small adjustment to the lift engine throttle setting. Approaches
_ were made where the lift ensines were started before the slide slope was =
_ acquired, and also where the lift engines were started after the slide _
slope was acquired4 _!_
4'
r 1971025733025
- 19 -
The first phase of the approach, the conversion, starts with the pilot
establishing the pre-conversion configuration; ioe., the gear and flaps are
lowered, and the airspeed is reduced to about 140 knots; the lift engine
pod doors are opened; the attitude stabilization system is engaged; and the
desired pitch attitude is pre-selected by the pilot. The airspeed has to
be lowered to 140 knots to avoid instability of the stabilization system
caused by a fixed gain system which was designed for hover, and also to
avoid high starting RPM's of the lift engines. Next, the lift engines
are started and advanced from a sub-idle to an idle speed _ FCU = 30);
this cycle automatically starts all 8 engines, and requires about 20 seconds
to obtain a stable idle. During this period, the pre-select trim button
is depressed to change the pitch attitude so that the added vertical force
of the idle lift engines (T/Widie = 0.35) is compensated by reducing the
wing lift to minimize "ballooning", and the nozzles are deflected to main-
tain constant airspeed.
During the second phase, the precision approach, corrections were
primarily made by modulating the lift engines with the aircraft stabilized
at the selected pitch attitude. The main engine nozzle deflection for
this phase was generally at 120 ° (maximum braking) to provide the desired
deceleration schedule; in some cases smaller values were selected to adjust
for head winds. When an altitude of about 60 meters was reached, the third
and final phase _f the approach commenced by rotating the aircraft to +5 °
attitude (through the pitch attitude pre-select trim system), by changing "-.,
the nozzle to 95 °, and making small lift engine corrections to attain a Z/
stable hover. Then the lift engine throttles ware adjusted to set up the i_ _
desired sink rate for the vertical touchdown. _j
J Ig71025733-026
- 20 -
Handling Qualities
The NASA handling qualities tests concentrated on selected stability
and control characteristics of the aircraft and stabilization system that
would be of general interest for future commercial V/STOL transports.
Figure 15 gives a detailed range and time history of one of the simulated
IFR approaches where the lift engines are started in level flight, a 7°
glide s!o_e was tracked to 75 meters of altitude, aftez which the aircraft
was flared to commence the vertical landing under visual conditions. These
data are presented as a basis for the following _Ls_us_on'..... of handling.
A major element in the success of the DO 31 to perform a precision
approach and to make a safe vertical landing is the pitch and roll attitude
stabilization system. This system has I00 percent authority and dominates
the basic aerodynamic stability and control characteristics. This system .-
is designed so that the pilot can command pitch and roll attitude and yaw
rate in proportion to control deflection over the major range of pilot
inputs (see figure 8) and over the speed range for the VTOL configuration
(from about 160 knots to hover). In addition, an automatic trim feature
is provided that permits the pilot to pre-select the desired pitch attitude
which is then commanded with a button on the stick. The system has been
optimized for the hover task, and it minimizes aircraft disturbance by
atmospheric condition, configuration change, or asymmetry such as produced
by engine failure. _:_
Because of the limited test time, extensive documentation was not per- i_ ' formed and the iuformati_n should be considered as an overview rather than a
detailed analysis, It should also be noted that prevaillng atmospheric
i'
]
i 1971025733-027
- 21-
conditions of winds and gusts were accepted and they may have affected some
of the initial and transient conditions.
Conversion: The preconversion configuration is established price
to acquiring the localizer. It was noted that little time was needed to
engage and verify that that stabilization system was in an operative mode.
When the lift engines are started in level flight using the procedures de-
scribed earlier, the conversion can be performed with little altitude change
and only a small attitude change (figure 15(b)). In several cases the
lift engines were started after the glide slope was acquired rather than
in level flight. An example where thls was performed on a 12° hooded
approach is shown in figure 16. The conversion procedures were similar
to those _ere the lift engines were started in level flight, and no major
piloting problems occurred provided that the intercept altitude was raised
to allow sufficient time for tracking.
No significant handl_ng qualities pEoblems existed in maneuvering the
_! aircraft to intercept th_ locallzer in the conventional flight regime.
_ The aircraft handles as a large docile fighter with light control forces.
*?_i The conventional surface deflection per unit control deflection is _educed with a gear changer as airspeed is increased to give good response, force
and force per unit acceleration characteristics at higher speeds.
_ ILS acquisition and tracking, lon_itudtna! flight path and control:
_ At the higher speeds, say above 100 knots, changes in angle of attack pro- .
_' dueed by pitch attitude change were very effective in (!) maintaining flight
path while the lift engines were advanced to the approach setting and (2) _
,: in changing the flight lath to acquire the ILS. For exak_ple, an acceleraticr _ 1971025733-028
- 22 -
of O.ig normal to the flight path was obtained in response to a stick step
that changed angle of attack only I or 2 degrees. When the pre-select
feature of the stabilization system was used to change flight path angle,
moderate pitch rates were produced (3 degrees per second). The longitudinal
control (stick) was also very effective in producing a higher frequency
input when the pilot desired to make additional corrections or to compen-
sate for gusts, etc. When large attitude changes were used, such as to
acquire the glide slope, a horizontal acceleration was produced at a time
when the pilot desired either a constant airspeed or a reducing airspeed.
The main engine nozzles were very effective and easy to use in controlling
the airspeed at these times (see figures 15 (a), (b), and (c)).
As the airspeed was decreased below I00 knots the longitudinal control _
became less effective in producing flight path changes. For a decelerating
approach not only does the angular response to control input change, but
_i also the flight path response to angle of attack changes; consequently, ....._ the pilot must continually re-adjust his gains as the airspeed decreases.
At the lower speeds large angle of attack changes must be made to develop
the desired normal acceleration and these introduce undesirable airspeed
to the flight path. Therefore, other methods of control were evaluated
I! changes because of the rotation of the engine thrust vector with respect , and documented, b comparison was made in the 60 to 90 knot speed range of l using 1) lift engine thrust, 2) main engine thrust, 3) main engine nozzle
_ deflection, and 4) pitch attitude to control flight path and airspeed while "_, J -_ tracking the iLS and correcting for wind and shear conditions. At speeds _
'i", ? ]97]025733--029
- 23 -
below 90 knots the pilots preferred to modulate lift engines for tracking,
but noted that there was insufficient control for large upward corrections.
The peak measured incremental normal and longitudinal accelerations pro-
duced by these controls are compared in fugure 17 with values calculated
from the thrust components. Figure 18 presents time histories of the air-
craft response to these controls; the incremental changes in veloeity_
altitude, and flight path were calculated from the accelerometer readings
and represent the change due only to the control (see Appendix A).
Modulation of the lift engine produced a maximum of _ O. ig normal
acceleration (-figure 17(a)). Only small acceleration parallel to the flight
path was obtained because the llft engine axis is inclined 15° from the
fuselage reference llne (figure 2), and the aircraft was flown near zero
degrees angle of attack. The peak acceleration was rapidly achieved he-
cause of the short engine time constant (about 1/4 second), and the measured
-_ accelerations agreed well with the values computed from the thrust components.
It can be observed in figure I8(a) that the normal acceleration decreases
rapidly after the throttle inp,_t. This decrease is caused by the damping
_ in heave (change in lift with angle of attack) at constant pitch attitude.
4 •" The result is a fairly constant increment in vertical velocity 3 seconds
_ after the input to provide a flight path change_ Referring to figure 18(a),
_,' ._, it can be seen that the altitude change produced by a control input of
about 60 percent of the maximum is mall; after _0 seconds the altitude
increased only 6 meters which is equivalent to only 1/10 of the slide-
__ P slope beam width at an altitude of 200 meters (ass'Jmins_ =-7 ° + 1° :_' figure 9). 1971025733-030
- 24-
For a main engine throttle step (figure 17(b)) the magnitude of normal
acceleration was similar to that produced by the lift engine. However, a
large longitudinal deceleration accompanied an increase in propulsive force
because the nozzles were deflected 120°. In contrast to the reduction in
the normal acceleration by heaving there is little reduction in the longi-
tudinal acceleration; therefore, a large unwanted decrease in airspeed
occurs. The magnitude of the speed change was sufficient that the pilot
had to compensate with a change in nozzle deflection; therefore his work-
load increased. He considered the use of main engine thrust modulation
with the nozzles at 120° to be an unsatisfactory flight path control. He
felt that the time constant of either the main or llft euglnes (about 1/2
second and 1/4 second, respectively) did not detract from the tracking task
during this portion of the approach.
Modulating only the nozzles of the main engines (figure 17(c) and
_' 18(c)) was unsatisfactory for high frequency control of the flight path
_' because little normal acceleration was developed compared to the longitudinal
-i acceleration. Since a large longitudlnal acceleration was rapidly produced, r'_ this nozzle control was useful in making airspeed corrections and in making
_; long period adjustments for large flight path changes (such as required
_ for intercepting the glide slope and adjusting for headwinds). The nozzle
_ deflection could be rapidly changed, and O.lg longitudinal acceleration
was obtained with less than 15 ° nozzle movement.
Figure 18(d) shows the response of the aircraft to an attitude change.
A normal acceleration of 0.158 was produced by pitching the aircraft 6 °. 1971025733-031
- 25 -
_is change in attitude caused a deceleration of 0.1g along the flight path
which resulted in an unwanted airspeed error of 8 knots after 6 seconds.
The combination of normal acceleration and longitudinal deceleration was simi-
lar to that for a main-engine step, and when combined with the large attitude
incremen_ required to develop the desired normal acceleration produced an
unsatisfactory flight p_th conrtol. Figure 17(d) she ss that the normal
acceleration and longitudinal deceleration changes can be approximated by
the lift change with angle of attack and the rotation of the resultant force
(A8/57._, respectively.
As illustrated in figures !5 and 16 good tracking of the glide slope
could be achieved by using the lift engines provided the pilot was initially
on the glide slope. In the process of evaluating flight path control, glide
slope offsets were purposely introduced to simulate situations that might
occur in normal operations such as caused by wind shears, turbulence, etc.
With the glide slope set at -7 ° _ 2°, offsets below the glide slope of 1/2
dot (1/3°) or less and any offsets above the glide slope posed no major
l problems. Offsets below the glide slope of I dot (2/3°) or more brought
i about expected power management problems because of insufficient normal
acceleration provided by the llft engines. This problem is illustrated
in figure 19. At t = 14 seconds the pilot advanced lift engines to the
maximum normal thrust level, but there was little change in glide slope l
_ error. At t = 30 seconds the nozzle deflection was reduced, but this also
did not correct the glide slope error because the primary effect of reducing
nozzle deflection was to increase airspeed. Finally at t = 39, the aircraft _ _"
attitude was increased; then the glide slope error decreased, and the air- T - speed also decreased. If the aircraft had first been pitched, the glide 1971025733-032
- 26 - slope error could have been corrected earlier, but the airspeed would have dropped to a much lower value than desired. Figure 20 shows a tracking run done where main engine throttles were modulated rather than lift en- gines. Comparing the glide slope error with that of figure 15, it is seen that eq_,ally good tracking was obtained; however, the pilot workload was greater when tracking with the main engines because of the undesirable airspeed perturbations with the nozzles at 120° (120° used for the desired flight path and deceleration). This effect on the open loop deceleration is evident in figure 20 at t = 39 seconds where the main engine thrust was increased to avoid going lower on the flight path; shortly thereafter the airplane decelerated to below 60 knots which was below the desired speed scheduled. Then the pilot decreased the nozzle deflection to increase airspeed, but the increase to 80 knots was too large, and the nozzles were rotated back to the full braking position to arrest this overspeed. Thus, one large tracking error can possibly force the pilot to modulate two, three or four control levers at a time when he prefers to keep constant as many parameters as possible.
Reference 7 presented a criteria for satisfactory STOL flight path control during ILS tracking as _ O.lg normal acceleration to be achieved in less than 1.5 seconds. This criteria was satisfied by the DO 31 oper- ating in the 60-90 knot range with lift or main engine thrust modulation, but the pilots considered either control unsatisfactory for tracking an
ILS. It is concluded that the criteria was inadequate because it only specified a maximum acceleration normal to the flight path to be achieved within a given time. It appears that the flight path control criteria
.L 1971025733-033
- 27 -
should include changes in alt%tude and/or flight path after several seconds, and should also limit airspeed and attitude changes. Such a criteria would bu analogous to lateral control criteria where time to bank 30° is specified with the maximum permitted cross coupling. There are insufficient data at present to revise the criteria for tracking and ILS with V/STOL aircraft; however, the following recommendations are made:
(i) The control of acceleratlon normal to the flight path should be
achieved with little acceleration along the flight path (i.e.
"direct lift control" is desired). When normal acceleration is
increased (upwards), an acceleration along the flight path is pre-
ferred over a deceleration; a deceleration along the flight path
greater than 50 percent of the normal acceleration is unsatisfac-
tory. The flight path should be changed at least 2 degrees witbin
2 seconds after the control and thereafter the flight path should
not return towards the initial conditions,
(2) Independent control of the acceleration parallel to the flight
path should have no appreciable do_ward acceleratiou_ and a small
upward acceleration is desired.
Additional simulation and flight tests are required to define the criterl_ and provide limits to cross coupling {such as unwanted airspeed changes),
Final transition and vertical landinKI lonKitudinal and_eight control:
The final transition to a vertical landing was shown in time history form in figure 15(d). To reduce the pilot workload the normal precision approach P • procedure was to fly the aircraft near zero lift so that the main and lift t 1971025733-034
- 28 -
engines settings were near hover values and to use a main nozzle deflection
o_ about 120° to decelerate the aircraft to an airspeed of 50 to 6G knots
by the time the altitude was down to 50 to 70 meters. At this point (t =
ii0 seconds in figure 15(d)) the aircraft is flared by pitching to +5 °
attitude with the pre-select trim system; then the lift engine thrust,
main engine nozzle deflection, and aircraft heading is adjusted to maintain
the aircraft over the touchdown area. The llft engine thrust is readjusted
to produce a small sink rate (less than 2 meters per second). As riteair-
craft descends below 15 meters of altitude, the sink rate increases because
of recirculation and relngestlon. In the descent between 60 meters and
15 meters, the pilot can increase the lift engine thrust to reduce the x
sink rate; however, there was concern that the resulting increase in gas
cloud could increase reingestion into the main engine and increase rather k than decrease sink rate. Figure 21 illustrates the suckdowm magnitude when
_ making a vertical landing in the DO 31. This time history of altitude,
_- sink rate, and vertical acceleration is typical for a low sink rate descent
_ when there is no increase in lift engine thrust Just prior to landing to compensate for (1) suckdown forces on the under surfaces of the airplane,
and (2) main engine thrust loss due to exhaust gas reingestton. The result
is a downward acceleration of approximately O.lOg one second before landing
with a touchdown impact of about 2.5 meters per second induced by the com-
bined suckdown and reingestion factors, For4example siren in £igure 15(e),
the lift engines were increased to the maximum normal thrust setting Just
before touchdownj and yet the touchdown descent rate was more than 1 meter
per second, Thus it can be seen that below 10 meters o£ altitude the 1971025733-035
- 29 -
landing commitment is definite. Even though the main engine thrust could
have been increased, the effect on the relngestion was of concern, and thu_
there was no "go around" capability in these tests. At any point down to
30 meters, waveoffs were easily accomplished by pitching the aircraft ± 5°
and reposltioning the nozzles to 65°.
The NASA pilots considered the available control from lift engine
modulation insuf£icient for descent control in hover. Sinc_ the normal
acceleration was less than 0.1g_ the vertical velocity damping near zero,
and the lift engine time constants small, these results are in agreement
with those of references 12 and 14. WLen this limited control was coupled
with the recirculation effects, the NASA pilots rated the vertical descent
and landing unacceptable for a commercial VTOL transport.
ILS tracking, lateral-dlrectlonal fliRht path control: At speeds above
50 knots large lateral corrections were difficult to make. The aircraft
# response to a lateral step with the stabilization engaged on all axes and
.'4. _ with rudder pedals fixed is given in figure 22(a). For this test the
and the sideslip is related to the bank angle. The pilots noted that large _ stabilization system maintains zero yaw rate; there is no change in heading, _V bank angles were needed to develop the desired lateral velocity, high side-
sllp angles developed, and it took longer to make the correction th_n was
desired. It was concluded that large lateral corrections co._d no_ be made
satisfactorily by only tranalating the aircraft. When the pilots used the
_" directional control (which commanded yaw rate) to coordinate the maneuver, i' it was impossible to find the correct input to maintain small sideslip angles; the aircraft responded as if it had no directional stability. 1971025733-036
- 30 -
Tests were also made with the yaw rate stabilization off. For this condi-
tion and no pedal input (figure 22(b)), the sideslip excursion was propr-
tional to the bgnk angle (AB/A_ = I), and the aircraft had a fairly long
directional period (7-8 seconds) with low directional damping. It was
concluded that additional augmentation for turn coordination was needed.
Reference 7 pointed out that it was necessary that A_/A_ be less than 0.3
for satisfactory handling, and methods to achieve theee levels were also
discussed therein.
The requirement for large bank angles to develop suitable lateral
velocities was not expected based on small-scale tests of the DO 31 (ref-
erence 8) nor on predictions made by Dornier personnel. The static lateral-
directional characteristics measured in flight are given in figure 23.
These data show that I0 degrees of bank angle are needed to achieve a lat-
eral velocity of I0 knots at 60 knots forward speed. The bank angle per / :' unit sideslip was 2 to 3 times that calculated from the tests of reference i_i 8 where the llft engine flow was simulated but the main engine flow exhausted
at _ero degrees raLher than 120 degees used in the flight tests.
Fqr small corrections to be made while tracking the iocaiizer beam,
_, the easiest procedure was to use the directional control (with yaw stabili-
_ zation engaged) and iec the aircra£L L_auslaLe ...... "_....._ t_o
held level by the attitude stabilization system. In this case the desired
_ heading was maintained by the stabilization system and this avoided the _- wandering exhibited by Qther V/S_nDL aircraft at comparable approach speeds
_' (reference 7). ----
r " i
_ 1971025733037
- 31 -
The lateral control, sensitivity, and response were satisfactory at
the transition speeds; however, the attitude command feature required a
lateral force be maintained by the pilot in turning flight, and even though
the force was low it was unnatural and uncomfortable to the pih>to
Final transition and vertical landinK, lateral control: Considerable
research has been performed by NASA and others on the maginitude of lateral
control power needed for satisfactory performance of the hover task (ref-
erences 9 and I0); however, there remains considerable postulation on the
effects of aircraft size and degree of stabilization, (references II and
12). Since the DO 31 is the largest VTOL tested by _.SA (45,000 Ibs.)
has an attitude command stabilization systemp and has a relatively low
lateral control power installed (0.8 rad/sec2), it offered a unique oppor-
tunity to examine the lateral controllability in f_ight in a realistic
environment and to compare the results with simulator prognostications.
The following evaluation and discussion of the lateral control power char-
; acteristics of the DO 31 during very low speed flight (at or near hover)
is in the form of reference 12 where the maneuver, trim (balance) and upset
requirements were discussed. It is assumed that the aircraft is being
operated as a commercial V/STOL transport; i.e., only modest VTOL maneuvering
_.tl ..... L- is required, an engine •az_u_u m_t u= co_t_ol!ed, and _,:=- =-__=ft...... lu_ o
be operated in adverse weather conditions.
First, the maneuvering charscterlstlcs will be examined. The effect
of stick deflection on the static rolling moment available and on the an- _!
gular acceleration is shown in figure 24. It is seen that the maximum \
=- T
= 1971025733-038
- 32 -
rolling moment is essent_ally independent of the lift engine throttle set-
ting (_FC_" Figure 24(b) contains several measured peak roll acceleratio_
values. These values are larger than the calculated values of_),ecause
the rolling moment is based on a static vslue and the stabilization system
initially con_ands higher acceleration to give more rapid angular response
to a normal pilot input. The more rapid response is evident i_ the time
histories of a stick step, figure 25, With stabilization (left hand figure)
the lift engine fuel control units (FGU) are commanded to very large deflec-
tions shortly after a pilot's input of 1/4 stick deflection_ and the initial
response (t < 2 seconds) is greater than without stabilization (right hand
figure). With stabilizationja 5° bank angle was attained iu 1 second for
this input, and the bank angle reached 90 percent of the steady state com-
_ manded value, i0°, in about 2-1/2 seconds. As the pilot input is increased
_ in magnitude, the differences in aircraft response with and without sta_ _i bilization will become less because there is less exces_ moment available
_.i to increase the acceleration. The pilot input for the bank step with sta-
__ bi!ization is a stick step and is easy to perform; the pilots co-sidered
_ the response, damping, and sensitivity to be satisfactory. The lag of m_
i peak RPM behind the stick input _eflects the lift engine time constant of
-i _im.2 e andhistory0.3 sebaseconds.d on aThenaturalaircraft frequencyresponse of closely2._ radiamatnschedper a second,calculated a m
ji damping ratio of 1.1, and an initial lag of 0.2 seconds; these charac- study i teristics were within the optimum areas defined by the simalator
Ii reported in reference IS. Since the pilots rated the lateral control S
2i.i .... - 1971025733-039
- 33 -
sensitivity and maximum control power (a value of 0.8 radians per _econd
squared) as satisfactory, good agreement was obtained with the study of
reference I0. The control inputs of figure 25 were taken to document the
aircraft response, and do not provide a measure of the lateral control
needed for maneuvering_ To make this evaluation the pilot was given two
tasks: on, ,,_ _o perform what he believed to be the most extensive lateral
maneuvering around the hover area, and the other was to determine the maxl-
mum lateral velocities that he _uld expect to normally use with a commercial
VTOL operation. Figure 26 illustrates the pilot input and control needed
to extensively maneuver the aircraft near hover. Figure 27 shows th_ time
history where the pilot slowly increased the bank angle in order to establish
and measure lateral velocities. From these tests it was found that the
maximum control power needed for control and stabilization during lateral
_, maneuvering was +0.4 radians per second squared, and the maximum lateral
i velocity over the ground that w_uld normally be expected in maneuvering ...... _' "_ V/STOL transports was I0 meters per second (20 kts). Higher lateral air
_-_ _peeds may be encountered when it is necessary to precisely position the
-_'-_ aircraft in crosswlnds.
Next, the trim, or balance aspects are examined. Some VTOL aircraft
_ have required large amounts of lateral control to trim lateral moments
ential RPMas lateral velocity increased, figure 27, it is inferred that iI developed In sideward flight. Referring to the lack of change in differ- _ti-" little or no control moment was required in sideward flisht for the DO 31 _" configuration, at least for velccities of 10 meters per second. Another
control requirement is to balance an engine failure. Figure 28 shows that ,:_ 1971025733-040
- 34 -
the static moment resulting from a llft engine failure can be easily balanced°
Sufficient moment is also available to balance a main engine failure; how-
ever, little moment is available if the remaining main engine is advanced
to a high setting to compensate for the llft loss. The dynamic response
to a lift engine failure is shown in a time history of the shutdown of the
#I llft engine performed with the hover rig on the pedestal, figure 29.
It is seen that with a fixed stick position the aircraft rolls only 2 de-
grees and within one second of engine shutdown itltlatlon, the bank angle
starts to return to the wings level position. During this compensation
the remaining lift engines are initially commanded to a near emergency
FCU level in the left pod and a near idle in the right pod to limit the
rolling; this represents about 80 percent of the available control moment.
+ Shortly thereafter the difference in FCU levels is reduced to maintain
static symmetry. A larger bank angle was produced by a main engine failure
the response to an engine failure was satisfactory; however, since the and it took longer to return to wings level. From a piloting viewpoint ._.
_ asymmetric moment was automatically trimmed out without changt,ng stick
position or force_ the pilot had no direct way o£ knowing that he was near a control limit except by reference to actuator positon gages on the in-
_ strument panel. From these data it can be ascertained that the greatest
lateral trim requirement for the DO 31 is produced by an engine failure.
No flight tests were performed with an engine failed, but since 3/4 of
the lateral control was needed to statically balance the remaining main
engine, it would be expected that _arginal control remained for a vertical
landing. For the case with the lift engine failed, less control moment
! 1971025733-041
~ 35 -
was, needed; and the remaining lateral control power should be satisfactory
for some maneuvering during the landing.
Finally, a few comments are given on the upset aspect. The control
power needed to compensate for £'zsts could not be determined; however, it
can be stated that the stabilization system was very effective in control-
ling upsets due to atmospheric conditions that were encountered. In fact,
the pilots remarked that the aircraft was very stable in a large variety
of conditions such as headwind, crosswind, and gusty air. This is further
verified by the ability of the stabilization system to compensate for an
engine failure (figu, _ 29) which would be comparable to a gust producing
an angular acceleration of about 0.4 radians per second squared.
It should also be noted that care was taken to keep the friction and
force gradient of the control system low (figure 4(b)). The values corres-
pond to those recommended in reference 12.
/ In conclusion, the lateral control power of the DO 31 (0.8 radlans
per secona squared) was sufficient to provide a satisfactory hover control il provided that attitude stabilization was utilized. It should be noted that ......
the aircraft could be flown by a research pilot with the stabilization
I, ' esystngineem ofailuff, resthe _workloadere not adreqequuatireedlytoehovalveruateand d inlandthesesuchtests,a craftAltihoughn com-
mercial operation were considered unacceptable even for an emergency oper- J ation. Based on these tests significant reduction in control power cannot
be recommended for this class and configuration of aircraft, 1971025733-042
- 36 -
Trim considerations in transition and hover: Throughout the flight
regime with lift engines operating, the attitude command stabilization
system very effectively performed the trimming function so that the pilot
was generally not aware of out-of-trim moments. Although this greatly
simplified the pilot's task, some warning must be given to the pilot if
control limits are approached. For example, the previous sectiol, pointed
out that a large amount of control was needed to compensate for an engine
failure, and yet the pilot is not aware of the remaining control because
the stick remained centered. This sltuation also occurred during a ver-
tical takeoff and transition, figure 13. Ref_z;ing to the pitch nozzle
positioo it is seen that at 40 knots 80 percent of the longitudinal control
is required to compensate for the nose-up pitching moment due to the aero-
dynamics and propulsion system, and yet the stick is centered. i Another aspect that must receive additional attention when stabili-
zation systems are incorporated is the complaint by the pilot that a force
must be maintained in turning flight at transition speeds. In this respect
_ a rate command with attitude hold may be preferable.
The ability to pre-select the desired pitch attitude and to actuate
this with the button on the control stick, when desired, was a very desirable
feature of the pitch attitude command system, because it reduced pilot
workload during the approach when discrete pitch changes were required. 1971025733-043
- 37 -
Terminal Area Operation
It was noted earlier that a wide range of ILS approaches could
be made with this aircraft because of its large operational envelope
and good control and stabilization system. This section will review
the approach in terms of constraints that may be imposed on a commercial
V/STOL transport operating in the terminal area.
Cruise letdown to pre-approach configuration: Figure 30 presents
a time history of a letdown from cruise altitude where the deflection
of the nozzles of the main engines is used for controlling descent
rate. The descent started at an altitude of about 2500 meters and an
airspeed of 260 knots and ended at an altitude of 450 meters and 140
knots with localizer capture; the engines were set at a moderate
thrust level, NF = 72%. The maximum nozzle deflection permitted from
structural considerations was 90 ° between 250 - 200 knots, and 120 °
below 200 knots. A heavy buffet accompanied the 120 ° setting during
the descent and would be unacceptable from an operational standpoint ...... , t' _
It was determined that an 85 o nozzle setting was about maximum to avoid
buffet, and this setting resulted in descent rates in excess of 20 meters
per second, a sufficiently high descent rate for rapid letdowns. The
use of the nozzles is considered an excellent method of establishing
a varying rate of descent during the letdown.
Conversion: Before converting to the VTOL configuration the
pilot maneuvers to intercept the localizer_ he then needed about 30
I seconds of tracking time while the aircraft stabilizes from the pre- _:_ conversion changes of gear and flap deflection. Then he initiates the
/,
°'! __ 1971025733-044
- 38 -
start of lift engines. By properly combining pitch attitude_ nozzle
deflection_ and main engine thrust, there _s little altitude or air-
speed change during this operation (figure 15(b)) even though the lift
engine idle thrust-weight ratio was about 0.35. The time to attain a
stable idle was about 20 seconds and the pilots considered that this
was too long because it distracted them from other flying tasks. This
distraction was minimized by assigning this monitoring task to the co-
pilot. Since there was no ground based guidance information_ such as
distance measuring equipment or beacons, level flight conversions
were difficult to initiate at the proper location except when landmarks
were visually used for position.
The lift engines were also started on the glide slope (figure 16i3
and in this case a better reference for starting the lift engines was
\ provided by reference to the altitude. Due to the length of time
t required to start the lift engines, the intercept altitude had to be
_ raised when starting the lift engines on the glide slope.
ILS acquisition: Referring to figure 153 before the glide slope
is intercepted, the desired pitch attitude for the approach is pre-
selected (-lO ° for-7 ° glideslope). When the aircraft nears the glide-
slope centerline, the acquisition is initiated by raleasing the pre-
[: selected trim, by advancing the lift engine throttle to a hover setting,
the 65 ° 120 ° j by deflecting nozzles from to (braking), and by increasing
in less than 4 seconds. It is seen that by following th_se procedures
the flight path is changed with little overshoot, and about 15 seconds _i the main engine thrust. All of these changes are performed by the pilot later the pilot is confident that the glide slope has been acquired.
Ii ) i ]97]025733-045
- 39 -
At this point (t=70 seconds) the aircraft has decelerated to about
100 knots and is at an altitude of 250 meters_ the pilot can then
proceed to track the ILS. By properly combining main and lift engine
throttles: main engine nozzle deflection and longitudinal pitch control:
it was found possible to intercept 7 and 12 degree glide slopes and to
track these slopes with acceptable accuracy while decelerating from
140 to 50 knots and descending to a breakout altitude of 70 meters,
However: the NASA pilots considered the workload imposed by the
numerous discrete control steps to be unacceptable for commercial VTOL
transport operation.
In order to examine the feasibility of reducing the time in the
V/STOL configuration glide slope intercept altitudes of 300: 450: and
600 meters were tested for the 7° glide slope approaches. The 450
meter intercept altitude was preferred when the lift engines were
started in level flight since this permitted a reasonable glide slope
acquisition and tracking time (approximately one minute). The 300
) meter intercept did not allow enough tracking time with the given
deceleration schedule. The 600 meter altitude was used when more time
_i was needed; e.g. when the lift engines were started on the 7° glide
slope. Because of the higher descent rates that occur during 12 °
approaches_ the intercept _s raised to 600 meters when the lift
engines were started in level flight, To obtain adequate tracking
time when the llft engines were started on the 12 ° glide slope instead
_i_ _ of in level flight_ the intercept _ititude had to be r_ised to 900 '
'_ meters.
j, 1971025733-046
- 40 -
Although these intercept altitudes may be peculi&r to the DO 31 configuratlon an_ will vary for other concepts_ this study and other similar studies (reference 15) show that with only simple_ situation
Information displays the pilot needs 20 to 30 seconds for acquisition
(time from intercepting the glideslope to confidently acquiring it) and 20 to 30 seconds for tracking to assess the approach so that he can confidently proceed to a landing.
ILS tracking: Once the glide slope was acquired, the ILS beam was tracked by modulating lift engine thrust. If the glide slope has been accurately sqquired and no large er_ __'_ ave been introduced, the tracking performance is good (see figures 15 and 16), and the pilot noted that workload is relatively low. The simulated IFR portion of the approach is ended at a breakout altitude of 70 meters at which point the airspeed has stabilized at 50 - 60 knots. The use of attitude stabilization contributed significantly to making these approaches successful. The pilots commented on the usefulness of the main engine nozzles to match the approach schedule with the desired ground speed. By giving the copilot tbc task of controlling airspeed with the nozzle, the pilot workload was significantly reduced during the simulated IFR ,_pproaches. For some tests the glideslope beam width for the 7 ° approach was decreased from + 2 ° to + 1 ° with no apparent increase in pilot workload or degradation in tracking performance.
The reason that llft engines were used for glide slope tracking and their limitations were presented in the section "Handling Qualities." i. Even though there was insufficient flight path control available from any individual control during the approm_h _nd landing s it is believed %
I! 1971025733-047
41 -
that the aircraft had sufficient normal ac,_.elerat_oncapabilit_ by
combining various controls; however_ f_rther resealch is required to
properly integrate these controls in a more manageable form. It
should be possible to integrate throttles and nozzle deflection, or
it might be desired to add a servo-system to provide a speed or de-
celeration command system. Consideration should,also be given to
displays that give the pilot a better visualization of the thrust
vector.
The majority of these ILS approaches were made at low lift to
reduce the pilot workload. These approaches were made at an angle of
attack of about -3°; the relationship,_ _._ dictates a pitch
attitude of -lO° for a -7° approach and -15° for a -]2° approach,
respectively. Since these approaches were _th a large Dose dow_
attitude and the aircraft decelarating, the resulting force on the
pilot (and potential passengers) is impractical for coH_nercialoperation.
Therefore, several approaches were attempted _th the fuselage more
"_ nearly level. These were notably unsuccessful. Insufficient ti_e was
available to explore the problem_ however_ it can be partially
._-, att_-ibutedto the decelerating approach where the lift at a positive
_ angle of attack significantly is reduced as the approach progresses.
. _ From static conslderations__ llft d_._, at constant _'_LL_U_I wOUld
_ be about 0.15 times the aircraft weight when decelerating from i00
knots to 60 knots_ nearly the total range of lift engine thrust modu- lation, in addition_to properly increase the thrust to compensate for
This problem requires further examination since fut,,___eV/STOL concepts
!_i: this lift deflcit_ the pilot must also modulate thn_st to track the ILS. _ _ ;_" _J
% 1971025733-048
- 42 .- envision the use of wing lift to reduce both nose-do_napproach attitudes for passenger acceptance and power .¢eq_irements for noise acceptance.
Final transition and vertical landing: At an altitude of 70 meters and an airspeed of 50 kn_ots, the !FR portion of the approach is terminated and a flare is initiated. The steps and procedures required to accomplish a safe vertical landing were discussed in the "Handling
Qualities" section. It can be noted in figure 15(a) that shortly after breakout the pilot went below the flight path_ this was done to assure himself that he would not overshoot the _houehdown area.
During the vertical descent it was difficult to see the touchdo_._ area, and reliance had to be placed on the vertical descent rate obtained from the radar altimeter. Figure 21 showed the magnitude of the suckdown which precluded low altitude hovering. This condition is
1_acceptable for a commercial VTOL transport. For complete IFR hover and isnding operation_ displays must be developed that provide additional s_tuation information.
Lo_w spe_d translation: Forward translation in hover can be accomplished by either modulating ma_n engines nozzle deflection
(fig'o_e31(a)) or by changing pitch attitude (figure 31(b)). Modulating nozzle deflection was attractive, because little lift engine thrust change was necessary to maintain altitude Stopping on a desired spot was difficult, however, because it was hard to predict the decel ration fro_ a given nozzle setting. Accurately and quickly selecting a nozzle _ position during the demandin_ t_sk of maintaining altitude while _,' maneuvering was also difficult due to the small size of the nozzle _i 1971025733-049
- 43 - position instrument. When pitch attitude _s used for trans!ating_ large attitude changes were required when stopping and visibility was impaired. EVen though the lift engine thrust had to be coordinated with attitude change to maintain altitude 3 the use of pitch attitude change as the primary means of translating for short distances was preferred over the nozzle modulation_ since it did not require looking in the cockpit to monitor nozzle deflection. The pilots felt that if
(i) the nozzle position were more clearly displayed and (2) the stabilization system had a height control feature_ nozzle deflection modulation may be the preferred control for translating 3 particularly for longer distances.
Lateral maneuvering in hover is accomplished by changing roll attitude and tilting the thrust vector. Figure 27 presented a time history of a lateral tral.slation. The maximum speed obtained during this test was i0 meters per second_ which the NASA pilot considered to be the maximum lateral velocity nonzally required for a commercial
VTOL transport of this size. The stabilization system reduces aircraft disturbances from the unusual wind conditions near the grou_ud and
allows the translation to be executed _th low pilot workload.
Environmental effects at transition speeds: In the course of
the test program a large range of e_ironmental conditions was
encountered. _e wind speed ranged from 12 meters per second (24 knots) to cs]m_ and the direction ranged from head wind to crosswlnd to tail- _ wind. Light to moderate turbulence was encountered on several flights. _i_ In the preconversion mode (airspeeds greater _han 140 knots) the air- _
cr_ft was quite disturbed in the lateral-directional mode_ however_ _
,' c' 1971025733-050
- 44 -
J when the stabilization system was engaged and lil't e_igines started_
the aircraft _as no longer affected by the turbulence and "it felt
steady as a rock." It _s gratifying to note that the aircraft was
relatively unaffected by gusts_ adverse winds and crosswinds. This
result was somewhat surprising in view of the large sideforce due to
sideslip (figure 23). Tests could not be conducted to isolate the
factor that produced the favorable response in gusty air; it can only
be surmised that the ability to maintain constant attitude contributed
to the favorable ride characteristic. This is borne out by recent
tests with an attitude command stabilization system in a light plane 3
reference 16.
In the approach it was found that crosswinds could not comfortably
be compensated by a sideslipping approach oecause large bank angles -
were required. At 60 knots a lO knot crosswind necessitated a lO °
i bank_ich not only was uncomfortable in terms of a high sid_force 3
but also because a lateral force had to be held by the pilot; con-
_ sequently_ the preferred method of compensating for crosswinds was to
•,_ crab the aircraft. In decelerating approaches the heading can be
,_ slowly changed to keep sideslip near zero as the aircraft slows to a
hover. With proper displays_ the workload should not be too high.
Thus 3 the crosswind problem does not appear to be as serious as with
some STOL aircraft where it is necessary to abruptly decrab the
aircraft before touchdown (reference 7). ¢
_ Environmental effects at ver_ lowspeeds: Figure 32 _illustrates _ h__ the effect of wind and sink rate on the main engine intake temperature •_
_, rise during vertical landing, The landlngs with increased head wind ,_
l 1971025733-051
- 45 -
and greater si_k rate generally resulted in less hot-gas reingestion.
For comparison purposes_ one short landing data point has been added
to the vertical landing data. Along with sink rate and wind_ it
appears the immediate shut do_¢n of lift engines and reduction of thrust
on main engines after touchdown is very important. Greater peak
temperatures occur when the power plants continue to exhaust their
hot gases after the landing impact is made. This landing procedure
is addea to the pilot workload during the final phase of the landing.
The techniques and procedures minimizing recirculation effects
have been discussed_ and the effect of wind on vertical takeoff
recirculation and hot gas ingestion is presented in figure 33. It can
be seen in figures 33 (a) and (c), that a 60 - 90° crosswind _t take-
off considerably increases the main engine inlet temperature on the
upwind side (20 - 25°C). The crosswind might be pushing the gas cloud
from the lift engines on that side into the inlet of the main engine.
The same effect can be seen with a lesser crosswind in figure 33 (d)_ y,'_, the difference being a somewhat smaller increase An main engine inlet
temperature. The vertical takeoff with a 6 kt. headwind 3 figure 33 (b),
yields the smallest rise in inlet temperature.
Each main engine inlet on the airplane had only four temperature
probes_ they were located, at the 12, 3, 6, and 9 o'clock positions I
and the plots show the average of theae. The limited number of z t temperature probes did not show temperature distortion across the 7
engine inlets that can cause compressor stalls. It should be pointed ,_'
out that no compressor stalls were encountered durlng the NASA evaluation. :_ % i • 1971025733-052
- 46 -
Comparison of approaches: Figure 34 compares two 12 ° approaches
and two 7 ° approaches with different lift engine starting conditions.
This figure illustrates the reduction in time and noise footprint that
can be realized by starting the lift engines on the glide slope. Less
time is spent on the approach since the airspeed is kept at a high
level during the first part of the approach, and the lift engines are
started after glide slope acquisition. This figure also indicates a
reduction of the noise footprint because the lift engines were started
while tracking the glide slope. By starting the lift engines on the
glide slope, the high noise level due to the lift engines is 2000 to
3000 meters closer to the landing site with an additional 150 meters
of altitude during the starting cycle. Reference 17 gives some
measured noise values for the DO 31. The approaches and landings
took 2 - 3 minutes_ these times are shorter than predicted in reference
16_ however, they are longer than theoretically achievable based on the
aircraft's performance.
_ Previous studies (reference 7 and 9) have reported that the descent
rates in an ILS approach should be less than 5 meters per second _L (lO00 feet per minutes) at altitudes below lO0 meters (300 feet).
The descent rates used during these test_ were greater. Fig'are 35 is
a graphical representation of the relationship between glide slope
angle, airspeed, and rate of descent. Considering the DO 31 approach _.
airspeed _ange from ]20 kts at g?.ide slope intercept_ to 50 kts at
_._ flare on a 7° approach, the rate of descent varies from approximately _ /_" 7 meters per second at the start of the approach to S meters per second /
._ at breakout. For the 12 ° approach 3 the rate of descent varies from ,_, 1971025733-053
- 47 -
13 to 5 meters per second. The pilots considered the high rate of descent at the beginning of the approach to be no problem because (i) the rate of descent was decreasing as the aircraft was decelerating on schedule 3
(2) the attitude co_and control system allowed more attention to be devoted to approach performance parameters_ and (3) there was a high confidence level that the aircraft could be flared to arrest the sink rate because s_fficient altitude was cllotted (70 meters).
The NASA pilots felt that approaches steeper than 12 ° were not practical with the DO 31 primarily because of the nose-down attitude.
Referring to figure 15 (a) it can be seen that the time to flare the aircraft and to land took as long as the time to acquire and track the glide slope. It would be expected that attempts to reduce the landing time, e.g. by not flaring to zero vertical velocity at 70 meters 3 would make the rate of descent in the approach more critical and these high values might not be tolerated.
The close-in pattern presented in figure 36 represents the type of approach that may be required in a restricted area. The lift engines are started during the turn made to acquire the localizer.
The altitade during the turn is held at the intercept altitude (450 meters), and the distance of the turn from the landing point was dictated by
length of time required to bracket _he localizer prior to glide slope
intercept. After ths localizer was captured, the rest of the approach
is the same as those approaches where the lif, engines are started in o_
level flight on the locali_er. The pilots considered this pattern to be feasible if the appropriate termlnal-area navigation aids were available. _ .i ]97]025733-054
- 48 -
CONCLUDING REMARKS
A flight investigation was performed with the Dornier DO 31VTOL
transport to evaluate the performance, handling, and operating characteristics
that are considered to be important when operating a commercial VTOL trans-
port in the terminal area. The DO 31, a 20,000 kilogram transport, has
a mixed jet propulsion system; i.e. there are main engines with nozzles
that deflect from a cruise to a hover position, and vertical lift engines
that operate below 170 knots. In this VTOL mode pitch and roll attitude
and yaw rate stabilization are incorporated, and the main and lift engines
are used to augment the forces and moments. The tests concentrated on the
transition, approach and vertical landing.
The flight tests showed that this mixed jet-propulsion system provided
a large useable performance envelope which enabled a broad range of simulated
IFR approaches to be made. Glide slopes of 7° and 12° were intercepted
at 140 knots and tracked while _eceleratlng to 50 knots and a breakout
altitude of 70 meters; the transition to hover and a vertical landing had
2] to be made visually because displays were lacking. The aircraft could be i r! ,i easily converted to the VTOL mode either before or after the glide slope
was intercepted. Once the glide slope was acquired, it was easy to track
because corrections could be made normal and parallel to the flight path
. (via lift engine thrust and main engine nozzle deflection, respectively),
_ during which time the stabilization system maintained attitude. The pilots
_ reported that the normal acceleration available from the lift engines (_+0. lg) _
wl_: insufficient for flight path control. Controlling the flight path by _
II 1971025733-055
- 49 -
pitching the aircraft was unsatisfactory because of the changing control
power and lift in the decelerating approach, and also because of the large
unwanted airspeed changes at the lower airspeeds.
During the transition and approach, the pilot's prime Job was power
management (control of thrust magnitude and direction); there were too many
discrete changes in attitude, lift and main engine throttle, and main engine
nozzle deflections. Most of the approaches were made with the fuselage
attitude nearly parallel to the flight path. This simplified the pilot's
power management problem. When the ILS was tracked with a more level fuse-
lage attitude, the workload increased and the performance deteriorated.
It was found that the current criteria for flight path control are inade-
quate because th_j do not consider the aircraft response after a period
of time, and they do not define limitations to the crosscoupling. Further
research is required to integrate the different longitudinal controls to
simplify power management, and to define appropriate criteria.
_- Several other observations were made pertaining to the transition and
: approach mode. First, when maneuvering laterally ac transition speeds,
_i the roll attitude stabilization (combined wit!; yaw rate command) created
problems, and turn coordination should be provided. When not maneuvering,
.i the heading hold feature of the yaw rate stabilization greatly assisted
_ in maintaining the aircraft track.. Second, with the stabilization _ystem
was relatively unaffected _e_ engaged and lift engines operating the aircraft/¢_ G ff_$s_ _ by turbulence. Finally, :rosswinds in the approach_compensated by the "_
"crab" maneuver; a "sideslipped" approach was unsatisfactory. L
J .
7 1971025733-056
- 50 -
In hover, the lateral control power (0.8 radians per second squared),
sensitivity, attitude stability and damply 3 were satisfactory. The results
are in agreement with previously reported NASA simulation studies. The
pilots felt that attitude stabilization is mandatory for satisfactory
VTOL operation. A sudden failure to an acceleration system in the VTOL
mode would be unacceptable. The vertical landing was unacceptable because
of the recirculation effects below 15 meters and because of insufficient
height control.
When the complete terminal area operation was considered, the time in
the VTOL mode was shorter than observed in other IFR flight studies. An
ILS approach, starting with intercepting the localizer beam at 140 knots
and ending in a vertical landing, could be completed in less than 3 minutes.
The pilot needed 20 to 30 seconds to track the localizer; 60 to 90 seconds
to acquire the glide slope, start the lift engines, and track and decelerate
hover and make the vertical landing. In a vertical takeoff the aircraft I to 50 knots at the breakout altitude (70 meters); and 45 to 60 seconds to _ i; has a high acceleration and a steep climbout. In just over 20 seconds
the aircraft attained 120 knots , a sufficient airspeed to shut off the
lift engines. 1971025733-057
- 51 -
APPENDIX A
COMPUTATION OF DATA PERTAINING TO
FLIGHT-PATH DEVIATIONS
In order to relate the pilot co_nents on flight-path control with
the aircraft characteristics and motion, it is necessary to realize the
significant difference between the change in aircraft flight-path
angle, _, and th_ glideslope error displayed to the pilot, _G' This
is schematically illustrated in figure A1 and additional comments
follow to describe the computations that were made for the figures
illustrating flight-path control in the body of the report. During
ILS tracking, the pilot's reference is the ILS beam which is ground
based, s b at the desired flight path, YG, and his position infor-
mation is in the form of deviation from the centerline of the ILS
glideslope beam, _G, displayed on the ADI. When the pilot makes a
correction with a control input, a normal acceleration is produced i! which integrates into a flight-path change, A_, which is not
directly related to ¢G. The A6 G is a function of the change in
altitude which is an integration of A_(hence a second integration
of normal acceleration). The integration of the longitudinal
acceleration produces a change in velocity that may or may not be
desired by the pilot. In a constant speed approach, these pertur-
bations can be observed by the pilot as a change in rate of climb
and a change in airspeed. For the decelerating approaches of the 1971025733-058
I "_....__ ,,_--_.-_,
400 _ .. /-8 3 dots high _""_'_'_..1, Vt Y!I //-r u ..L_ ,'
200 _ *O°t ;:, m _croft path Ah _ _0
to" __ u ection
,., o i i "_'_'_,.j . A: _,h200 _nslmhlcorS responas,o 3 2 I 0 _. Distance to touchdown,km
_., Figure A-I.- Fligh_ path relations
,i 'X', t.
'' -2 1971025733-059
present tests, the perturbations are obscured in the situation information
(ADI) displayed to the pilot. In addition, these changes are also more
difficult to measure by normal i'light-testt_chniques. Ib was found
that the radar data of the aircraft's position was not accurate enough
to d_termine the.changes in flight pat_.produced by contro.l_r.I_ats,
and it was uecessary to integrate the aircraft accelerometers.
First, an analysis will be made for the simple case where the
aircraft is stabilized in plt._h_the angle of attack is near zero,
the initial flight path is -7°, and the aircraft is decelerating, For
this case cos e = cos y _ l.
The p,_.Tturbationof the aircraft to a control input is obtained by
_ using the observed no_mal and longitudinal acceleremeter readings
.q (Az and Ax) minus the initial reference conditions to obtain the
i 1971025733-061
- 54-
Aaz Aa x incremental values, ---- and -- For these decelerating approaches, g g
it is assumed that the initial accelerations would have existed if no
con_l _,p_"_ "_ had been applied during the 10-second period of the
integration. In figures 17 and 18, the ordinate of normal acceleration
is inverted to have the direction of the curve in the sense of the
aircraft motion. The change in _-ercical velocity is now obtained by i Av,_ Aw = #Aazdt. Then A_ = - .-_-× 5i .3, and since the attitude is fixed,
the _'hange in flight path, AT, is equal to A_. This is also presented
on the figures in comparison to the observed angle-of-attack cha_s;
recognizing the poor quality of angle-of-attack information at the low
speeds th_ _.greement is reasonable. Conversely, it can be seen why it
would be difficult to use the angle-of-attack information to compute
flight-path changes. Next, the change in altitude is obtained by a
second integration of normal acceleration; i.e., Ah = ## -Aazdt2.
For comparison, the change in altitude that corresponds to a 1-dot
glideslope change (EG) at 200 meters altitude is shown for a glideslope
of -7° with a beam width of +l °. The change in velocity, AVx, is
obtained as At'x = fAaxdt; it should be noted that this does not
correspond to the change in velocity that would be observed on the
airspeed meter because the initial condition of the aircraft is
decelerating flight, and Avx represents the perturbation due to u control input only. _
"t
t 1971025733-062
_ - 55 -
For the case where the flight-path change due to pitching the
aircraft is evaluated, the computations are similar; however, the
accelerations are transformed to earth-fixed axes before computations
are performed. Since the cos e = l, change in acceleration is retained
Aaz in the form of _ to be more easily compared with the data at g
constant attitude. The change in longitudinal acceleration (along the
flight path) is given as A dv/dt. Because of the changing attitude g
and angle of attack, Ay is presented in lieu of Av z.
It should be noted that the previous derivations are only
approximate, but are sufficiently acc_u_ate to assess the initial
motions of the aircraft to relate aircraft characteristics and
pilot comments.
_J 1971025733-063
- 56-
APPENDIX B
MISCELLANEOUS ENGINE AND CONTROL RELATIONS
Thrust and fuel flow.- The thrust and fuel flow characteristics
of one main and one lift engine are given in figure 37. The relations
between throttle position and engine speed were presented in figure 5.
Control deflections.- The variation of the VTOL control deflection
and of the conventional surface deflection with the pilot control are
given in figures 38 to 40. Also included are the maximum hover control
power about each axis for a nominal hover configuration. Since the
conventional surfaces are simultaneously deflected with the VTOL
control, the control power increases with forward speed. In the
conventional flight regime above 155 knots, a gear changer reduces
the surface deflection per unit of pilot control to maintain a better
stick force per unit acceleration. The reduction in control surface
deflection is illustrated in part (b) of _hese figures by the
_ reduction in the maximum surface deflectL;_n with increased dynamic
_ pressure. For each axis, the maximum deflection of the stabilization
system actuator is 40 °.
\_' i'd_
"_ 1971025733-064
- 57 -
REFERENCES
i_ _ Fry, Bernard L. ;.and Zabinsky, Joseph M. : Feasibility of V/STOL Concepts for:_hart-haul Transport Aircraft. NASA CR-743, 1967.
2. Marsh, K. R.: Study on the Feasibility of V_STqL Concepts for Short- Haul Transport Aircraft. NASA CR-670, 1967.
3. Anon.: Study on the Feasibility of VJSTOL Concepts for Short-Haul Transport Aircraft. NASA CR-902, 1967.
4. Anon.: STOL-V/STOL City Center Transport Aircraft Study. McDonnell Aircraft Corporation, FAA-ADS-26, Oct. 1964.
5. Waldo, Richard K.; and Filton, Peter D.: An Economic Analysis of Commercial VTOL and STOL Transport Aircraft. Stanford Research Institute, FAA-ADS-25, Feb. 1965.
6. Kelley, Henry L. and Champine, RoSert A.; Fl_ght Investigation of a Tilt-Wing VTOL Aircraft in the Te_-minal _ -" _er Simulated Instru- ment Conditions. AIAAPaper No. 71-7, J_ ' ' /. Innis, Robert C.; Holzhauser, Curt Ao; and Qu±_. " C.: ! Airworthiness Considerations for STOL Aircraft. NASA t_Q4, 1970.
8. Smith, Charles C. Jr.; and Parlett, Lysle P.: Flight Tests of a O.13-Scale Model of a Vectored - Thrust Jet VTOL Transport Airplane. _ NASA TN D-2285, 1964.
Flight Investigation of Hovering and Low-Speed VTOL Control .Requirements. NASA TN D-2788, 1965.
i i0.9. Garren,Greif, RichardJohn F.K.;; Kelly,Fry, EmmettJames R.B.;; andGossett,Reeder,TerrenceJohn P.D.;: AandVisualGerdes,
for VTOL Aircraft° NASA SP-II6, 1966.
_ Ii. Anderson,Ronald M.:Seth SimulatorB.: ConsiderationsInvestigationsfor Revisionof Variousof ControlV/STOL HandlingSystems Qualities Criteria. NASA SP-II6, 1966.
12. Anon.: V/STOL Handling, 1-Crlteria and Discussion. AGARD Rep. 577, 1970.
13. Rolls, L. Stewart; and Gerdes, Ronald M,: Flight Evaluation of Tip Turblne-Drlven Fans for Lateral Cmntrol and a Hovering VTOL Aircraft. NASA TN D-5491, 1969.
i 1971025733-065
- 58-
]4. Kelly, James R.; Garren_ John F.; and Deal, Perry L.: Flight Inves- tigation of V/STOL Height-Control Requirements for Hovering and Low-Speed Flight under Visual Conditions. NASA TN I)-3977, 1967.
15. Reeder, John P_: The Impact of V/STOL Aircraft on Instrument Weather Operations, NASA TN D-2702, 1965.
16. Loscke, Paul C.; Barber, Marvin G.; Jarvis, Calvin R._ and Enevoldsen, Einar E.: Handling Qualities of Lift Aircraft with Advanced Con- trol Systems and Displays. NASA Aircraft Safety and Operating Problems NASA SP 270, May 4-6, 1971.
17. Flemming, M.; and Schotten, R. (Dornier GmbH_: Noise Problems of VTOL with Particular Reference to t Dornler DO-31. Royal Aero- nautical Journal, August 1969.
i 1971025733-066
- 59 -
TABLE I
AIRCRAFT DIMENSION ANDDESIGN DATA
General:
Length, m, (ft.) ...... 20.60 (67.6)
Height to top of vertical fin, m, (ft.) ...... 8.53 (28.0) _win :
Area, m 2, (ft.2) ...... _ ...... 57.0 (613)
Span, m, (ft.) ...... 17.0 (55.8)
Mean aerodynamic chord, m, (ft.) ...... 3.415 (ii°2)
Aspect Ratio ...... 5.05
Sweep, deg ...... 8.5
Airfoil section, root ...... NACA 64(A412)-412.5
Airfoil section, tip ...... NACA 64(A412)-410
_ Incidence angle, deg ...... 2
\! Dihedral angle , deg ..moooeo...oeoo.o.$_._.. i . 5 Taper ratio, deg ...... 0.615
Flap deflection (max), deg ...... 45
Flap area, m 2, (ft.2) ...... 6.64 (71.4)
Flap chord, m, (ft.) ...... 0.85 (2.8)
Aileron deflection, deg ...... +25
Horizontal Tail:
Area, m2, (ft. 2) ...... 16.4 (176)
Span, m, (ft.) ...... 8.0 (26.2)
Mean aerodynamic chord, ms (ft.) ...... 2.13 (7.0)
Airfoil section ...... NACA-63A-010
Aspect ratio ...... 3.9
Elevator Deflection, deg ...... _25 1971025733-067
- 60 - TABLE I. - Continued
Vertical Tail:
Tota 1 area, m 2, (ft•2) • • • • . • . • • .... ° ...... 15.4 (166)
Span ,m, (ft). • • . • • • • • . • • • ...... 4.4 (14.4)
Mean aerodynamic chord, m, (ft.) ...... 3.61 (11.8)
Airfoil section ...... NACA-63A-010
Ru dder area, m 2 , (ft.2) . • • • • ., - • • • • . • ...... 5.59 (60)
Rudder deflection, deg ...... 530
Mas_:
Maximum conventional takeoff, kg, (lb. mass) ...... 24,500 (53,900)
Maximum vertical takeoff, kg, (lb. mass) ...... 21,800 (48,000)
Standard empty, kg, (lb. mass) ...... 16,594 (34,300)
,i we__IX b_ :
Maximum vertical takeoff, newtons (_. force) ..... 213,000 (48,000) .....
Moment of Inertia for 20,500 ks mass (45,000 lb. mass) and gear down:
ixx, kgm 2, (slug_ft. 2) ...... 385,000 (284,000)
Iyy, kgm 2, (slug-ft. 2) ...... 277,000 (205,000)
Izz, kgm 2, (slu_-ft.2) , , , • • • . • • • • • • . . . 606,000 (447,000)
Center of ,GraviJ_:
Percent of mean aerodyn_c chord ...... 23.0
I 1971025733-068
- 61-
TABLE I. - Concluded
Propulsion S_em:
Main engine, 2 installed
Rolls Royce Pegasus 5-2 turbofan
Maximum thrust per engine at S. L. S. for 2 1/2 minutes
newton (lb. force) ...... 67,200 (15,100)
Emergency thrust per engine at S. L. S.
newton (lb. force) ...... 76,000 (17,200)
Weight_ per engine, with nozzles i newton (lb. force) ...... 16,000 (3,500)
Lift engine, 8 installed
Rolls Royce RB-162-4D lift Jet
Maximum thrust per engine at S. L. S.
newton (lb. force) ...... 18,700 (4,200)
Emergency thrust per engine at S_ L. S. i newton (lb. force) ...... 19,600 (4,400)
Weight, per engine, with yaw nozzle
newton (lb. force) ...... 1,570 (350)
Total maximum thrust at S. L. S.
newton (lb. force) ...... 285,000 (64,000) ] 97] 025733-069 1971025733-070
a_ i 0
,,I / F \ E -] -_/r- _l _°_ i0 -_.o
j • _e-
.__
' " t-' U
q P A
.c _ 1971025733-071
Rotation: Translation:
(_ Longitudinal_fick, 8Mp, controls (_ Onelift enginethrottle,CrFcU, controls elevator(BE) and pitch nozzle (SpN) thrustof lift enginescoliectivel._' Q Latera!stick, 8LP, controlsailerons Q Two main enginethrottles,O'TLandO'TR, (8A) and differential thrust of lift controlthrustof each main (ngine engines (FCUL" FCUR) (_) One mainenginenozzlelever,O'M, C,_ Rudder pedal, 8Np, controls rudder controlsdeflectionat all main engine (SR) and differential lift engine nozzlescollectively nozzle deflection (°rNL+°'NR)
Figure 2.- Schematic o£ control £unc_ions. 1971025733-072
Q '_' m ...... #.' 6r_ier p'ilot-" . , . _¢ . NAS. /), evaluating -
a) Control layout.
Figure 3.- Cockpit control and display layout. 1971025733-073 1971025733-074
160 I I Control limit__ 30 120 lj
Pull 80 ) - 20
z 40 I0 ,_ d ff o .,_ / o _-- 40 "/J
2O Push 12oj-_Control limit 0 47 9I:5 13l 9mm -12 -8 -4 0 4 8 12 Fwd 8Mp , deg Aft (o) Longitudinal control 80 15 40 S _o ..
, -- 0I 5r1 I( _.,omm 80 / I -40 - 20 0 20 40 Left 8 LP, deg Right (b) Lateral control
300 --.
200 A /"- _ - 40
z I00 ; _ " 20-- e / e '° o / o_ oo I"__ _o _
/ ,40 _ 2___8100-60 -40 -20 0 20 40 60 80 %" Left 8Np, mm Right _
(C) Directional control ,._,_
Figure 4.- Pilot control force-deflection relations. ;
i 1971025733-075
I00
.i_ Emergency 2-1/2 minute
90--- _/ -- rating , _ Maximum / continuous NF' % ..I / 80i--. / / 7£..... / / I1 6030 40 50 60 70 80 o-T , deg
(a) Main engine throttle, length of lever = 220 mm
I00
-12,700 / - _ _ Emergency -12,4tom00 " _" - - } 80 / Maximum FCU, deg - 12,400 _ .... ----""-" continuous ' ." 60 -12,000 / / 1 - I 1,000 / ,,o / -io,ooo / l (" Flight idle
20 -f -- Start / >.
0 20 40 60 80 I00 120 -° "_: O'FOU, deg .._
(b) Lift engine throttle, length of lever = 240 mm .._
Figure 5.- Thrott]e relations. :_
i 1971025733-076
Mechanical control Aerodynamics system of elevator r "_--] i _---t I I Stick Servo i I Pitch Stick I signal Control motor I I Boost nozzle Limiter _ _ r -_ +
Transformation Aircraft and integration motion : L____I "- (a) Pitch axis
Mechanical control Aerodyllamics system of aileron
I Stick _-_Servo ,,,r--' I 'f,,.... ,t, -1, Stick I signal Control motor / i engines LimJver I
L'-'-J _LpL---'J _ ' _l I _ IIIII Transformation / I Aircraft I ...... , and integration _ j _ I "
(b) Roll axis
Mechanical control Aerodynamics system of rudder
! Pedal <-I Servo itI -'! 1, IYaw ---_,, - _ 'I i _Control motor I Lever Boostnozzle Limiter I _, Transformation Aircraft I i_: and integration I motion I
tt, , (c) Yaw axis
' Figure 6.- Block diagram of stabilization and control system; VTOL mode...... 1971025733-077
i
JStick q_
--_ Elevator
(a) Pitch axis
,.,_Lift enginethrottle \Electric o_ -
Litrue_ngni_oe(Fl ._CU)_ _ ,,,,..
(b) Roll axis
q¢o Electric stabilizationactuator der
on lift engines (D Hydraulicactuator
_TriCenteringm motorspriningcruiseconfiguration " __"_ (_) Gearchongera, function_f dynamicpressure ._
(c) ¥a.,_xis __,
Figure 7.- Schematic of control system. 1971025733-078
30 Control /limit
" I I 20 ...... / _, / Limit
Preselect values _..-J-Y T, / switch
0
/_ Trim range
-,o i1 /' I _L // T
-._0 -I_ -8 -4 0 4 8 12 BMp , deg (a) Pitch L
'--'_-i "
_.l .._Control limit
_" / _-'Limit switch _: I0 • ' /
_. _, deg ' • • , /
'_" 0 /
"_' ' /
'_ -I0 " : /
-ao_• "/ I :,, -40 -20 0 20 40 i 8Lp, deg
Flsure 8.- Steady-state conditions commanded by pilot control position; i VTOL mode. (b) Roll •j: % 1971025733-079
0 .-
limit
I0 // _ switch _, deg/sec .... /
0 / i - /
"__!l -- 21 ) -40 0 40 80 : 8NpI mm _! (c) Yaw
Figure 8.- Concluded.
"IF X
1971025733-081
1971025733-083 1971025733-084
8 lift engines at maximum continuous, 2 main engines 2-1/2 minute rating Same, but I lift engine inoperative, :5lifts at emergency 4 lifts below maximum continuous I Main engine at 2-1/2 minute rating, 4lifts at emergency rating 4 lifts below maximum continuous 1.4 .. 1 T _1 _AII engines
1.3 '_ "_-. _ _rating \
1.2 _ T l lift engine _ W failed
1.0 _
I mornengine _, • failed _ "_ .9 _
/Standard temperature ,.
-I0 0 I0 20 30 40 TA, oc --:
Figure 12.- Thrust-weight ratio in hover out of ground effect and in lateral _ :. balance; m = 18_500 kg_ h = 600 m, am = 95 ° , e = 5 ° . '_ 1 J 1971025733-085
I. Stabilization system engaged and checked at t = -50 sec 4- o 2. Prcselect tO 8 1,t- 40 Niain enginestoT0%J -- sac 3. Nozzles to 75 at t=-3o sec 4. Start lift engines at t = -30 sac 5. Lift engines to idle at t=-20 sac
' 6. Main engines to 89% "[ t = sec Lift engines to 68 ° FCUi 7. Release trim 8. Nozzles slowly to cruise setting of I0 ° 9. Stop lift engines, I to 6 Stabilization system disengaged at t =+21sec "--'---'8 9
160 I-' . , : , ,4oi. , .. : •>.>. ,ooi--.12oh. ': ': i L.._._ " . . h,m o_ I...... ,,.,,....,-_."T" • , _v I • J- "" " . ' 60":r. ../.= ..... , • , • .'."'"! i='" ; 401-" i_" " , : _'" . : i : , ! 20 '' "i" :' i : ': i ' :'" Oft'' : !.--..-_--± ...... _------, . , ---J
120 _[ - .L_I.J_-II.]_L_.__...• T..-[_...... _-V_-J_.-_.__..tJ-2b"__ , . ,""'
I00 'l . I : : : • " l : " i ' ' ' ---- , " " ' .."' _ '
• ; • i , " ' . i I ' ; . I ' V, knots i " '-!_ " "i' '., ', "'. i "" I " . i .'. 60- , ,= , • , _ . I ' i I , . " , 40 __ir_e from takeoif,sec! ; '.: ; 20 ": - i ..,._...... , ...i.:!. j., ..-' ' i'.
O )i ! : I i : i 'i ' ' : i I " i'" L ; .__1 • " • [ "i.'..__hJ 80 ...... -,-r---r_ ...... _'. ! ' : ! • '. _ ! • : • ' ' _ " • " I • : " ; l i l • ; ' ' ' ! 60 1 . '..%,.. ,..,... ,.....,... ,.. , I. .., ...... I' -..:'_i. , ''';1' I '.. _ ' ',i, " '_."'I = . I : , , I I , ', ; ! ". O=M,deg 40 It-.... I . . . "=Is.. I" I _ . I .= I • ..; • '. ' .' • !
20 • ,. I ! • :.,.. ., ., ..!.. !;_ ., :_ :. oL; .LI , I' ,.].I, ;' I.' _.,iti _ I I ,I
LI.....,_r ' ; --, ...... _ F'_Tt'"-.--I...--'- ,'.j.._r'-,'"l',_J ]-r. i _,"'1_'"', "T, - -r-----_•...... I : . 1, i
,oi-...... _:I..!::.:" ..: '...... i.._.,..j : ...... ,...... _..,. ,-q1"" ,,,:r-_...... ,.._-'-_, . : ,.F",....,'...... ;..: .,_' O,deg 5 U_r-i:-!-[,, , .... .! ...... ,. i'...... {. !".{.!','"".":{,:..l.,...... :. ,' It..' ! _'_i: : I -L.; , . I , I { • ,.. I..-L.--J.J 0 I00 200 300 400 500 600 700 ',_ Distancefrom takeoff starting position, m ( a ) Range data '?j
Figure 13.- Vertical takeoff; m = 21,000 kg, 6-12 knot direct headwlnd..L:_i'_. l 1971025733-086 1971025733-08";
1971025733-089
20 -V " " " ; • ' IO
Ah, m 0 ...... _--./ . _ .... , ...... _ ...... : =10 " " :'" • " "' _._ -20 .... !...... ' , : , , i ....
20 S:(;ble.:su....l) idle "r --. . . . Sh]ble ...flight.: ;die.... O'FcU'deg I0 I..i,t en,]iqe throttle ! 30 _r/ . / ...... ' ,/! ' 0 _ i i , • . , '
60 • ,,_
O'M'deg4200 ._ ...... :"_ ...... i'. .i _. i , , E ,,, . . ,...i i i ...... ,/,.,,,..,...... il .d...... =_., :,, 6o r. 50 • .....,,,. _."_",i'c. r_'. ,.,..,.....'.... .,,,-"-...... "" : I " 0"1-, deg 40 I.....:...... :...... :_.L..-- .....-.--_.L_. : i ! • . 30 ...... L.....i ...... I......
I0,- _. •, n ...... • : _ • , • , _---, • . r- .. !._--r-,_-,-,.,--, ..,-._._" ." • - • " |.. • .= . I , l.. ', ".1 , " ''_i _ " : ; .. . : . - ...... : . .. . = ...... :.... 5[I- : ', . ', : •_ " ", ; ...... , . . . . : . •. :\ • '_k.', . " ".., :! = ...... ' • i • , i • i • : • " • • OTRIM'deg 0 _. : • i"...... ' " !., i i'. .'. i '-.: : ._;..,m. -i,,;;,,--..,,,__ - 5 L,....L_'._L_..."...... 2...... :..J...... ___L.-..i;._..1...... _L_....:...... I...... i - =_.;--.i_.f._i....-...""1".' ,'".."n
%. O,dego • "-,,-____--- I0 _._. . -,5 _ I ' ". ' "
5 _ Longitudinal st,ck-- ...... "...... 8Mp,deg 0 - --'-- - "'-- -;--- • -" - ' -5/ I i I = I , I .__.J ' I I . I. ! 0 2 4 6 8 I0 12 14 16 18 20 22 :. :. Time,sec
_" (b) Time history of lift enginestort
_,_.. Figure 15.- Continued.
J_ 1971025733-090
4OO
h, m 5200.00 "_" "--_._.,.....,._,. .[_":::!',:,.;r' I00 : ...."'"---_"_,,.-..,.,.
...... 140 ,,......
V,knots iO0 _ __ 80 120 "-__ _ .,,._..... 60 I0
ct, 0, deg 0 ....
- 5 ,"" "7,...... ,...... "" ""' "';" ''# -I0 -- "_= %_, =_" "=,Nel _i" 6o •
I II'.gii ...... -- ...... 2 0 .o'-," deg _'',...... , ""-.2 _G'
_G'dais -I f" n "' -.4 -2 "/" ' "
5 . ::.:;.l..lil,j,",i.J,tr',C.,'li]i" BMp,deg 0 .._.,:;--' -- ...... -.L-;,._,.,;=,." ..,...... -5 = 60 ......
_ ", , , • .... , • _ o'FCU,deg 4030 _ ."I.ifl engines I(' o,.__'r(:..'.:,_,:h,,:tt_n.:] :' 120 / -- ,oo
60 '-':: "'...... ' ,_ 40
"- Ax'g . 0 :"i" ! ill: ""[.'.! I ':ill.....""'. ill ,i;'i"• , i _'._L,' 1. : ...._ - _ 7_ ...... i, "i'" --o5 , : ...... " i . , n ...... j
-I.5 . _ i I._ ! ' • l " r:'""" ..... r._'""': .... , :': ...... : _...... : i • .. : • :• • . ": ...... Az'g -I.0 , , ...... I...... '''.. ._ ' :_"_...:-i ... • : ...... i i ,, i -- ,5 ' • : ,,, ,.,_" , , , , , , : I
O, i.... :, "." [:'i_i .... !"J';"ri:'"""1..... ,' "" ' : "'] : i ! I ,, , N ",i I 8pN, deg _10_' I I" " ,'J-_--"-'..' "" i" L'-:"':'i '_':" '" I • :: : , - I , I _ --.-_ i_-'_'i -, n' -2% • "' :'I' '.: 'i .. ! • . .! .i ...... !" "i !'U i ' Time from lift engines stall, _ec (c) Longitudinalparametersin approach
Figure 15.- Continued. 1971025733-091
L'dots I .i ...... ;..eft _ • . "---,...... 0 Hight
tO - • • : . . ... : ,,.=._,j.---,..,..i'_'_'."...... , . ,./.-_.,,,:.:_...... ;:....=._ f .. _ • -I0 ......
5 ...... _ "-:- "-- "-"--;" P, deg • " " _i_'" : _, ..... '_ " ' "-- ': : " - .... o • • __..,.T _ # wll___ " .-:...... -5 ' " ' • : " : " ' "_: " v\'"%':
; 50 ...:...... i -7. :...- ...... : • ..._ -:...i • ._ .i ...... _ 20"..":_."-"---.....-'_,-..--..--- ,--.,4.-'__--,-...,,.., ..,.___..... % deg ....i ! ...--. __:.-,-: :,,,.'..::....-:....,.:. .' -2o i _ ',._.,. •....., :.,,...... ! , 5 _, deg/sec 0 "'-_t_.;:,.=;:_=._..... v ".,,"_;_,_ ='"-_4,,Jk...... -'..--:.-.'P-_ .,,,._,,:,"" '"" -5 lO .@,deg/sec _p.,=,i_wm,.jL__ _ ...m_,_.=,=.td_._. • ...... =...=.==mh.=...... -I0 ...... "
BLp, deg ,o0 ._I Jt" h:,artJI S,l(i ,V,. . ' -I0 20 ...... li_lltli J 8Np _mm Rudder .... a,_,. 0 .,'_,..l_._:" :.;_":' _ , ___ ..-i,.....- _._ -20
I0 '.=:'F-: " I"._._.::_.'. ....:_ I : : I..::,=.."_-....-T_...... :.1 . ;...... : ;_".. I '..1_.._". :.1 .:. :i:': |" .i':::::7!.-..i :....'q..-:.:. I :'"_"_..".".:':: : :'.' 1":.-"-1:.:"-..'i"..:":::"" deg i .."'_,_ " • I .. i %1 "' "::' " "_::"i.:'!' : ::'."':':" "::.:':'l'i.:'l J:[" _ _ :l' ":':':i E' "i:: F FI" ":: -k' 0 "" _• ...... I • "" :'" '"i """"'''" ,.iI ...... I ,',",...... •...... , • I' i ....•.: ...., ....•. ,: ,".:'_:, .I,..i__i,.:,%_i,,,,:, --:..:.E. "" "i._=i.,,':.,.;_.,.,"',-..i.' :L. _i__-..I _I',: , • .. i"...... i . ...,• , i • i .:.:.,_:,."'_ .I_...'.4..... I.:..,.'.'_..._..."_.'• _ --10 i . I 1 .:-..-" .. i .:.:. i' '..::::._.-::.!..i. ":....::.'.; ". _ :.'l.h'i:-. : ".i" I':.%!:..::.' .... :'1 iil ....
' I0 _.,:: i".. -'r.....I • ".T"i ...... ,,.!r.-.-: ,-_.....'-,'_:'-r-,.._".':",,:,"":.._' ...... :.:.,:_71;'"!!."-q_..•...... ! :'":....'_":" '.."FI"_.'_F ":•'":i ..... i I" '."• ' _N' deg n :' .... "I . '.....'I ...... I • i' ":,...... 4..:._..j.... :.'. "" ::": "'" :.' ....L_l_i"_."_'_,_I...... ':::: .'" ".:l "" '." ":. _:'":' .:'.:: ":: I" :..... I,. • :..:'...... :::...11.'.I..:.:.:"...:."L:."'.i"'. :::!:...._" I:!i:.... -I'_ _ - I0 ...... :.""• .h "_...... I "'"'. I .i• " ,.',I" ". '• '..:'.1'." . .",'.:.: '.",.:..I l."J.'":..L'I .'1": ,".• :',.:',.."I::,"',l' " _" 40 50 60 70 80 90 I00 I10 120 _ Timefrom lift enginesstart, sec "_ (d) Lateral- directionalparametersin approach
"_. Figure15.- Continued. ._: 1971025733-092
9O
70 '_' :'"" '" 6O 50 ...... "': :'':!'""" h, m 40 '_-,, ,_
2050 " "-_,_..,,,_ ' ' I0 "'_,_......
6o -- :....'4-i-::- "-.... : g:- ;F:-:_ -- AT oC 40 20 - . . I Oil0 ::- L.J:I.:I,?7.__1it: : :-:-k :....I I1-J-I:...... I t I.:l::::l::J:L,l,#.lJ_.L_'qh '-
V, knots 8060F_ _zT'TTFFFFFlqTIFT_'-_']; "_'_-FF[---FFViq,, 40 ..--"_--._....._ I . 20 : " "...... "."T-"_t : . " O ......
, , , . R/S, m/sec 4 "_¢V' : . . : .. ,"._ 3.:.:,"i,,r.;, . 6 : ! -I.5 F_,Ti-l__ :_T.F _. N Touchdown.--:7 Az, g __3"__*_'1_ --I'OI--.5 o L" ' I _. __[,' J l, _ll[ll,/l.I ,, II ,--m_r-_.:1 '..,_.'. i
70 i;J t t I [.ifi e-nginet'-fir_til-_[F I-I-]IJ% ':l{IormalIT] 60L-q-j...... :.TrW_ttt,t...,._t_.i_- - f11.-7_1:' -operahng_,. .. E1 o-FCU,deg 4o50 I.:..rl' I:]Eq' Y/..I,'l-:-'lH:'_LL_L''7Pt'_._U:LE.L.-F-q'"L.,W ...... hff. i:::
6O50 I ,i I-IIJ[irTl:ll::l:I _:.l!" :lil'! _I"tI,i '1!':•:-_]'qi;" liiTdTi11': ' '[_q' ;" 1:7
o"T, deg 40 LiJ---l:.._.Mainenginesi , ll_rottles.:. : "i ,I !1..... ;2 so!-ih: ]H _r:r:I:rL_L__U_LJZ_L...LIU_.:_":.u _, J,2oooE__i...... :.F----_JLI-L- .+4-i(_T.I:_.t.T--Ftt l :t O'M,deg 80"'.... "'iNoz_":le._!ever_i' : ]'': ; :'i: !!l i, .'. I.,.-.J...... 60 ! J...... "'7:' :..,. , ' '''I '';=....., ,I':
,0 [T;,m ;eI_'TFt::] L '__!t::[-[ I:F_t-[ 1!_] 5N,:NI ,_t I : ;;,:,:tIt hi,:;;,t ::_J.:_Eit:it_:l:lHJ!t 8, deg 0 "TII-_!...... I I-I I ITrrr,. . i r Iil.... rr'," ", .... ._,'q""r,TrT'. ..rr l , ' -51,;.r",',',',',',_,:!.;.;,i :t !-ri"..'.! , I l' ' !_i : ' "lJ_l" -I0 1 I J_L.L.I_L.J__z.__L_t1 . , I_J._J , '1 ! ' l I00 120 140 160 t80 200 Time from lift engtnestart,sec (e) Flare and vertical landing
Fisu_e 15.- Concluded. ,-,_ 1971025733-093
Figure 16.- 12 ° simulated IFR approach starting ].ift engines on glideslope; m = 19,500 kg, wind 15 knots from _ight, YG = -12° ± 2°. 1971025733-094
_ooo .... /_......
"_;-.__.___.. h,m 600800 400 | _._,. ;r O,,.q,, . 200 / / --"------t •
160
V,knots I00 ...... -.7"--.._ 80" "%" --... 60 I0_-...... T ...... :......
• .'-..,,,,,. ",.. ". . ir -I0. ". 1 • " I -15 ' ,;_; t "_.,,,,,,"_ .,.A .,,,,--, . "-.. : -20 :
I, . I . • • EGt dots o- f, _7:1_.._.,_...:-'_*s=._..,_'...... _--_--z....Z o: EGdeg' I .--'' • • • _t4 60 ...... r .-- "".'.i) " / I :. .'; /'-"'O"J"'k,:....,"_--.--", ..... \ "",,._._.:_ o'FCU,deg 4;). I • , : / ...... 30.. • 'i ' "_.._ -- ..... • . Z " . : _ • • L,f, C:,,,I _._ ,) (..,; (..... I .,C., .£f 20--'-"r . _. . - . '. ,oo, / _ : l,:,.._2_:_i: • . • :'. .... i:...,..Nozzles.....t{) o;;[:,c-:.',nset;.::qgi • o"M, OT" , deg 6o,.-_: _. -- ...... _-" i:! :." _..... : : . i :.....'. Z i. i.: ;_,-_': r "_ " • ..... _ ...... 50 • _ ' "" ...... '' ^ ' i"= _t 40 I tJr l...... • _. ii r,,e _rr : - i . • ,. ,,....,,,,...... -I,5,-'1 1 -,...... _--_------r . , ,-,'--'T--_, Az, g ''' . it • : ..,...; :';"• • • i',i ;':", • _ -_.oF"-_.'_'"e_ _----_. ; -.5_ -- '-U .v..' ':'' : : :i . i:: ;'; ....i!i
5!--r-; I 7! _ .... I ...... '-'--'-T".""aT:-!-. Ti P_' Ax'g 0_,=_=,_-__ . "i,.',...._L.,,.': ; ; "' !'., i • ,,• I • •i , ..... Ill.. ,.... .' . , ....-___,. __L-.,_ '_ -.5 ...... I i , .....-: :1 ".', t ,. I ...... ".I ...... '"'-. " R.. - " I" = I- - ,--'-- . "r--r ' ",..... _" ! --_ " " i • "'l ' • _P_i deg ..i : '. | • ' " , .,. ! , , ...... , ' , .I • ' U I_ J,,. "= .... . ,. ; ' ...... '. _ . : :. .: .; .,
-,o-.._:,,.111,1 ;"?i_,-, , i _--'_--:•J____ I - , .I_-,,.:_,, _.i • , ,, _20 [_i'1'i: : i . Ii.I , ...... [-'.7].!.'..'.'. :....'.'..:")! '7. I ;.! -10 0 I0 20 50 40 50 60 70 80 Time from lift engines stoHi sec (b) Longitudmolporometers,nopprooch
Figure 16.- Continued, 1971025733-095 1971025733-096
.... i. - .... I • --'--- i t --i O_ • N ' 0 , _o L_/ ,:.---- / J "r- O
"_ ' i ...... :...... • 7 '' 'I0 e 0II 0 !. ' / g) 0 _-i
Eo I ' 'I °I <_ "_ ,i_I • ": _,_( -- _ # 0 / "'" ' I o c"= l" ' / o _--- r-4o_0 I rd [:. .H 0 ,¢1 0 '.- 0 _ N .,.4 ' 0 "_-,-",, •\@ I uz_ • (;]; 10,J ,.I-Jg_._ _5 o o \: .I _ ./ : ..jo _:_(D ',." .... I_, ,, ,.. _ n.._,.,._ " _ .... _zo'- ,. , I:. ' ,, _f- "_ ._ 'I 0 "" ;" i'":.... . ": !I _ ...... : ?' _...... o
_o_O00 0 ,-I o
t I o
:-. \ iu'4j_' I, i_o _._
_ , i ...... 0 __ u o _, !I I_ i ._.. , • ,_,..., -" 0 • q-N ' _" k.,, I " I.__ I i ,' L ...... _L ...... !...... ! ...... _0 OJ e_l 0 -- (NII ._4 I" I I j _ ; N . ,,x 1971025733-097
I . • "I ' ! i i
, ,t , tI ' ., /i.!.i / ,.d.•
Increased I...... i_• ' '] " ...... I ! ;I ...: '_.i,'-/'-.,,...,.i; ,,,,,'<,..._._).,,,,,;., ,,%• accupwardeleration .I iI • ,, , i •I .,. ./: , ... ..'r...... q. ..b(;._.,l._.,v , :.....:,-,_.T_ , ". ; . _,. • .].," i • (.)u.) ...... ' ' :. . !,...... "/.i._,'. .,,,d,-,....._; ' ", •". • '. .':;:i l"', i _:I " ' ., -Aaz o ...... J!i...... --:_ ._-.i...... !...... i.....:---_--:-.- g i .,-"'/• " i'. _/., ':i " '"f: ' i _. ! : ! .. i _ ..:...... I i, :,._ .... ij • -.I_ ! :.,! : i .. i ' , i , •
...... I jo/,! ;. ,I ...... ,i , •: .... ,=i ,i,". ....,i I , i _ J I' • "_• •:..,ii"_i"!.["i "::i:._fi_...... ,i. ! i_j ...... --. CLaq S(,&e)" AO - 57.3 2 r ! • I : , ! - i- -i..... ! • ...:..i ... • / , I i I ' I ' , ' i ! i | .!. ":" i' :'.!" :'.:.i .." :... • :'": _ , ..!. I '""'1 • .. _ , ': • 1 , : ..... , i _. •" ...... "'" : ....I ;.:.,l---I-; ...... ,...... I'.':'" ' ":": • ...... • • ,:'" I; I ...... , ' '' I" :r" • , ': " ", "1,, ",..... " " forward , LJ i,..._r=:-L..._i .';::' ,t"!.:....!.'::lr.:'...:.' I:.:... ,_:.....' :,1 '. ":!..• • J:..... acceleration ...... I !"- ,'. 4, :,,I; .. , _ :,'i.....:• ,I ..... _ .... I' .I ' i : "',I_;,iJ'".. "- .: ;. "" , ".... • "" ." . . :" . • : :I:" "" "', i.....: .::'...... t"'i': I...'I _.":.-" I.:!-:" " •:._-.::-,:.F---:t.:...... ":"...--.I :-I ..I ' :..'.I • ,...... dvx I ;. .I _i_.::::..1:': !_ :l: . :...i.., i..... ,. A _ 0 II:--m---4=.:1..._:, !:"F 1:_i.:_F:.;: .&.,.i1.:I,--.-,..l:u--_-F-_.:-I_.-;-' -:r.i : ...- ."I -::.:: •I --g I:I ':' ':.'"v "i '-.:! .1r:.':..''l!"1-:: : ::r"i'1::_ :'._:",i-,..:,."'..,..i'" _. ";:...... • ";:'" . ! "'.':" "1 ._•:' ....• , ' •I . .• -. ,:."" ..'...l::_'.,..:•.:. :.....,_.i .. • ",.--." I. :" _.-' . _:...... " ' ""i,.":...... : _I...... -, I" q" • • i . ' " . ' I I .i "N_. : : , : • • • ; ,:" _ " , il ' "i : I"';"'I .... .I "'; .I ...... I'. v: T_'.;."': " ":'" .'" ".I. "' ; " ' : ' ': • .., ": :I ":" "'! .... :. ' : "! .'. ; I '.• ! -.I _ ...... i: ...... t--'----I--'.:".-'_-:::r-:t'"'i ""_.':":-:+"';' :-i:-':'-', I:,',,: " i,, ,;, , : ,,,i ' I"! ,P'_,: I , r, ; I I; ':l " _ l':I :" i- :::,._T!I • ,_.::: ,:_ iI:.:::.:l::r. _t IF_'_:-:.,'";;....., --:-.... ;"...... '"'::l" : -'i" • _ ' l"i" ; I _ I i : !' . • ' '"_, , l .... • i. ': ... ;...... ::..:., • "_L',/ .21 ...... I,..• i ,,,.:, - ...., , • ,l I, ,, I J| ..-J "8 -4 0 4 8 12 _: AO, de_ _ (d) Pitch attitude --
Figure 17.- Concluded. 1971025733-098
/Start of pilotinput
O'Fcu ,deq_ ..... :....."...... '...... " 50 .. "
2 12,000 ./...... - ...... NI+Nrpm5 , II,O00--13'000 L ,..-_-..
-L_az 0 " ...... - .... '- • g -oJ I- "°'go ....'--......
V,knots 60 ."'-" 48o0 ! • '. 5_-]7.._ • , = •, a,dego AVz -5'...... I.. " .-! ' ! ', ...... "-z_y=-q--x57.3
0 ---= 1 ...... i-- ...... T'--!" O,d g : . ' _10 i I • •
"-_c=-- _ pN,deg -I 0 i1"_4l "-- ."_"'"=-__" .. : ..=..:.l.... :.. • "• _.: -20" i l ...... _ .. ,. ' .
• . : : .... :.: .: .... : • • • i .... --_V_P 2 I' 1 ,, ,: _ " • I I
_"::_<'_'_- ,- f-Z_azdt, II. I = ,_.= w. :! !..,.._ .t-?.-p.-..•_...... _: m/sec 0 = ' ' _ • • " : ...... -_" I I i • ' • I • = ' "...j " " °! i .... :l ...... ' ' " ! "'=i...... • _: "I I...... !...... ; ...... '.....'1 ..... : ......
--- . Ah= |0_.r_--r-_ ,.---, --r-- : , .,-.---...... /I dot EGaI 7°'t"I°at h = 200m _-- " U L--'_"_l _ .. ':...... --I ...... a " s fj'_Z_azdt z, ,.,t ./i... , ,:" , ,.."i i..--..i: .-=_/ ;,._-._: -'---.'i'.'/-_ (._ full scale deflection) ,...... •I ,l'"i. ,..... ,i :, i I : : :. i .: : m -,ol...... i: i...... I :...I ,.I ......
.. AVx= 5i' ,. : .i,. , = i, :': i= i ...... i I0 ,/'Z_axdt, u^ ,-=i :l-,--=`,,-,-,:_-=_i"i,'!--=,--"!""::._i--.--l'::i"I,-.---: .._i""i",.,._;'"!"",_"_,:_ .," vr_AV, knots _I "l=;.i :i'i..!.=. ; : '. _ i , i ;..i I0 r_T_=_''' m/sec -b" J--'-O...... ""'2"_":"':""'='"""4:'"-'=...... 6 -- ....8 _...... :""10 ...,_ Time, sec
,_ (a) Lift engine step with fixed attitude
,_., Figure 18.- Longitudinal response to different controls; Yo = -7_' M = 19,500 kg.
N ]97]025733-099
,,,.Start ¢' pilotinput
40 .... i"
9O I '
70 ...... i...... :......
-Agaz_ 0 -+'/- ...... "...... -.I I
Aox i T o .....",,,, "...... 2i.;2:.-.:..-..i:-:....- -J • ......
I00 r-_ : . ..
V, knots 6080 "i !"-" 5 ! " i ...... "...... ZSvz Q,deg 0 i ..-J__i..... " - ' . ,.;../A-y= _ 5.73 .,,,._.,,,._:,,_ ...... _ ...... ;,...... _' - 5 • • _"""'" "'PL'''_''J"'_i"I'_'I"Ple"F_""' '
0"-- ! 0, deg -5 ' • ' ' i • , I -10 " ' '
o_'pN,deg 0 .----,- • . ,_. : "- :!' : : ". ' :i .' . /..":
21 1 ...... ', L,, L-- .,-F----. " • "'" " I " ; .,._=i -'m" • ": • • I ! .AVz= I I .... ;,*'" " ..... " ": : " :" "'""" J-Aozdt, . I _ •: , iI . "" i ' : •, ,i • m/sec 0i. _ . i . _ . I ' .i .. i -I ". /i i...... ; : ....',.. ' .. i...... !
Ah= 20 ---I---T- ,'-- "'--1 ---,,-- "1..... :":_'" • • '.....: !'" I , ": "'1" ; i ..... : I.....Itlnt E^ nt "P:I:I e(:it ]'J'-Aozm dta, I0 ,-- I I ; . . _ : ....,. ,,..-"i,_, --'-..-"_ :' ,,L.-.,.v_,",e,, ,...,,, i • l , • • I _N _ • ' : : "
._V X _-" 0 rI.- -1-.] "=,,I-" =-+' .,,.ii, !.._ .._;.'_: --."",L...... i... , . • ...... I • I, 0 ' • •: " •i" :i . ' _4: ..I",,-•_ i-,-,,..l • I • • .rAaxdt, -5 i...... "{ i ...... ,...... ! •...._. I.... -10 Z_vx,knots I "/ : l"!:",":"l'"i " "! :i • :I m/sec -I0 ...... •...... ='=-20 0 2 4 6 8 I0 : Time,see _ ,
'_E,; (b) Main engine step with fixed attitude
_.,. Figure 18.- Continued, "
T_t,
, , ,.'j" =-:r .,_ 1971025733-100
Start of pilotinput i20 l IE'-- ...... - ..... -i....i-i...... : , , -
-,_"g -.Io,.z.....J I"-..i...... )_...... ! .... "
i i I I _axO...... { "L i...... :...... :.... T-,1' • /':-:.;;- ...... ;..i.:::...... : .... ::", ...... j 80. , I
V, knots 4600 i.i-_,i..... i...... i ...... _'-T.-'- ! • j 5 _._q_<__ , _,_e___o!...... !...... -'...,:;\.-...r::..-..-_=_..::]...... , \ _Vz =., =
8,deg-lOi-151IN--...... '_lI...... ,., I.. • I :....._ t...... ! _pN,deg
__: -201..... :..... I...... I .. !.. ! ...... I...i..... I...... :_._1 ;"' 2 • -...... [- , ... -Z_Vz: I ,-. • -. ....---- _ "r.=:._--.---..:"T'." = i" • _; I-z_a,dt, / _ [_* "_' : : " : , m/sec OF.-----T.,T ';. i,...... i.! "i - I L.,.-',.---L--:---..L...... 1.... :.... I .....--_a__ .I 20, , 1" • : '' ..! • • ' ' ' i Z_h-otaD- . !....J .;=,.I ...._ _,__;-_dot_,_Gat7"+)° - j,j'_azdt_, ), • ) = • ' " " " ; ,_,;_". 1 m 0"r .'i...... [ ....._ .L.,2,-4-_._. T,.._" , i. i....,...... j. at h=2OOm
" _vx= I Ol-r-"-r-_ _' _ I ='" i r" ', 20 J'axdt, -v : '"'r "' ! , "'," . , ", ,'_,,,.,.'. _" , OF-I: i Iil.. , ...... _;,._'=J,'"_ ...... -j I0 AVx, knots m/sec o...: i -..--__., _....I...... ,. I'.. J:..!....!"....." [_, /..' io _ -2 0 2 4 6 8 I0 ', :_ Time, sec _ (C) Main engine nozzle step ' '
;I Figure 18,- Coatlnued, %,
./ 1971025733-101
5 r ...... [ ...... _ ...... i " i " o...... "" ,-:; ...... : ...... O'Mp,deg 0 r..... -k ,_," I I%J" ...... ! • .. : " "5" ...... L......
.... / r,,. _ ,l \.....,,.4. . ,,...... -- | U Ir'Y-_%_'M_"_k' ,/ "-al..- -.m.,..._ o_L...._i.J.,v.._,..r-_ _ m% _1_ .... =_ _pN'deg I I_/ , , "_' ; • • -20 ......
5 " "j' "_,. _.,.,,_ '"',.,.'%_,_.% a ,deg • :I ' -'_-. e,deg 0 _%_,e_J ¢-....._ .,.._f,.. J: / ...... • -5 __"_l _r,,: // .....) -I0 _, • ,/ Ii_'i...... 't..",,...... _ .!.-*-,:'-,,--,_-...... ! I /"--"='-.. O,deg/sec OiL...... Li_ .V , :..... , \ , _ ...... , '.....,
,I " • :/" .... "" .... ; ob • +. ,. • ...... : ,,,," • : ..... ,_.,,,;,,,. : g -.I / / "',,"""" : '
Adt 0 "" i i
g _j i...... I...... I. J ...... I...... _.:".':'.-:.:t...... ; i /' ,.
V, knots 6080E_ .i...I.._i'1 ...... "...... I.....".....I...... I ...... " :....;I.__,.....I.... :_...:.T-::.....i"t_-_-_-_-_-_._J 5 .__ i i, • : , _...... 7-"'_.....:-" A'y, deg .... •...... _ "! ...... : • ,,r _ . • . : r.-.:._.._=.:.:._.d...i--: ::[..!.::i_.:;...... :...I...... !....-.i
0_"1_'_1 _ _,'" ! ' i • • " • ' " • . " :* ...... _ "/ :":' :'""-'_'-_.-"'"'i ...... '" "'. :.... i ::" :": _-_-dot_I AGe,at 7o +_P AVx m/sec -I0'-r • ::"/I ...... • I : _ , • "_.k'i,.'_ _...... ,...... :..! :.. . "'_•'_• • - ' -20 I- .; / :...... I .-"-,.,.-,,, :' ...... h =200m LI'"!...... I' I._-.....:" Ll':.. _'...:: ...... I;'i .'...... __J," "".. "]'"",--' '4"":'"1_._"'. -_J
I ...... ;...... ,.-.._ ...... ,'if...... ': "-. _._;"._..•i.....i..:--_ I0 '/ ih:J_":-"'e"_.'i i::"; :" '_" _ _ _" Zlvx,m/s 0 :---. 1"•'_i ...... , ,.... _,...... ," .... • :n , .... ,' : :":,"-- 0 Avx, knots 5/i'J';'li f • / ..... I ' : ; I • i :. " !...... ' .'..._,i ...... _ . I.-:....=. = :• i
i...'- .! .1,. " i 'II 'l__l"' Illl ' il :i,;' --l, " ' 0 -2 0 2 4 6 8 10 12 14 Time, sec (d) _ttit,J&. _*<,,fixed throttlesandnozzle '-, /
-_ Figure 18.- Concluded. -_ 1971025733-102
2 . . _ .... i"
,'liq,I " • ;" /"\ ..,,_.._.=,...., 4G, deg 0 I .Io, • ,...... ,... _ . . 0 I : .• . " ,_ . "
n n _G, dots -2-1Iy'_f.. ,.. ". . ." ,_ .
40011 ..... -"_- 300 i ..... "'; .... :
h,m 200l i...... _" i.....-.... '_ ,...... _ .. "1"" i 4". "i i00 I • : : ...._ 120 ...... -,_. .... , ",_ l::rspe(._{1 n'lCr!,;;.._(i (IU(, _ t1".; nozzles deflect,on I00 " ; • . • i. ,----'-- • V, knots _k..,._ . _ .... _.,%.. 80 " . :-.'_,._. - _ ...... ,------q.'• 60 ...... "
701 .'--" i.... _,r_--..----.----_ ---'_:" .... :"_ :i-:...,-2-, ._._, I : _i .... Lift engines to normal: %1 _" o'FCU, deg 50 i • ' operatinglimt • • • i . := : 40i ...... 8. • . _ : ". ' . ! ! "
50 _,_:., , ,, :!:___! : . .: :: , ' ' -:::! ::':•L:_.:-•:::":.:::';:::.:...'", ' ...J O'T,deg 40 i : ...... L " !
" ._ ...... _ ...... • • i 120 I .:: =i =i" J .:.. '':" ".." .'.fl_. ._ :.... i...... I•rl'"': .-- .._,..._--._,., .. . . : _. """ I00 I.::'...'.,...; ...... /,..:...,:...._,... •...... ,...... ,...... :.:..,..: : : • O"Ml deg II :r =. ="=:.'i: Nozzes _ i=: . '. • ' • ; i ' . :.. 80 I:i:i::_.:... :...... i _ .: defle• . ctedl _.....= : = :i...... -.....;...... •=...i...... ',.. .i., I'" i-:"I., i u !.' ....i."_ "'_ : ; ! • ;':":.':": .... 60 '---'--'_" •...... "...... '.... '. "".... •...... '...... ' '
' • _'-"_.,,_ ..... ' • ". I =. .. a, deg - 5 I.i _i ..."&.."_,_:iJ.-:i..!.'_[.i..:'-"I.,..:_,,,,i. _;;.,-_"!',.'.,'MJ,-,,..:-...._,'_.-'--...,,_" i_:_J_.:i...t!,',t. I:' '.'"i!.T'....i ".i• -IO I":_:i ...... !'i!; I ;" !':_.l.:;,i
O_ n r..i.., i " .... i _i ! • ._:._ : '.i i!:=!. ....I I"LI"'_ "'" ;"," " !': i ...... "" "." ." "." :' : F..= .... ! ...... I e, deg -5 I-...i.-._i.,_.t..' .,::r. ::.,"...... ' " ".-...." !....1.,,,._' " :..• • • "...... '{ ....._'"'"i.... ' 5 1 ...... i...... ; ,_, o I-..,,__.._..--- .-_.._._.___,...L _._."'Z--..,...... ,....'___ _ : 8Nip, deg i -5i • , ' = .. I.o_gitudi_la"St:Ckl : u : I ...... __ ', ; _0 o 60 zo oo Timer sec ':.
Figure .1.9,- Cor£ect:Io_ _o_ 1-1/2 dot: lo_ e_ro_ a_l:e_pt:ed by ltfl: e_gi_e81 ' _'_ ""G = -7° +2°" _;: 1971025733-103
2 • •'1't"igt'I " i .i • . !iI. '. ,...... •.. '1 I .i : ..... :._ . • . " : : : : • : " : • .. deg •-"'?''4-.,,..".,._ • " : ' " ' .j,-----....4-. : _G_ eG, dots 0 .... _.... L; _..--'_,-...... :.•._,.._,..-::;,"T,_.._.._-_.. • ..... T•...... - : • --__: .- .... 0 -I ,:,, il.o_,• ,: ..... ,, ::i " _• ' : , i. : ! . i : _ .... "...... I -2 i.. . I .. : !.... , • : •
600[.,._ ...... !...... _- ...... : i... ::..: : _ 500 :' ':,'_'_<'i...... :__' ,' ....:,,,• ' !•• " • ._'_4._: " . . " "...... '' " h, m 400 _ , .,""'.-:-4_._,.: , . • • i i i _ . ":."",,4_.',.,j : :..:.:.... • ...:...... :.. . ; . i ! "''"_,,,. _ . :. . ".::: 300 • ; ". .... _ _ ' • ! ..':"_-L_i . 200 : : : .... ' :. ":''+.:"-,-...... -
120 :...'t_: .. : . : ....: • i • •: • • : •.... ' • ; -_. i ...... : • • . : I00 : ....-_ .,_..... • Airspeed IncreGse due .....i...... " .... _ ::" " :: i ' I " ".1 I " V, knots 80 • :. i!. : i .:i_i • •!o .inozzl!es !.,defl:ectio•,n i ..'" :_ :.... _...... : 60 ...... ! ' '_' ,i ' • ...... i ..'r--,, •-,_.'.-,!.;_f._"_.i .-.,L='":.:._...._...' " ". :i. ....:r '._. • ,I" • .'.! i: ;. :l .. ": "'" ! :. _I ." qI ...... -'.l...._.'_ "'" . : • • I. "." "• 40 ! _ I .. ! . ' : _ " ' . ! .... ". "'" I ...... , ,'. !' .,:!.: .!
60 .i: " _ :.: i ' .._.,._-.._,_ ;..---,-_-_._.-,,_._ _. : _i :::i: ; :. : • • !,,-'_F_"- i ' -_."" ...... :. !'_ ' ': '. "..::; : • O'T, deg .-..___- ...... , .. i • ,.::.i. i:.:.i:.:..:._.__:..::: 40 .. " i: :..i i.. : :..... !:.!::. : ..i:. Main engine throttles...:: 20 : .... "'"" !'''''''': '" ..'" : ...i...' ..:...... , ...,:...,..:::.:
120, ..... :. ",'" :. • , ' , ".,': ,. / _-_,;-_--E.'-."I'.':"I: '"."i ':a".T...... , O'M,deg I00 I:._.:_,..i_::...... ::I...i_i :,.I:...... :Nozzle,,:i...... _ !.:.,s:._.....deflected.l.• :.....-."r.::..!...... ,_.:l_..-.4-_....]_-J.b:: ___..-.._ _.-I.:::::::!::,:_: I...... "" tI ...... • : : ...... • ,, :.::, ..::...... • _.._I_., ".12!!.:, ...... , " 801...... •._ :'..,:..,'...'_!'I ...... : [...... •" •: ,I : " .. 1...... •I" I...... " 1...... •: , ..: •...... ,.I ":: I•.".L...... I: I : .:.. :.': .... • ....| : :: .I,..•
0 I.:" I "'.1 ...... • • !.. '_,_ I'"_'": :__i": ::' i_l"'.';:ii_...... ii:. .' ' il.._.i..A'_ii r :_" • •! :.:,..I:. _....::..i. • • " " • .''. :.::"" • .'" .:" ":..... • :I , deg -5 _'...."",:;:_____;",:",;':'",",::::;":"= ..__:iI' :::_:;;" '":":: :".....:'"'::':: :::'""....;' • '" I ...... _ I-;: _--_:=_:::I _:= ...... I '.....' • --tO I'_.:_ :::::;...:. I :_..':.:;.:._" L .:" "_. 4::..I .: :.. :F::.!.=:,: : •
l ...... ,if • I:.:.,:.,1:•I::::...• : ....,.,...,..i..:...... :., ,..i.....•.,:.,• 0 •..... ,...... ' ' ' ' I...... " '" ' " ' "...:" "'.'!:!:! ...... :: " ;: ":"
i ...... :"" "'l." :;.::': ::'"'....I...... " . .I: " I::...... ':;: :: I ::;I!".'i'l':: :. ':"'"_:l: ....":'!.. • ::,. ..';":".:.'....: '" ;..:_::....: I ::..'.." _I'':'. _ "•!':t: {:i ..I_ I "ill!!.'if;""" ".. "I : !': : : : : • : : I " I t " " ::; " ;" :'"'=:' "" : ...... I.'" : ...... : ::" "::::"!: ':L:': . . :: 8, deg -tO-5 _!.._ :-I._:_-"_! :J,':1...:!._:_.,:: .:.
__i_1_u_i_@_!_i_i__i__ 17. 8 MP' deg _ _tgtll tttUtth'ttt_tttli,llt}lftt_tt_tlft_t!_t_l_,iii_, tttt V _. t _, _i_tlil_ttttttt!tll _lllfllltt_ i_itt,_lllTtfilllttllttilili!l!_lilllll_l[tlilit_l_llllt_ _. ,__i__B_i_i_'i_i_i___T_'']_._{_ _i_t_ _ _tl , -" ___i__i_i_ _,liiii_i_n_ii7;_i_i'"'_'ii_"_lt_ l ' ' iI " ' " i ii_ ..,,. I'tl_lllil%llll I%111i]llllt'_l_l'I .*_-%-.,., -10 tt_Ht!ftt!tti,_ttttt_ttttttttlt,t,ltltllt!titt_ttllfl_t_il!ti,,,...... 'F.',,..,.,...,.,,,,,,.,,,,_ <'< 0 I0 20 30 40 50 "60 70 80 _ Time, sec i_ Figure 20.- Effect of main engine thrust modulation on glideslope tracking; ' TO = -70 -+20" 1971025733-104
20 ' : ...... i ...... _; • ! , • 15 • '_.... ' .i . :.. : ...I '.
• : . " ' "_ = i .I : ." 5 ' •.... "_ ' 1 ' " ...... ' o ' : __,.-_._.l__.... _:L__J
0 : ii •" ! !F' _.." _ . " . ' Sinkrate, m/sec I • -; ...... ;___i : '. II : ....i: 2 • _ :; l . . '' : " "__ I I _ _ ...... _ .!.:J ..i...:.' . : " 5 : i I.._ .....
4050 .....' :J: J i'i" :...... ! j ' " ..I_]r_',_."!...... :.../Q"il'_Lo:'',_'f,.,,,f...... " •...,,..! ' •: ._ " . '1!..:x"'_'r_."_R,ahti- '..,..• ...._.., : i • " '. . . :'II. "_.1 'Y'_,- ": "'i
T I ,°C 20 .-...: ": . : ;....i. ! ' 7,/1''/"i ......
'°0 i___,_--._-_,,_, [, ,,,,,,,, ,,
, -1.6 I -,.4 ii -1.2
Az' g -I.0 ,...... ___-_ .. : i i.. [ SuckdownP" 1 o8 " "' ]. : . m .6 . • , • • Touchdown= ,,,d .""P '
i,4 ... | ...... • .. ,.. ..
. ' - i : ; : " ': _ _, 50 ...... ,...... J...... | , ,...... ,.-..,, O'FCU, deg : ' '. i !! • :, • .: ! : 40 ._'Li ft enginethrottle_--'-_"_'_: '""i'_ i!"i!__i':!:::!!!!!]J 20[.--' ....."...... • , i...--..j _" , 'L-- I0 8 6 4 2 0 "2 -4 ,_. Time till touchdown,sec '.
Figure.21.- Lift loss in vertical landing. 1971025733-105
2Oi-rT, rq-r" ,,%,...... ; ,,,,,r_. I T.I,I',l I,_ I _-r-l'r.-r_i I.i I I....Ii /...i'r ..... T ;-i3-"T-r,.,i;;...... _-----F-r--I-,-['. ,.,,llkj.: . i 1"1:.
I_Lp, deg -ioOi.I,.. _ ...... ,_-. rI_T .__, . i.., i.i I .. t'. I .... _,.]
20 1 ',.,_-[,,R.dder pedol I , 1' __-T_,, r-_-;-rr-<7_-rT:...--..q"Tm..r-=-I--r-._Tl - ' I ' , i _ '- .... 7 "i-l _ ,"!"2T_.,rrr...... r"--r ,,'_'I-="_,:.,'T''''?-., ,,, _,,pmm. -20O1<1:-'' _i I !! :_; i,!ll.l: i.,...I., I i.i., !_',.1:..I..:.. i, ...... __,...,...i: ! .:_: ,.i
_, deg/$ec I0/I-TT:--r:I__T...... L_T-I1 f -IO [I_L].AX_ IX-J'-Jl LI I-L'-J
,o_LI_I1_ILL[LC]:]I:/_-[t_tT/t:I ....
AY' g .5Of)_-- . 4i-iT.:" l-rI i i _TI:] TI, Ft._I::I_T_... .._...... T:FI__ _.._..,..,,..-_..:=.,,=,,.;...,, .,,,,,._.,,_ . . ,, -,5 i
O ,Or.rr-i/-L]-i-Ti-l-[-[_-rT-{Tfl-f_-1-_! _.[#71' " i ., ..... ' ' ' __[L_l ]_L;.!.[Lit__!]-[111]]I.!L:Z"!!f[;"t •;-. -. ,. .,
• ' ,,,l,,de<_2o___o[__.. I!!l!ll:li_t:iii:.!_it_i..[iH._l_.., I__,]_L,..F!LL_!....111.[t,!J-I_Ll!_l_ LI!.q '..II'_-')#jl....,oh,'.o :' " :'" ':"I . " "•. , i i " .!i 400 2 4 6 8 I0 17 14 16 0 7 4 6 8 IO 17 14 16 Time, sec Time, sec • (a) Yawstobiliz(itionon (b) YawslabilizQticnoff
Figure 22.- Lateral 8ttitude steps w£thout directional control inputs; V = 70 to 90 kno:s, m = 19,500 kg.
A
=_ 1971025733-106
, . i. • i" ::.. v • . : #r • • .. • l . . • •
.: . . ! 4 "'"" I ..... • ' • ! • : "i ...... _ _ :: :
' " i . ' '. _,deg 0 ...... • ; .... •'...... :,j =i...... : . • , •
• : • ...... • . . '" : ";
-4 ...... ,": • • ...... J '.
-12 L I 1 L
40 • .: ...... h ':" ". ' ... ' : ...... i 80 Right r • ::. : :.. : .:" _ "' • I... :'.! : I • • " ' ! " : " .. '.:.L",..!;..
'" i • • " .:.' ...... i._i .: ",_.•,..:'...: ;zI40 |, " .... ' : .... • :• '' • ___.._.'' T". I"" | I : i_ :"": ...... : _.,...._ .,,?"_. • ". . : • " • : :i .":.4_-_'7.!:. '._ _. : "
(_Lp,deg 0 _-_•-EP. .--'--'.._,,,,__:.F--_..._j.. • _-4:P; _-_.-.• .F : ...-_ ...... :,:"...... • . 0 _Np, mm ,4_,.--r_,._ " : • '" "B • ": I "*:". '. i. . • i " : .i .. ;JP' ii'. -20_-: . • : • ' i • : [':'i :" ":':..'":""l'"':i.:i":.-.-40 , !i _,': L',, ' :," i'::'!: _ , " ;, Left -40 _L_ .'_-...... I L_._. .".; ._! ._:"":":..::""l.i_ilii . ". [!._'".:i'T_i..i;.,i: -80
F..... i .... .: i., , .- •• I :." • I ;!. • i...... ::!. • 0 FCU.L.-FCJD ii ..• Right roll IO ',...... • .: . , ...... : • • •.-.....i l ,..--; ..-- ..... ; ..... !..... I ...... ;'...... FCUL- FCUIt, I:• l" .....' .: _l" " ' I._,': .•..:._I. .._"i" ," i'" I...... :::'":....:', • a• ..r.'s, ...... ,., .... deg r']•L..._..|...... i."r,:_, 1"_,3-_'3"_£q_ . , .._._" -I-:--_....-- - | . 'iN'...I "..: . .- ..... I...... I." ;.:: .!l; ".] • '': : • i" I:...; .." : ....
I ...... : ....:: :: :::.... ' •, i"•: :" : I "":='" .= ,.-.',. "_":" ...... =i 'J:':".. ."I "'" :'"'"":: I ...... : ' I' " SA ,deg .,oI ,:..,,,_ ,..,:. l--f i,, l,_...... _..Ji
l:j.'.. I:.:..pI." :..ZI.j ::.._.i:..1_:....:|.: .... : :1' : I n;I. Plf_.:.iji RighI yaw I0, I :....'i :1,.:i:.•,i _:; l.: _, i. r_.,i... •!, • -." .;-:i:.---,----":.: i i m0.-.0-,_".,- + 0".,_i:,;'i :: I_"i.m'" .:"":":';'"'_".T"'....":""i':"1...... •,..:...... : . ; L.J OR ' • :.,".,i O'NL + O'NR, I"_'!::i • ...... r_,. ,_ .,:...., _..i::":' ',,: deg 0 Ii_'l" i i:":"...... '.,.:. " ..., ..:..-r....-'4_"-- ' ',-_r:. .-...
-Sdeg I0 . "12 "8 "4 0 4 8 12 16 _0 deg
Figure 23.- Static lateral-directlonalcharacteristicswith stabilizatlnn *,_ engaged;. V _ 7.0-90 kn_ta, TO = -7°, _FCU = 57°, NF = 80 percent, _m -- 120°' m = 19,500 kg. '" 1971025733-107
400,000 , 1971025733-108
2°-- y ..L,. • !..,.,.....I,(: (j.: ,!k';K ti
8LP, I00 ]....._._",.,,4. ,.._...__I ...... I,,[P,',..,-,--.._.'.,,.... _ """- .- deg -I O I ...... ---=-_",...... -2o I I _-'"
20 t i ...... _ .... -,
,,_,deg 0 -:_"I" ...... • . . _,,.,,"-""'"-- ...... ,,...]..,,L_' ""_ ....."..+_..:_-. - ...... -I0 I ' I 80 _ 70
FCUL, 60 : '_ • : _ .... : p : (! deg 50 ,"_ _" ;_" " ::":" " • ' 40 30
to j I: ::...... ':
FCUR' 6050 :,,-i, '_/""_" ' [ • deg • • • ' i I : 40 ., ' j::";; i . : .:..... !.. .i i ...... • I
12,000....:::.. ,," ,i t ' .'.'" " i...... 1 • • • .. N,, ' ': ..... '"':" -I;- " ' : -'_" _ " """ ...... _ rpn, ,o,oooI I.)00 F_-,I:.:-,;. ,..:'..:_!_-£_-H-W_--!I-, ...- ..:. l -_' A , ..... i ...... = " :• • : • ." • •:.. N_, , ...... _._-:.----,: . :,
,:,'/ :; rpm II,000 . ;"I _ i .'. :.... i . " " _ .,!, ,, . io,ooo_2_::'0 2 4 _., 6 _..8 : I0 -2 ..[.0,..:i.i 2 ;" 4 ' 6:-I 8;.!l I0 Time, sec Time, sec -__ (a) Stabilizationon (b)Stabilizatoioffn
-_" ' Figure 25.- Response to lateral control input; hover rig on pedestal, ',
,_- ,;:
j,,_: 1971025733-109
20, .... ':-- • .-- :
i d ,-.= _'er(. _,.h ,,>, I0 [ ' " "_':'""_ -,II 8Lp , deg oF-->-I .---t...... ," _i ...... 1...... "_--x...... :...... 1-:...... L
-,o n . : .._.._.:._._.__:_ -20
20, - - , -- . , . : . I0 .',_ "-_ .... : " : ...... - ...,"_1_'k. -/ deg rv ",.J:i _ : .... ___..J'_._n o,-______v__...... " __A_f=..__.._......
-lo _ i ="" ] "i_rk l I v : i: -20
20 ; ----
I0 -"' .... ""
oi! ...... ;:.s:...... : "-..<...... LL -20 a
I000 | " : n • , l : , .4 : I m,-m_, rpm [ • _ _: • ,:...,..J,_.•..., • : • , .....' ,/,.% _, rad/sec2 O| ?_. i "_. %_...__r • • i i i '.:- i [_.,,. _,,,. 0 Ii ...... ". "....-'.k .... " ...... : . ,. • • . :• ...z :• , ...... , ...... -I0000 2 4 6 B I0 12 14 16 18 20 Time, sec
Figure 26.- Lateral maneuvering in hover.
-: $:
._, ...... 21: 1971025733-110
++
IO _,p,de0 t -+-Ll-.-zu'-L,I-J_ateml stick._., +4-.,..+__L_I_L-._._+L.L_._, i,.,...... ;.--; ...... • ...... " _'++".-..:...... "::,....k-._-,._ oi_.,/_..,_ ...... _ ...... i .__._ .._..i . . . . = . ' -I0 ' ...... ' ' ' ' '
IOOO 1 I i. I I 1 I'"1 I -I- f I ____-'_-1- _--1--]--_;_--_ -,
NI-N5'rpm O_0!-'-_."-."I ' • ._'_• ._W__; ,'" • " i • . . :, _, ..4_,0 rod/see2 I I ' : [ i i . : • _, _1OOOI , " • , ' ! , = , ! i i •
4, deg IOI i I i+:!i:.1' .!..I I [.TI-r'I'T]''i". • i • iI :Ti-r-f-.• 1-r-:TTT-7-r" _-_-'_ i-F:T.._--- -,oo.I____ ;, ;-i, ,__-_-____:__+.._.___...... _ ...... , ;-- ...... _,'-T'' -_----' 7 • • ;_-T-'_/,_.:.-.
]-_ i +" • • -1..... I-. i ...... 1+ i ..... --_--_'"i_'- 1 -i _' " - ' i, " " ' ' ' "" ' i .... i ---:_ r ...... I" • • i .... 1" I+
--.x.- .-='--:--b--..-+- _-_.__,._- -+++-i-__. I AY, m = ' ' ' i- l,: I . iI.....+ ! ...... , , _: : ' "•"--+-+-+I i I :.... I...... ! " __' +' • I: I. , -_5O2ooL._...... r---l---:--,ii_i,....,....' I!!i i :+1I-:, .!,:1::".4_-i,, , , , , +___,i..:...!.._.,:.-+;___ .....
' 1-i-3" ' '!'.='i ..... ' I o vy =Ay/AI, m/see °l-,_I_'_''_ _ _I,.-.,.._...... I-,=.._ ! i ! _ ,I . , , ! ,! i.r-! =,..!. i . , _ _ , . i Vy, knots -51 ...... _ "''_ ...... _-'; i !='" ..... = +il i l _"!10
--I0 Li-t-J--. if,+-+-1._.-I-_- ._ I-+_:_,J__4.._:-_L-._-.._I_.L-----L i. a,L__ " " " +_ ,L !_-1+-J- -i. I-.+.I.t-.:t.-.-1+ ___U 0 4 8 12 16 20 24 28 32 36 Time, sec
Figure 27.- Slow buildup to maximum desired lateral translational velocity in hover.
lj,
+ _., ,
, ++++_+.<•
1971025733-112
12000 =- ,.... -- ...... :...... :...... ; - "-"-
i .... ! I 10000 • _ "k7 ; • ': i " " i '..: • I_, ' . .; . ' : _ • oooo ': !\ ...... ! " :i _: ':, :. NI ,rpm Iiii ; ....I _. I • . i 6000 : , 1%;. , " i _ I . : 4000 ...... i I= N! ...... : I ...... : ; ,_ .... " I :.i i ' i I ' 2000 il " _' .I '! "_ I: •...... I:...... L...... '..._:....!' ...... J
2 I • " .; •' ' ' I "' i I" • I :___1 . i _ __.._.. ''--'-..L ! _j____.: A_,decj-- 0 1 . II'_, . -.,,_,"I-- . i' '_ " _ ' : 2 [ , , L ...... ' ,i...... • ,.[" ...... :.:....,..:.II ;_.4..I __..._,,i i
•"i ." ' ' • ,I ', ..." • I 41 • Ir a• • : • .I . ': I ': !. i. ..! .i": .. 'I
--- _}Lp'deg-4OF.._--._&_..L' .i-I. i.q-,'-', '..' I i' '! • [---'. '-=-'i..... ' lI _-J:
-20 _--.-: I:" '" ; I"" :":: "[' " ..'i :'.-"..T.[ [. ,. ' "'" i "' !. J" • .'.'" I...... :' _ • "'" i i I ...... " , i .... i . i ,.. 0 .:.:_l _ • " ' , ";'• I ...... A, , ', • 'i , ; i 'i" '" _'L deg 20 ': .k._.l.-..i...:::...-4---..,.-.i-.:-..:..-;:-:.-:-: .!..:-.. , i..:!;; ,' . ul •[....:. i :i:.:...... :1....1 :<": i : .":.,! = '-_'.-i 40 -.:...'.--: ,. I.:.::...I_ . ." .. :. • , ._.:. •.,. g...:...../,_7__l_t/._,{.i//tf. ,:t 'i""" ":'!:','"i'Actual°'.... !:.hm.tI....[ .:I 60 :' i::. ... t .:1: 'lq:i , I.... ["-._. • i
' I10 ......
90 - i.d_ _._',_l_.a, '' i-merge,-:y I,! I I ;at .4'" l I , " _" : .
FCU,deg _ . . , . 50 .... .,r,_---""_,b-.'--:--" _-; I i " " | i " '
2400 '" \ 2000 \ \ h,mleooI- "X,\\
0 .. 0
R/S,m/s 2040- - 40008000 R/S,fpm 300 I 280r-: ...... \ ' 260i • "',..,
V, knots 220i" : ,,, : 2ool. '%, ,8oi ...... ",_ 160 _ ' :".,,,.,.,....' C40[- _.... " " 1201...... -.. _ I I...... ] ....'
5 , "- -I "1"'- .L ...... '" ede_, -srE'X! ', _/\: - _/ : -,oF-15' ...... ,',./,! ,_--.L._J I0 -,- : , . T "-
- 5t_.:...._...... I...... J_Lz....L.;.. I....:.• 5 _._.L-JLongil----udino---T's-]ick"-q.-['__ -5 •-.--_-.,-r_--r-_'_ . _'.? , "_".-F, I00 P " ;...... i :, __ '. i. :, : "if'., 80 • :1'...... • ."...... ,,' • _,_ o-M,deg __ -. ._,j...• .•;_..iL.._:CL...... • J _,..--I 0 20 40 6_ 80 I00 120 Time,sec ,.._, , '_.,,:,
2igure 30.- Letdown from cruise to pro-approach using main engine nozzle i deflection;NF = 72 percent. , .'_ 1971025733-114
i ...... I t !. ;: I i i if't (,_r,(_l[l(,-,a ,r _i'rr'; ti,_" i " ' " " I • j ...... _ ;
°-FCU,degcoOJ---ff-'-,., •l :: I! "i , ...... ,. 501i .i'"----'---, : ....i. • I I...... ] ...... i" t--t-:-q
5 i"--'" I ", _" i - ":-- "'-q ...... , l • '. ' ' i orloitudtn,:.!! st,ckj I : " i " i " i'" • ' , ...... _1 BMP,deg 0 r"_-"-_...... "' I_--....,,,=...... "'" / '-',,'_, -i ---'"_'• ,:_.=:•--'_,-=_ '""-...... - j;.... "_-;"_/ , ' : I '. . [ I i "" I .... ,. , ! .• i, I• .I .._, | i ,' -5 I__".....i...... J...... I. .j ...... I...... ,...... ,...... j
I0. I...... " ...... :...... ' ...... ¢ ; • ' " ";: ;" " l I " : . r. . I • " l • i ' " " l ' ' ...... i ,. I ! i I : . l ; ] i ' •.... I "...: ., : i • .._. 8,,deg 5 •r.....• i,...... , • ....t-,..._-._ . ,,_"_,,._ ,,,,I,,..,":,,,- ,_...... -._.---,,r,...... ,'I...... :",..': ..-., .:. ,,.,_!,._.. ; . ... _.---'., . : .. . . ! . : : ..... I ! • o ,.... _4_...J_,-. _,-,j,...... ,z,w],b J.,a,____.L..._.i_-....L'_.J• ...... • • L..' . _.•.I...... : ,.J
o-M, deg I0080
25r i , .-, • ,. ' .... ' • ,'. ' ,..; ' • • • I"': .... : "" • ": i.. .:. . ! .,.1'"' i i'" -i ' j" • " ":' i ' • ": ' i' " • ' I _ ' .' • ! .: •.. " " I • ":': ":" : ; i • . : ,I . • 20_'-- ...... "-"_ ...... ' ...... t...... I ...... J ...... t' I.':'"l• '";': • .,':1I t ':J,:: .:...... !...... , '"" :l •• i.':'"l: ' .. .' 15l[.....J'.'...... ':l',i" : 'i "",:!• ,i'":.,..i':'"["::... i "'" . • ;,r:: : ..,. : .. .., ...... • ,",x,rn I.:_!...... I. :::.• _ : :!:...... :::!:::._.:'I::..:._1:...... ,...... :.: . _ :.' ,,.,I.'. ,I.,.:.1....: ...... 'i.',.,,'. "i :1 ' .,....":'" " I".:,.." ,t ,,,• ,_. .::.:...... , :. • ,I ...." ,: :...,.:...::,I: i'' :1...... ,"'1...... : ...... ;.:,,r ...... , .;:...... :,I • ...... ,I : ...... , = i, _1 ...... I", I .: .'. I,,"'., L.,', ....I i !....• , •i "1, _' I _--'IT:.."'_..'_"::'"I.-..:. :..:l.:'::.:_'."':::'-:f-'.':"l":-"/...... ,' "':'" "1 ...... '_":-l""- •, ...... ,,_', ! , , •...... t,,, I • ,: ...... iI ...... : , I : ,, , .....". ....:, , 0 ti::'.'._'=_l.!.....: '.:.',' "_.-u_..'.:i..'.._l.:..:.L_ :.I,...1::.,,'"..LL"..._...! :.'.: '! :';:__...... ,.:.|'::...!'.:...._...,.,.,_
I+H+tt_H +H-itff H-I_+t+t+H+H+t-_HF-I+_H H H-FH-I iHI+H+t] I_:PI_]]H P H-HTCJ
, 0 2 4 6 8 I0 12 14 16 18 20 ,..-I Time,sec •_i (o) By mm,, enginenozzledeflection Fisure 31.- Longitudinal translation in hover. 1971025733-115
_0[ ' I I' , • "I-- ." ' ' I ...... ! '..,._" .! '. ! _',L '_ '' ': "- : ". "", ..'.'" ! . .' ; °'FCU, (leg 50 I-.-!r, ..... ; • I .... , "-" _-._"_-'_.. i ...... _ "" ! . i • :, ' ' :......
40 I...... '...... !., i,, I ...... I .... _ . _ ..... ,...... -" . .... :.....
_MP, deg .i '...... ' !"" "...... =t'. i....:.':";'_: i it..ongiludinoll --st' lck..,.'. ,.I"..!"! ...'.. ., 0 _,"-L ...... ".... I..... "-I---"-,-I',...... ;..-.I..-...-..-.II...... -k.I ._".4._.i....,e_',-._.4.""_'-' _'_ "..._....,._...._,.,,,,_=='D.I"--" _ ,,.,4t,_._ -""=L-._ . . . , ; .: , : .;..... ,j...._.._r • ! • : : i .! " : i i I "'" : : "-'_'.,_-_--_L-_".'':L._" J-,-_.._l'-: !" " '! ;' ' .i :" ! :' + '" ' "!" : - 5 ...... _.--_-.....i...,..__.."_--,,...-7_.._... :...... I.:.....:.. :...:..=.1...... i...... ]...... -_.:=.....,
I0 5 8, deg " - - " ;' ' '...... 0
-5
{b) By attitude and nozzle
Figure 31.- Concluded. 1971025733-116
iii 0 0 0 0 0 0 0 0 I_-
_2, \
,,j 1971025733-117 1971025733-118
800 - _ Conversion'onglideslope .... Conversionin revelflight
7" _T.._.7' Lift engines s|cirw 600_ _"_, / Approach h,m __ _ thrust
Lift enginesstart I . "_ Breakout 200 Approachthrust "_ J "-_ Noisef_otprint sav;_lgJ'_ t I ...... i I I.. I . i I I I ,=._ 08000 6000 4000 2000 0 Distanceto touchdowq,m 800 -
Lift enginesstart
h, m 600 - _._ ...... [- --__... 400 /
Lift enginesstart _.. 200- ___ atTimesavinghighpower }___ "...<
O __._.L__. I I I ,. -_ 200 180 IS0 140 120 100 80 60 40 20 0 Time to touchdown,sec 160 - =,//.
_ 120 - - ...... i: V, knots _'._
$ 80 %%%%_'...... _
200 180' 160 140 120 100 80 60 40 _0 0 i O-- 1-1 I l I ._Timel to touchdown,I IN¢: I l ,r"-'_-| (a) 7" Glideslope
Figure 34.- Comparison of time and distance _.or diffe_en¢ approaches, 1971025733-119
o\ - ---.. -12 y \\ _Lift engines start
BOO=- -_-'--"'_Approach thrust
_oo- ..... _.... ______-:OConversiono°wr.in level,oflighno°t _,,,.,o_ h,m | Lift engines st(Jrt 400 - Approach thrust "_. X 200 - Noise (ootp_int _,_ 0 I l.... I savingl . I I ] ! 7000 5000 5000 100,.3 0 D.s!.:.".,'_ to touchdown, m oo-
h,m I \._
'_ 400 - Lift enginesstc,rt _',,_-\
200 - ..-p,l Time saving __.. "-."_.,_,.%. { at high power l- "_,,%._
'yl, 200 180 160 140 120 I00 80 coO 40 20 0
%_"¢ Time to touchdown, seC _-' 160
_' V, knots \
'_, 80
40 | i_ 200 if]O I60 140 120 I00 60 60 40 20 0 -'" Time to touchdown, uc -.# (b) 12= Glide slope
" FCgure 34,- Conclude4. L
_ 1971025733-121
_-_.._ )", deg
4 - - _- 8oo
R/S, m/see 3_"_"_ _- 6
_X_ R/S, fpm -
160 20 40 60 80 I00 120 140 160 V, knols
Figure 35.- Variation of rate of slnk wlth airspeed and fllght-path angle, zero wind.
|
J 1971025733-122 1971025733-123
IOO
80,000 ___.__j
H+hcN_'q-H L_H__ :"I:Il:IHq_+H4 _ IlH I:lI:IIlIIH4_ H+H+_t __ "
• 4' z 60,000 __E____._ ." ,,, __,,.H,..t_ • -- I_,__ .-m_m_H_._'. :::::::::::::::::::::_.+._.,__._.-H-_.+
I-- ._ ¢_r%O I:t_I_Nq4-H4_4-1Hd4Hd444444444+N44444_H_d4_t4444444444_I44444444444_
zo,ooo_ - ',,."- -." i _i!_ :30 40 50 60 70 80
O'T, deg
(a) One main engine ,_
Figure 37.- Thrust and fuel flow characteristics at V = 0, h = 600 M, _;_ Ta = ii ° C. "_,_ '
t % l 1971025733-124
60 . I" " I .I ;: I '. i'--: ["--r"-'."[" i • ;...... ,;.... ,...... , :I...... i /I I "• Ii. [i.. ; .;,...... I ! , i_ :., : i • ,• l • , • I[.|_, I : . f . I . , I I i ' I . , . , .l_.'il_ 5 0 |...... "...... "...... =- -..':v/_.- lii_ --- ..... , I . • , ,.... ,. ".._ .'. "_ . ".T" ; I ' . ! i . . I ' .,L_, I _ " I 401...... ,,,L.....L..--...... :, ,I-',_...... '- , r...... :...... ", :...... ,,,,,i, I ". I ,,_, ,,,", . . ., i , I I . • • • P ,. I "I, • "" ' • J-,,--i +, ...... ' ,_--' "",,'..- •;4 : "IIz.. i, I I !_.I.-, I Fl..... t, , :i[: I, , , _30 F.... • ,.:I'- . .':'TE"'_." _ti ...... t ...... ' ...... " ...... "; i _ ..... _ ' " I: "_ ...... • _'I " " " I_ = / _ _.:.....,....!. I....:...... i...:l..;'"'it .... ' ...... i" "1" i • ! i .' • i. :!:.:1 / "1' ; ; I " ; !" ;' "" "" ' "I" -6 20--.I, ...... _... ;...I ...... :. i .', ...... i. i.. ,!.....!...... i ..... ' ...... : ...... I I, /i/' I • ! •I I I. •i . • I i, " .... . ,.... ; I I:'" I '" I ' • i "; ": • :' " " " : : I hi. I.;..i i I ...... i..i.', i ...: I i!! , I:. ; ..... i ',..! _i I 0 F'J" ":! ..... : T'"'I"" 7: ...... i"...... 1.... !'. '1":"I": 7-: .... !.....I:...... i .... I:--:r-:..-1 I ..:. 1. !..!..i., .... : ! ::F ! i ..i. 1..:.1..!...i. !:.:!:...I .!:...I '", I • I .... i i' " "!.";!: -i "'1", I • I''1.." .... :.,."' "'." • '':'"l ...... • I ". J- ...... " ! ;' : i! ."_..._..J: _ ...... L...... :l I .... "'!" • i • . • "" ! .• J
22,000 [",'.7-:T-,.-" I. ;_:__H_IllI44__[[J__'_ _1_ I" i: "": ' " I'":'. 44+H+H44_HH-I-HH-H-H4H-H_H-14+HH+H'H+I-_H+I-H-H'NHffH+J-_I_id-_4_
...... - ...... _ m+H+l+m-I+t+_mm___:_ _8,ooo__..R+H4_ __: :: :: :: : _:: :"_,_;4: :H_Ht ___tH+H IHFFH_...-H H+H_::H_:I:JiCi:Lt tt.t 2[ J z _r_4:__ ___t_
'...... I:H4"H:I V" ...... uu_ u.._u._..u. ,_ _.-_ _-_- _, __+_ w.-_,-_-_+++_ X+H+H+I+IH ',: :: H : : H : ..... : ',: : H H H : H H I-H-H-H-H-HI-'_,'FI+H+H'I+H+HH-H+H-H'H+I+_H H tl-H-rl-_, H+H+_HH+h'H'H'_ F: "'"'" ...... :',:',::',',IIIH: ::;%fHI:H ...... -_
...... :.,_-______I (_ _('_(_ R4_I:=I' ,.,:II:IIIII::'.::H H::t+H4N+I+H+IH-i+H-t+I-t:h,,; ...... _:::::':HH:::H! •-'---30 40 50 60 70 80 90 crFcU,deg
(b) One lift engine
Figure 37.- Concluded.
3:
t "197"1025733-'125
3O
20 Control limit
I0 (Pitch nozzle)
8E _SPN' deg 0 (Elevator)
-I0
-20
-3°12 -8 -4 0 4 8 12 81_P, deg (a) In hover maximum angular acceleration is +0.22 rodian/second squared at NF =80%, m = 20,000 kg ...... ,.
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8Emax deg __,i!!:::!!ii_i_
_155 knots_
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(b) Gearchancjerrelations FiSU_:e38.- Lonsitudinsl control relal:ions,
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(b) Gearchanger relations
Figure 40.- Directional control relagions.