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r^^ (NASA - TM-X-62144)- LONGITUDINAL HANDLING N77-33151 QUALITIES DURING APPROACH AND LANDING OF A
POWERED LIFT STOL AIRCRAFT (NASA) 68 p 'HC
A04/MF A01 CSGL 01C Unclas G3/05 51362 LONGITUDINAL HANDLING QUALITIES DURING APPROACH AND LANDING OF A POWERED LIFT STOL_AIRCRAFT lames A. Franklin and Robert C. Innis Ames Research Center Moffett Field, Calif. 94035 bJ y,`^ w"r ^,'-T March 1972 LONGITUDINAL HANDLING QUALITIES DURING APPROACH AND LANDING OF A POWERED LIFT STOL AIRCRAFT James A. Franklin and Robert C. Innis SUMMARY Longitudinal handling qualities evaluations were conducted on the Ames Research Center Flight Simulator for Advanced Aircraft (FSAA) for the approach and landing tasks of a powered lift STOL research aircraft.
The test vehicle was a DeHavilland of Canada C -8A aircraft modified with a new wing incorporating internal blowing over an augmentor The investigation included (1) use of various flight path flap.
H and airspeed control techniques for the basic vehicle,_ (2) assessment of stability and command augmentation schemes for pitch attitude and airspeed control, (3) determination of the influence of longitudinal and vertical force coupling for the power control, (4) determination of the influence of pitch axis coupling with the thrust vector control, and (5) evaluations of the contribution of stability and command augmentation to recovery from a single engine failure. Three pilots, all having flight experience in powered' lift aircraft participated in the simulator program. Results are presented in the form of pilot ratings and commentary sub- i stantiated by landing approach time histories.
NOTATION Lift coefficient CL c Mean aerodynamic chord, ft Longitudinal column control force, lbs.
s He Elevator hinge moment, ft lbs Hs Dimensional elevator hinge moment derivative due e _t
rad/sec 2 f
to elevator deflection, ate-, rad Ie aae Dimensional elevator hinge moment derivative due H* to elevator deflection rate, 1/sec gee ,
Ie d
h ,hf Course and fine altitude, ft.
c Elevator momen± of inertia, slug-ft2 I Aircraft moment of inertia, slug-ft2 Iy Pitch attitude feedback gain to elevator or to KB nozzle, deg/deg Ke Pitch rate 'feedback gain to elevator, deg/deg/sec Kb Column feed forward gain, deg/deg Column rate feed forward gain, deg/sec/deg ' Kb K Airspeed feedback gain to nozzle, deg/ft/sec { Ku Longitudinal acceleration gain to nozzle, deg/ft/sect` M Pitching moment, ft r lbs Md Pitching moment derivative due to throttle, T I aM, ,
rad /sect/lb
a Z, I L y, iil M Pitching moment derivative due to nozzle., , rad/sec2^i^ I d^ Y Aircraft mass, slugs m< Normal acceleration, g's
n
Aircraft pitch rate, deg/sec qB q Dynamic pressure, lbs/ft Wing area, ft2 S S Laplace operator Hot thrust, lbs T it l/T Roots of the longitudinal characteristic equation l/T, sp2 sp l nominally associated with the short period mode (which in this instance are real instead of complex), rad sec Roots of the short period mode with pitch rate and 1/Tsp , l/Tsp attitude stabilization, radfsec l^T tl Roots of the short period mode with pitch rate, pitch l/Tspl,
sp
attitude and airspeed stabilization, rad/sec l/T' Low frequency roots of the longitudinal characteristic l/T'-, P P2 i equation with pitch rate, pitch attitude, and airspeed stabilization (nominally associated with the phugoid mode), rad/sec Numerator roots of the elevator to pitch attitude 1/Tel , l/Tg2 transfer function, rad/sec.
iv Low frequency numerator root of the elevator to l^Th i , l altitude transfer function, rad/sec l /T- Low frequency numerator root of the elevator to altitude transfer function in the presence of i airspeed stabilization, rad/sec V Airspeed, ft/sec, knots Aircraft gross weight, lbs W Longitudinal force, lbs X Longitudinal force derivative due to thrust, X
` ax
ft/sec2/lb
M a s 2,'
Longitudinal force derivative due to nozzle deflection, b-V
ax
I ' ft/sec^/deg
M
a
j in.
xV , x Fuselage station 16cation of the Pegasus nozzles
Z Vertical force, lbs ZST Vertical force derivative due to thrust
l d Z
ft/sec 2 jlb
m bST Vertical force derivative due to nozzle deflection, Za
1 6Z
ft^sec2^deg
M
c^-^J '
Water line location of the Pegasus nozzles, in zdl z e p( Angle of attack, deg Flight path angle, deg Incremental value Longitudinal column deflection, in Sc Elevator deflection, deg
Se
SeA Command to elevator surface actuator
^r Commanded elevator deflection, deg b ec Stability ty augmentation actuator input to the elevator dre Stability augmentation actuator command if e
sc
Stabilizer position, deg !rg Flap deflection, deb; Of
Hot thrust, lbu
Throttle position, deg
Pegasus nozzle deflection, deg '
Pegasus nozzle command, deb c Stability augmentation input to the nozzle
d^^AS
Pilot's nozzle control deflection, deg
pilot
E
.,Tlide slope, and localizes errors, deg
gs, Eloc
^ , GJA Damping ratio and nat^iral frequency of trtp elevator
surface actuator
,gip Damping ratio and natural frequency of the,phu^;oid mode
P
pp Wp Dataping ratio and natural frequency of the phugoid mode
as modified by pitch rate and attitude- stabilization
Damping ratio and natural frequency of the elevator
s' Ws
SAS actuator Damping ratio and natural frequency of tire ,numerator f Lt^e
roots of the elevator to pitch attitude transfer
presence of airspeed stabilization
function in the
Air density, slugs/ft3 deviation of atmospheric gust velocities, ft/sec Standard gust
Natural fregiioncy of the elevator-spring tab system, xad/sec
I
k ^ INTRODUCTION The pilot's control of an aircraft capable of landing at the slow flight
speeds associated with-STOL operation is complicated by problems which
are generally more severe than those of conventional aircraft landing at higher speeds. Longitudinal control of pitch attitude, flight path, and the low speed, high wing airspeed are all adversely affected by loading, f and high inertias typical of the STOL transport class of 'vehicle.
In
addition, the availability of powered lift for the pilot's control and
the associated influence on lift, drag, and pitching moment of engine
power setting makes these aircraft respond to the application of power
in a fashion considerably different (and not necessarily favorably so)
These problems are from aircraft having conventional lift concepts.
general^ recognized (not necessarily in order of importance) as: • poor longitudinal static stability sr, that attitude and speed
tend to wander during untended operation
• unstable flight path-attitude-relationship associated with
thrust required curve
operation on the "back side" of the
• changes in speed and angle of attack with power setting where
speed and angle of attack are not uniquely related as they are for aircraft using conventional lift concepts.
changing flight
while
• attitude -changes required to hold speed path with power which are opposite those of a conventional aircraft.
DESCRIPTION OF THE SIM IATION As part of the program to develop a flight research vehicle for demonstration of the augmentor wing powered lift concept and for research on STOL per- formance, handling qualities, and operating problems, a real time digital simulation of the proposed vehicle was developed for the FSAA.. The basic augmentor wing aircraft shown in Fig. l consists of a DeHavilland C-8A Buffalo airframe modified with a new wing incorporating internal blowing W
over an augmentor flap The aircraft is powered by Rolls Royce
Spey 801SF engines with offtakes from the compressor section for wing
blowing and with direct hot thrust which can be deflected through Pegasus
nozzles for thrust vector control. Pitch control is accomplished through
I the Buffalo's existing manually actuated elevator spring tab system.
Roll control and stability augmentation utilize the modified aircraft'6
blown ailerons, spoilers, ` and augmentor flap choke which are integrated
moment relation to cockpit control
to give an essentially linear rolling
deflection. Directional control and stability augmentation function
through the Buffalo's existing power actuated two segment rudder. Lateral-
directional stability augmentation provide roll damping, spiral mode
stabilization, Dutch roll damping and turn coordination to compensate for
i
the objectionable handling qualities of the basic aircraft for the STOL
flight condition, of interest.
r .: j The vehicle simulation was built on the non-linear aerodynamic characteristics as derived from static tests of a powered model of
3)
the vehicle in the Ames 40x80 ft. low-speed wind tunnel.
(2 '
A downwash model, based on finite span jet flapped wing theory and correlated with data from Ames +0x80 ft. wind tunnel, tests was used to determine the contribution of the horizontal tail.
Rotary derivatives were estimated, using jet flap theory where appropriate.
Supporting data for the downwash model and rotary derivatives are unpublished. The models themselves appear in Red's. 2 and 3.
TEST PROGRAM This vehicle simulation, with modifications to the longitudinal control system for stability and ,command augmentation, was used to evaluate the influence of certain vehicle characteristics and control system configurations on handling qualitiesduring approach and landing. Specific consideration was given to longitudinal handling qualities with emphasis on: •'studying the use of several techniques for the control of flight path and airspeed for the basic vehicle 1.. flight path control with attitude; speed control with thrust vector' 2.
flight path control with thrust vector; speed control with attitude 3. flight path control with thrust; speed control with attitude ' • assessing stability and command augmentation schemes for pitch attitude and airspeed control 1. pitch attitude command and stabilization 2. patch rate command - pitch attitude hold with varying degrees of control sensitivity 3. airspeed stabilization • determining the influence of longitudinal and vertical force coupling for the power control (variations implemented by using different trim thrust vector inclinatio;)` e,determining the influence of pitch axis coupling with the power and thrust vector control (variations implemented by using different thrust line offsets) • evaluating the contribution of the pitch rate command attitude hold and airspeed stabilization modes to recovery from a single engine failure In the approach and landing, the pilot assumed control of the aircraft,
trimmed for descent and aligned with the glide -slope and localizer of
a 1500 ft . STOL runway, The approach was initiated at 1300 feet along I a`7.5 degree glide slope at an airspeed of 60 knots. Flaps were set at ^I 65 degrees, Pegasus nozzles at 87.7 degrees and power corresponding to own 7160 pounds of hot thrust. The pilots generally introduced their; I conditions, disturbances, offsets and abuses to the task for evaluation. IFH random gust disturbances, wind shears, and crosswinds were also included as test variables. Time histories were obtained for the approach, and were supplemented by pilot commentary. Pilot ratings, based on the Cooper-Harper (4) .,were obtained for selected configurations.
v,^ f _6_ DISCUSSION OF RESULTS Longitudinal Handling Qualities Of
The Basis Augmentor Wing Aircraft
To provide a better description of the handling qualities problems
path and airspeed control of a powered lift
associated with flight the basic augmentor wing aircraft to
STOL aircraft, the response of
i-
vectoring is presented in the first few
elevator, thrust, and thrust
Fig. 2 illustrates the response to a step column input figures.
pilot. Somewhat sluggish pitch response associated with the
by the
be noted. Strong phugoid excitation
may
low short period frequency attitude excursions is observed which will
and
in the form of airspeed for precise attitude control.
require attention by the pilot Furthermore, with the low level of static stability and the non-
with this trim
linear pitching moment characteristics associated condition, nose up and nose doyen pitch disturbances produce considerably Unstable flight path response to attitude different responses.
side of the thrust required curve-
associated with operation on the back
flight -condition. Performance data of "Fig. 3
is also typical of this
the relationships between flight
more graphic description of
provide a
- 7.5deg, V = 60 kts)
airspeed The trimmed approach condition (r
path and
V curve (d t /dV = 0.2 deglkt),hence is well on the backside of the
-
path corrections at constant power through
attempts to make flight changes in attitude (and speed) wj.-.l produce a result; opposite to that of thrust to change flightpath, without any
Use
which was sought corresponding control in the pitch axis will produce an unaccustomed change in speed; that is, increasing thrust reduces descent rate and decreases airspeed, whereas speed would typically remain constant or increase with increased thrust for a conventional aircraft {at least in the absence of a large nose down trim change with thrust. Furthermore, peed and angle of attack do not bear the same relationship as for S conventionally configured aircraft, i.e., C Vtrim p L ac) since a significant portion of the lift required for steady flight is contributed by ;power, not angle of attack. Consequently, changes in engine power setting can either result in steady state flight path and m speed changes at constant angle of attack or flight path and angle of attack changes at constant speed. It may also be observed in Fig. 3 that the changes in pitch attitude required to hold speed while simultaneously changing flight path are opposite to those attitude changes normally associated with flight path corrections for conventional aircraft.
-Responses to step increases or decreases, in thrust level with no compensating longitudinal control are shown in The changes Fig.- 4.
in speed anticipated from the performance data areevident. Some variation in angle of attack caused by the thrust trim change is also present. Landing approaches where thrust was used to control flight y path are presented in Figs. 5 and 6 for high and low glideslope offsets as set up by the pilot. Attitude was maintained essentially constant up to the point of flare Speed and angle of attack excursions with thrust appear as expected.
Performance characteristics associated with flight path control with
thrust vectoring are illustrated in Fig. 7. Because the thrust vector
is oriented nearly perpendicular to the flight path for the trim con dition, changes in vector angle about this condition have effects- similar to those of thrust control for a conventional aircraft in that speed and angle of attack are directly related and a change in vector angle simply causes -a change in flight path. If there is no 'trim angle of attack change with thrust vectoring, speed remains constant.
Responses to fore and aft thrust vectoring are shown in Fig. _8. No longitudinal control was used to compensate for trim changes. The vector aft and forward responses are considerably different due to the non--linear static angle of attack stability associated with the trim condition. For forward thrust inclination a stable nose dean pitch response and an increase in airspeed may be observed due to the trim change associated with Pegasus nozzle location below the e.g.
Conversely, aft thrust inclination produces a nose; up pitch response which drives the aircraft into the region of longitudinal static In either case, the need for the pilot to control instability.
attitude is apparent if flight path corrections at a specified speed are
'
be accomplished. A time history of a landing approach for which
to
vector angle was used to control flight path at constant thrust is In this case, attitude control was used to maintain
shown in Fig. 9._
reasonably constant approach speed.
Pilot commentary relating to attitude, flight path, and airspeed
control are summarized as follows for the basic aircraft:' Pitch attitude control A • poor static stability; attitude wanders during unattended operation • sluggish response; difficulty in making rapid and precise changes in attitude Plight path control • unstable Flight path-attitude relationship (backside operation) • sluggish short-term flight path response to attitude changes • ` inability to flare precisely to low sink rate through a change in attitude • flight path and speed response to attitude changes occur with nearly the same: time constant
• changes in angle of attack with thrust which require the pilot's
attention to insure adequate angle of attack margin from stall _ to 9 flight path control with thrust vector angle changes similar to effect of thrust change on conventional aircraft • pitch coupling with thrust vectoring requires pitch control during flight path corrections r thrust vectoring not sufficient to flare; thrust difficult to
modulate precisely during flare
I " a reductions in power to steepen flight path undesirable during final stage of approach - high sink rate, low, power, longer- time lag if increased thrust is subsequently required Airspeed control a changes in speed and trim with power setting not related as for conventional aircraft • attitude changes required to hold speed while changing flight' l path are large and opposite those of conventional aircraft Preferred control techniques • flight path corrections with thrust vector, speed control with attitude • constant vector angle through flare, with thrust increase to augment flare with attitude change : x4 ._'.""'nT .tea ..v....
rt-^=-'.v
--.. ^ .. .".'. 3
' i t
-
11 -
Pilot ratings given the basic airplane for the task of a straight-in, constant speed approach under VFR conditions are tabulated below.
Pilot Rating
Control 'task A
B
Pitch attitude control 3.5 4-4.5 3.5-4 Flight path control - with thrust
3.5
- with thrust
3 3 3
vectoring Approach in turbulence 3.5 (overall rating)
3ftIs'
Crg= - 12 Fitch Attitude Stabilization and Command Augmentation To improve control of pitch attitude and to provide attitude stabilization for unattended operation, an attitude stabilization and command augmentation To mechanize system was incorporated, in the longitudinal control system.
the attitude stabilization and command features, a power actuator was used to drive the elevator through the existing mechanical controls, - including the spring tab. A block diagram of the system is shown in :Elevator-spring tab dynamics are described by the elevator Fig. 10.
hinge moment equation ._ Qr _ f = a e * O e + A e 6ec e _ -f- + H ie * O e where
H6 - .332 q
q = 1/2PV2 e _ u 2.04q + Ae - 2e 13.35 + 17.44- Ae
i+ .lob q
The elevator and SAS actuators are represented by second order transfer x =. functions where z t O
j
s
A=
i A ^s ' 6 _ C.y2 I —^S
5' ♦ 2 ^GJ S S + CVO
The system can be tailored to provide pitch rate command proportional to either column position or force with pitch attitude hold when the column is neutralized or the force relaxed. Besides permitting improvement in precision of attitude control„ this ,system reduces the pilot's workload somewhat by trimming the airplane at the desired
attitude. If the gain K of the control input integrator is set to
d zero, the system reverts to an attitude command control in proportion to column input. By suitably adjusting the control input gains, control sensitivity can be tailored to the pilot's preference.
A comparison of the pitchcontrol characteristics of the basic aircraft I with a typical: pitch rate command system is given in the following table.
TABLE 1. Comparison of Pitch Control Characteristics I - Basic Aircraft Typical Pitch Rate Command l/T s p = .`62 rad/sec 1/T'sp = .93 rad/sec l/T s p = 1.2 rad^sec ` 1^T'Sp rad/sec = 2.99 2 Cal = .22 rad/sec Wp = .27 rad sec
P
p=.15
I: P- .94
1/T^ ;= .18 rad/sec - 1/T - .37 radlsec ^ ,
8 rare -ec Q .064 rad /sect/in
=042 5 / ` /in
/
Fs /nz ,.2 lb
57 1 9 lb/9
; =" , 75
Fs/n
Ke _
deg/deg
K6 = P deg,/deg f sec K, _ .5 deg/deg C Kb = 1.5 deg/sec/deg Response to a step column input is presented in Figs. 11 and 12 for the pitch rate command-attitude hold system the pilots found to be most acceptable. This particular configuration has the characteristics shown in Table 1. It provides the pilot with _a steady pitch rate response for a column input and good attitude stabilization when the column is centered or when no force is applied. No attitude overshoot of any consequence exists and control sensitivity is favorably increased.
I Attitude stabilization against trim changes induced by thrust vectoring,- I is apparent in Fig. 13, thereby reducing, the pilot's workload when using this control for flight path corrections. In Figs. 14 and 15, the system was used in conjunction with flight path control with l thrust. In Fig. 14, ,flight path corrections were made holding attitude constant, while in Fig.15,`path corrections were made while attitude was changed to hold constant speed. In either ,case, the precise control of attitude required for the particular control technique is advantageous.
-In the process of tailoring the rate command - attitude hold system to the various pilot's preferences, various degrees of stiffness for attitude stabilization and various control sensitivities were evaluated. The system configurations encompassed ranges shown s in Table 2.
x.
i" 17 - TABLE 3. Longitudinal Control Characteristics Comparison - Effects of Speed Stabilization Pitch Rate Command Speed Stabilization Basic Aircraft i 1.10 rad/sec .62 rad/sec 1/T11 = l/T -
l
1/T sp = .94 rad/sec `1/Tsp = 1.2 rad/sec
.39 rad/sec
1/T' p _
= .22 rad/sec
(.tl p
_ l 1 /1' .81 rad/sec _ .15 = P P 76 rad/sec .06 rad/sec l/T'h = , % 1/Th; 1 .
- .68 rad/sec l/Ta = .18 rad/sec GU8 •99 ^8 =
- 9 deg/ft/sec
U K.
U-
o xe = .8 deg/deg;
deg/deg /19
=
- 2.1 deg/kt ^w = .18 deg/kt ^/V -
-.38 kt/deg
2.5 kt/deg
^h = =
V/6 J Y Il s 4 — L 18 - The speed control system functions to stabilize flight path response to changes in attitude as well as to reduce speed excursion associated with path and attitude changes. Short term flight path response to attitude changes is also more rapid when speed is stabilized. These characteristics are apparent in Fig. 17 and 18.
Flight path corrections were made with the longitudinal control I in Fig. i7, while in Fig. 18, path corrections were made with thrust.
Note that the adverse speed changes with thrust associated with the basic aircraft are no longer present However, the changes in flight path made at constant attitude still involve significant changes in angle of attack and hence the concerns-re regarding operating margin from g g g g ^ p ^^ I the stall associated with power management for the basic aircraft continue to exist.
_ I A typical landing approach time history with the attitude and speed stabilization systems operating is shown in Fig. 19. For this approach, the longitudinal control was used exclusively for flight
j path control and flare'
C Pilot ratings given for the preferred attitude and speed control configurations on the VFR approach were PR 2.0 in smooth air and 2.5 to 3.5 in 3 ft/s rms turbulence. The applicable control technique for these ratings was path control with attitude.
^ i i 19 - Longitudinal-Vertical Force- Patching Moment Coupling Effects of coupling on thrust control , A number of configurations wereevaluated which possessed various degrees of coupling between the longitudinal and vertical forces and ;pitching moment associated with thrust control.- Since the dominant influence of thrust, so far as powered lift is concerned, is in vertical force the test configurations are defined in terms of the incremental amounts of longitudinal force and pitching moment produced for an increment in vertical force.
These configurations are listed in Table 4 and are displayed in Fig. 20 along with the characteristics associated with various ranges of configurations.
Variations in the ratio of longitudinal to vertical force produced by a given change in thrust were obtained by inclining the thrust vector for the trim flight condition.
The ratio of pitching moment to vertical force was altered by changing the longitudinal thrust line offset.
Thrust vector inclination - Some effective forward inclination of the thrust vector (X^^ f'T negative) was found to be desirable , for use of thrust to control flight path.
This .favorable coupling served to reduce the speed excursions to which the pilot objected for the basic aircraft.
Fig. 21 shows the response to a step thrust
increase for thrust inclination
X /Z^ _ -.QhZ( (_, 76.4' deg) .
t - 20 - TABLE 4 Thrust Control Coupling Configurations
Confg. c za X^ df
Sr
G
ft/see 2/lb
rad/ft;,
deg
deg. lbs.
Basic .0008 -.0018 .0164 .23 65. -.024 87.7 Aircraft .00096 o. .0333 0.
go.
l 7520
.o0143
.08 100335 3 -.o8 4 -.16 .00572 .0081 5 .24 - .00558 -.o64 -.16 76.4 644o , .00333 7 -.o8 f
,00105
-.0012 9 .08 I -.002 -.0023 -.15 -.024 .23 5190 1 6o.
o - Ll - When comparedto the basic aircraft response to a similar input in 4 the reduced speed excursion is apparent. A side effect of Fig.
the forward vector inclination which adversely influenced path control - was the reduced thrust ;level required to stabilize the aircraft on the -75 deg. 60 kt flight path. Particularly for the configuration with = 60 deg. (Xd' _ .15) the increase in thrust response f /Z6'1"
tin ge lags associated with the low thrust setting made precise flight
path corrections more difficult. In addition, not enough incremental thrust was available to satisfactorily correct for,.-offsets-'above glide slope. Furthermore, at the lower power settings, th" r'.leron blowing coefficients were reduced sufficiently to seriously degrade ,.
lateral control.
In conclusion, considering the favorable and adverse characteristics of thrust inclination, its net effect on flight path control with. thrust'
may
was negligible. It be noted,, however, that if -longitudinal-
vertical force coupling were accomplished by interconnecting the thrust
and thrust vector controls,_the undesirable consequences of thrust inclination could be avoided and a more favorable tailoring of the thrust control for flight path could be achieved.
- 22 - Coupling of thrust and pitching moment can be Thrust line offset - expected to affect the speed 'response associated with flight path corrections due, to the changes in trim contributed by the path A range of pitch coupling configurations were explored for control.
two levels of thrust vector inclination (Table 4 and Fig. 20). Is tended general, forward offset of the thrust line from the c-.g.
Some to exaggerate speed excursions accompanying changes in thrust.
aft thrust line effect was found to improve speed control. Too much offset forced the pilot to control excessive attitude excursions Figs. 22 - and thereby again increased the pilot's control workload.
present responses to thrust inputs for a range of configurations to 24 Fig. 22 corresponds
tested at a,1= 90 deg. (
XS, J2a - .033) •
_ T Fig. 23 represents the extreme forward closely to the basic aircraft.
offset condition, and Fig. 24 represents a large aftoffset condition.
No longitudinal control was applied by the pilot in any of 'these cases.
Fig. 25 and 26 show landing approach time histories for the forward and aft offset configurations respectively. Thrust was used to control flight 'path in both instances while attitude was used for speed control Smaller speed excursions and less as the situation required.
longitudinal' control activity are apparent in the aft offset configuration of Fig. 26 as compared to the forward offset configuration of Fig. 25.
i 23 - r Figure 27 shows an approach for a forward offset configuration in ^ combination with forward tilt of the thrust vector. In comparison of i Fig. 21, speed and attitude excursions are again observed to be greater for the forward thrust line offset condition.
I I Effects of coupling on thrust vector control The dominant effect of thrust vectoring appears as a change in longitudinal i force, hence the test configurations in this sequence are defined in terms °,f the incremental vertical force and pitching moment produced for an f increment in longitudinal force. These configurations are listed in Table 5; and are displayed in Fig. '28 along with comments descriptive of P the characteristics of the various ranges of configuration.
u Thrust vector inclination Inclination of the thrust vector over - a range corresponding to trim Pegasus nozzle deflections from = 60 to 90 deg. had no apparent effect on control coupling so far as the pilots were concerned. For nozzle deflections of +20 deg. the increment in vertical acceleration was low enough to have an insignificant
= -20 deg
influence on flight path or speed (An d _ .028 g's for Al)
at the ^'^ = 60 deg. condition' However, the reduced. level of thrust at the more forward vector inclinations made the vector control less effective and hence not as useful to the pilot as a math controller.
} _ 24 - L Control Thrust Vector TABLE 5.
Coupling Configurations I
X
Mb
!— x-0 — Config.
C ft^sec2deg deg/ft deg d eg lbs
.0153 -.094
.o424
Basic -.024 .23 7160. 65• 87.7 Aircraft -.101 ..
0.0 0.0 .
0.0
7520.
9o.
.o16
.24 .032 .48 644o.
75.
76.4 ^1 .0157 .23 -.024
.o167 - •059
.615 .23 -.024 5190.
60.
Thrust line offset - Again, due to the changes in speed accompanying trim changes associated with pitch coupling, the vertical offset of the Pegasus nozzles from the e.g, influenced the pilot's ability to use thrust vectoring as a path control. The series of configurations only encompassed nozzle locations below the c.g. (M /X6V positive as S.V 28). The pilots found some positive pitch coupling j indicated in _Fig.
to be desirable in that the aircraft's attitude led in the direction of the intended path correction. Although some longitudinal control was required to counter the trim change in order to hold speed, the control force levels were innocuous to the pilots for the level of pitch coupling associated with the basic aircraft (F. ^ .2 lb/deg) I At the highest level of pitch coupling tested (M & 1% = .032) the !
longitudinal control necessary to trim became objectionable. In 29, an example of response to a step change in thrust vector Fig.
for an uncoupled control configuration is shown. Essentially no change in speed with the change in flight path is apparent. By contrast, an approach for the configuration having the greatest pitch coupling is shown in Fig. 30. In this case, some longitudinal control was required to maintain the desired approach speed. Sustained' column forces did not exceed 5 lbs during the approach prior to flare.
I r
CF
t
a
i Recovery from Engine Failure Considerable evaluation of the behavior of the basic aircraft following a single engine failure and development of suitable control techniques been accomplished during previous simulator investigations for recovery has at Ames. Results of these tests are anticipated to be published shortly.
It was of interest during the current phase of testing to determine the effects, favorable and unfavorable, of the selected attitude and speed j stabilization systems on engine out recovery.
II f _ L In summary, the basic aircraft's initial response to the loss of one r engine consists of: • an immediate increase in sink rate • an increase in airspeed and very little yaw prior to configuration change • some roll I i increases thrust on the If the landing is to be continued, the pilot remaining engine and vectors the nozzle aft to re-establish the is maintained at or slightly above 60 knots.
glide slope. Speed Lateral_ and directional controls are used to counter rolling and yawing moments due to nozzle deflection and due to increased thrust on the remaining engine. Acceptable landings can be made if i sufficient altitude is available to arrest the increased sink rate.
I A typical single out landing is shown in Fig. 31. Engine out waveoffs are performed by increasing thrust, vectoring the nozzle full aft, } i I.
i and raising the flaps to 30 deg. Speed is allowed to increase to 75-80 kts for best climb performance. Typical altitude losses during recovery are 100 -150 ft. in excess of those experienced during a normal two engine waveoff. A time history of an engine out landing with the attitude and speed stabilization systems- engaged is shown in Fig. 32. Whatever improvement exists over the basic aircraft lies in part with the precise attitude control, and the ability of the pilot to have the airplane in more precise control on the approach prior to the engine failure. Speed stabilization has the unfavorable characteristic upon failure of an engine of rotating the nozzles forward (vectoring the remaining hot thrust aft) to counter the increase in airspeed which follows the loss of powered lift.' Sink rate increases even more as result of the nozzle response until the pilot can counter with
increased thrust. If the landing is to be continued, the speed hold
at 60 kts can ultimately aid the pilot as soon as sufficient thrust is applied to regain the glide slope. If a wave off is to be made,, the pilot must have either the capability to override the nozzle comand from the speed control system or the ability +.o quickly and
precisely- select the desired climb speed of -80 kts. Given this
capability, the speed hold system can assist the pilot in establishing his climb condition and thereby relieve s(m.e of his work load.
28 - CONCLUSIONS Considering the difficulty in obtaining satisfactory control of pitch attitude, flight; path, and airspeed typical of powered lift'STOL aircraft, j this simulator investigation has provided an indication of improvements which can be made in the aircraft's attitude, thrust, and thrust vector controls to make the aircraft more acceptable to the pilot for the STOL approach and landing. The results relate to; • stability and command augmentation for attitude and airspeed control i r • longitudinal and vertical force and pitching moment coupling associated with the thrust and thrust vector controls • impact of stability and command augmentation on recovery from a single engine failure i i I Specific conclusions regarding each of these categories are: Attitude Stabilization and Command Augmentation • improves precision and speed of response for attitude changes improves control sensitivity • stabilizes against trim changes.and external disturbances • pitch rate command preferred over attitude commandto relieve pilot's
` trimming workload
Airspeed Stabilization
• stabilizes flight path response to attitude changes
• provides more rapid flight path response to attitude changes • reduces speed excursions associated with path and attitude changes i y ! • reduces speed changes with thrust f -29 w Thrust Control Coupling
• forward thrust vector inclination (negative;X&. /-Z &r) preferred
j in order, to reduce speed changes when controlling flight path with thrust
• adverse effects of forward vector inclination associated with
lower trim thrust can be avoided by interconnecting thrust and i thrust vectoring controls to achieve desired coupling thrust line aft of c.g. preferred in order to provide pitch coupling (positive M ` ) to reduce speed excursions when
i -
sT^
controlling path with thrust
i Thrust Vector Control Coupling • thrust vector inclination with respect to the vertical of 30 degrees has only minor influence on flight path control with thrust vectoring i (reduced thrust reduces effectiveness of vectoring for :flight corrections) preferred 0 some nozzle offset below the c.6. (positive
%IX&d
for vector control of path Stability and Command Augmentation On Engine Out Recovery f
• attitude' and speed stabilization effective in permitting precise
control of attitude and speed prior to engine failure and helpful in recovery to a satisfactory flight condition Following the initial transients engine failure • response of speed control to 'transients following
adversely affect flight path control by ;increasing rate" of descent
y , pilot must be 'provided with the capability to override nozzle
commands from the speed stabilization system and the ability to
select a new commanded airspeed quickly and precisely
-30-
REFERENCES Quigley, H. C., Sinclair, Nark, and O'Keefe, A Progress Report 1.
on the Development of an Augmentor Wing Jet STOL Research Aircraft, SAE Paper 710757, National Aeronautics and Space Engineering and Sept. 2 8-30 , Manufacturing Meeting, Los Angeles, <Calif. 1971 Rumsey, P. C. and Spitzer, R. E., Simulator Model Specification 2.
for the Augmentor Wing Jet STOL Research Aircraft, NASA CR- 114434, December 1971 I Cleveland, William B., Augmentor Wing Jet STOL Research Aircraft 3.
Digital Simulation, NASA TM X-62,149, March 1972.
Cooper, G. E. and Harper, R. P., The Use of Pilot Rating in the 4.
Evaluation of Aircraft Handling Qualities, NASA TN D- 5153, April 1969 I ,x ,f 4' r` ^j i I ill r.'
r t^
Weights
Maximum Gross 45,000 lbs.
Maximum Landing 43,000 lbs.
Operational Empty 32,600 lbs.
Inertias (Maximum Gross Weight) 380,000 slug ft' 1X 207,160 slug ft'
I
5 52,6 10 slug ft' Iz Center of Gravity Limits (Horizontal tail incidence of 0 0 , 40,000 lbs Forward 24Y0% MAC 33.0% Rear MAC Areas square feet 865 square feet Wing area, total including ailerons flaps and 111 of fuselage 187.10 square feet Wing flap area, projected, including ailerons aft of wing line 46.30 square feet Total aileron area aft of hinge line, including trim tab- 233 square feet Horizontal tail area, total 81.5 square feet Elevator aft ofhinge line Vertical tail area, total 152 square feel` Rudder aft of hinge line: 30 square feet ore Trailing 30 square feet Dimensions and General Data Wings: 78.75 feet Span Root Chord 12.58 feet 7.74 feet Tip Chord 12.1 feet Mean aerodynamic chord Aerofoil section NACA 643A417.5 (MOD) Root 632A615 (MOD) Tip NACA Sweep back at 40 percent chord zero degrees Dihedral, outer wing only 5.0 degrees (Note, Leading edge sweep back and dihedral each start 17.6 feet from plane of symmetry,) Aspect ratio 7.2 Fig. ld.
Aircraft Characteristics
Ailerons: 1 1,50 feet Span 2.01 feet Chord aft of hinge line 33.70 feet Distance from plane of symmetry to centroid of aileron 20.0 percent Aerodynamic balance Spoilers; 11.30 feet Span 1.18 feet Chord 62,4 percent Position of hinge line percent wing chord (average) Flaps; Span 55.70 feet 3.2 feet Chord aft of hinge line Horizontal tail: feet 32.0 Spam 8,33 feet Root chord 6.25 .feet Mean Aerodynamic Chord Aerofoil Section:° (MOD) NACA 63A2`14 Root (inverted) Tip NACA 63-212 (MOD) (inverted) 4.8 degrees Sweep of leading edge zero degrees Dihedral j 4.4 Aspect ratio I Vertical tail 13.60 feet Span 14.00 feet Root chord f_ , 833 feet Tip chord 11.41 feet Mean aerodynamic chord NACA 63(215)014 (MOD) Airfoil section 22.6 degrees Sweep of leading edge 1.2 Aspect ratio 28,7 feet Overall height 93.32 feet Overall length (with probe of 16 feet) 46.3 feet i Distance, wing MAC, 1/4C, to horizontal tail MAC, 1/4C 43.4 feet Distance, wing MAC, 1/4C, to vertical tail MAC, 1/4C +2.5 degrees Wing incidence, angle +1,0 degrees Horizontal tail incidence angle (adjustable) Control Surface Deflections and Rates Flaps 6 5° down to 75° down 4°/sec extension and retraction ;.
Pegasus nozzles 18.5° to 116.0° (down from aft of aircraft) ' 90° /sec l Ailerons ±17° about +30° max droop angle 30°/sec j Spoilers -500 100° /sec Augmentor Choke 65% choke gap area closure at 75° f ► ap deflection' i 30°/sec Rudder ±25° forward segment I, I +25° trailing segment I -50° /sec Elevator +250 -15° _ ft.
zoo- loco WIT
Ft
I^R.
A/@A**%
4"AAA
9r 7, *, O Trim u_ W i in /7p ra, >.A .0 G1.
v • K t's `rn r 4 r4%pow 4%L& oc /^RlIGIG ---
—^^
...v •e V 25 ^ : ubo Ids /0 d.?/set
-u
- -.u.
1 a ftC ro E^.•Taa -/o
-a6
^taec Nor TIO¢aolt.
flow.T.^.rsr s, o. ooa Tw,r it's ^iao fts A101"
ruP
,4•^rrce ss
^'^
so dad .__
a
&- e Gr_1 oe
co"wo
P POSIH Els dog APT eowm N FWD COLUA40 * Figure 2. Longitudinal Response to a Step Column Input - Basic Augmentor Wing Aircraft i ^r ]
l^
R
a
•^ 4 / s
g
V
.Y
F^
i- j
CM- ^- rr b ru.dd l rtq
Figure 3. Augmentor Wing Aircraft Landing Approach Performance - Thrust Control C^,oe j Acn nae
L—
I del + ► + -4 r' ^^sr DCJC.QSO.s^ T^wsT ,Iuccc^tr 1 ,MO
Q A.a
Logo
.JG /AIM r ^y ,dlOi c h 9s - 1000 K 1Zo AN^.e
A,tsp^-o
Cw QO Tr I!, Con i - t. on - 20 40,000 !bs GM f0 Rdμ 660 kts VC, RTT 11 t o -7.5 deg j d631ML Clio, oti : 61 d. y - 10 . 7 - 1+0 = 7160 Ib5' bT 65 ,ley J^ : EAVArae
&ev 4rb e
-20 -/o
der WV{/NN r a _ T1f^sT_ F h^s i fl4V (et &1) 46w v^8 , 1" T ;wr aT ^ er I ^A! ^ ltU _ 4 0 9'4p d^ , o io mo Ass 1:tee r -ro !
Longitudinal Response Figure 4. to a Step Thrust Change - Basic Augmentor Wing Aircraft 14L77 rwF C06W^A41 46ITIO^- -e QA7L 164D J o / m m
qv
kh
_^ w
cw ^o aq Vo (A' kLr ^i1Glf P Tr l+ _ -T.'^ drC - _. r^ r* 4° a . Iro 0 ^^ JbY l^[/11L 0147 Am C _ s Ar hbU Iba Ac a0 (05 •i.y b ^/lY/iIUQ FLEV^9fae S co r'I SAS { ^n16[G .- SAS_ —s- - ^^ dog
Cv
10 sec. i Sn+eiu a No Ac° T , uH.,.- d(QT.
^..,.eU.
^9^ . oeo-..,.per - -,o
Nom Lup
4,16Le Qs(rnov do?
010Y ^+ 4 6Uoer LwAU *E*- dal ` Ih —•^ LT —4 Figure 5. Landing Approach Time History for Flight Path Control with Thrust - Basic Augmentor Wing Aircraft A i COLIAA,U nrwo
A
low iIVAD
^^ _ _
-- o-rte ;^. . ;.
low
4sv64c 2 0
A7mc< -o- v all 4 #.)
-24 Trim Cowlit lev Sv7-Y7jdj GW - 40,OC<) to aq vo - 60 kts Ar- M hr ltbr'lb5 P1 T b 5 to • -7-5 & 1 ima m ddy /sea Cky `O C41 io Mt ASILie C9 Sr 4 ^ Dr-
7-
' 4 7
IT.W er 41ZLA r,-4p Al 6 1 . 5
+
LacAu eak 6U0W We- U.
day Ah L:r -4 Landing Approach Time History Figure 6.
for Flight Path Control with Thrust - Basic Augnentor Wing Aircraft
i
f i
o
i
-^ ^^ 1 s i
b
e4 er ^
O J
Q
Figure 7. Augmentor Wing Landing Approach Performance - Thrust Vector Control ------^ de! wl -4I Gp^....^QD ^^Ciltt.t O NO'Ct^• ArT 7 A.DO rC A S LOAD
Fr g t °F
ct g
Im r b 2.O ANela A 1ts^0 - —,.
OL -- .°^ ,.
r ArrACK -o,^ t v del Cun. { t „_, - 2 o Trio, 40,nx1U 11, GW to p 60 kLa V° v Rr^..l rlrLM t o D - -R^TL- W. der.
^o m = bSv'y b+ F^vArne F.^evarce ^ 2/±)- °°9 -ro "C .--.
Poop Wvr CbcurnN
T
_ ^...
t r
lba mj -4 0 so
TAM
— ^Rbi.L^ 1 na.J 40 - - —
dQq
dy
/0 -rte, 1...
loo ^,alu=ce Figure 8. Longitudinal Response to a Step Thrust Vector Change - Basic Augmentor Wing Aircraft _.........—....^...__ 7000 .^ ^I Cot wAv ^ ^4Lnn^
dos i naJ -
v _e O^ low — -- - ?
Q•v7E 144D nj /io
Aj6ce m
4 es^,a oL
-- -
_o.
Am yx
v
-sO Trin Con,l1':•tj aw 40,0 .) v 6,, = e7. ► e1 ^TC* Q V0 60 kLr.
A RM P rp . .%.i dry• = (05 dog e ~ t ^ + Ffmn b+ Cloy
e
-^o c^a9 -so
i a xc
to
to ll------
—^TMSOLf r^pT RNA' tx TihA94Ar - b N
d y «qrT-W RT
A° Nbu.ci r', n p b [B ^^^ M y ^ ^'QT, TIOA!
dal 1369 J RT .^ b + UXAut--0- 6U cc
ES
CRA011i day Landing Approach Time History Figure 9.
for Flight Path Control with Thrust Vector - Basic Augmentor Wing Aircraft ^ f ^^ 0 ion q ^ ^iwc I ( Ekvaioi- ^t ^e^ Eleval-o^ ^c oC ^^n Tab naw,^cs 14Ct ua ^or I*Y e
l^pna^ acs
SAS
5A 5
Actua+vr
11(21 KS
7i'
Figure 10. Block Diagram of Pitch Attitude Stabilization and Command System
i i ^ E s
r too 1060 Axi nI;9e A- 6 rorvDo too -3 RA rf or A/04vtoitc.
Agi m
2C Oq "q $POD 10 OFA MVA Trim Con4ition GW 40,000 Ibs.
& T z VW.) wn v 60 kts Am."
q*rg- O s 7.5 deg 6 4.
to - -T , dsy dap /sec • Stabilization Attltad 5-0 K K. -A . - D de, deg dam 0 .5 KO Kb C-Ur.64rVe l.
;-Ug/see —J aw '^j -- 0 ' .
d4 9 to M& --,04r r^4&11, r Aar 7'00w &,& or) Odrpr) 5, M C 10410 Ib3 '7;4r lbs - . 4 -- - a --- .- low , 4 1 w C. LI O :Z,, ',j as Apr 6LJ Oc dowm-^
d. 9
A . M Figure 11. Longitudinal Response to a Step Column Input - Pitch Rate Command/Attitude Hold System COL&WS) Accnrwa ^ ► ^^ I , ► -^.
-Pass - -e O —_ 2 .0 Qq^ {t^ww oao KI Ali . !^i fll
NN^
,11 M1 O !jp
A
MbC^ ^
A,ta^ zp
--- -
WR '"- Asrwx -o Ii ►
k - I n
to" L i,
I
—^ Trim Condition -all d'r /p CM - 40,000 lbs. by7 bo lb,:, br l o Vo - 6o tta
p
i 1GF^ ATC* d t ^^ = bS APB) ro ` -7.5 deg
Il'^ IIIL ; r
jf, ' II ^ I Attitude Stabilization der .24K K ^ b deg doe I.S did tw, CL* V.4rw- KQ `?A
- - Q" ,g me d
to eg/sec Kd _t4erbr c ews dog HP— to ZrA81LI eeie 4r dO°D r W &my) ... .... mrrmmmmn,mmmm KaT QT ^bn
'747
-^o a eo Ivbzr 'ZAP -- -- _da^ o ^+ bur LMAUec*- Lot dAf A-+ d7 Lt-4 Figure 12. Attitude Changes by Pilot Using Pitch Rate Command/Attitude Hold System
f^
„^^--- v + Iboo NI • ♦ ^ LOW P .ri k/^,M 9s - Moo iZo 1 n ANS ^c OL ^9 v 4b - 20 Trim Condition i0 W - 40,000 lbe . i M.l s 6 I .
Yo 60 kts Sr = T,ec "I" -7.$ deg rlt, ro dw dejImL B p -IG Attitude Stabllisation Ke ' O 2. K a . a0 d deg d Qz,.%roe Fd^VATLQ.
de do ryc , Ko -a•p deg Kb ,^Ib(!l /sec W S ^ -ro io sac Asod Cbt t l^.tN
¢ War
- rCStTlq^/ T K (Rt• E^^^ 4OW -4 p /b qo r ,r w
4 v
d d °9
^ b
io ACOfs roD -10 Figure 13.
Longitudinal Response to a Step Thrust Vector Change - Pitch Rate Command/Attitude Hold System
W 1 1h
ALnrUW I
Coww y
H um
Il^lj^
-e o Pa^6 ! &D - n Trin r,nittlin 0V 40,000 lbs.
A i e =O /Q Yo 60 kta d eE ^Ii ► rt ^m n -^., eon v'tilt(ffRrtrtrlRRt11 g wAcjc IIII^^ (^I ► I^IIIIlilllYo k* dl 1 dry 49 0^ _ !l 1 2WO MW111111T1TRR1 MdmW --4 Att1t-Or Stabilization
dy
ddy/jW, Ka -90
a^ e a_ ^ ._" de Su K j ..
& deg rile G v*me &EV14 I.S Kd ^Y, A &u — 0 dog aec , L e
1ro
: aw smelt.; Nor dam Roue __ 2 Tor &:r —JO O _ _ N MATS r4.4 p aO - dog o' er4 ♦ WCAU etc s L10 6UO6 i dy day Figure 14. Flight Path Control with Thrust r Pitch Rate Command/Attitude Hold System Op- l ^ ^^ ^I^I!!!!^^i^ji . I I iill^;;ll'I^l,Ijl^ll Ilf s ui^^^^^^llli!ph I ^ I I,!IIIII I
I ^
.^', 1'''I A,_nrwc "11^' 1 '^ I' l ' ' llp COLAAv i, ^ i ^ 1 ^„^^ 11 °III — - ^^+1ll1Wuu.,,t, ^^lu^^^L 1, r^s•
la+o ^^ _ o. n3
"'' ► !m"^ —" I II1 iit^^^ s ,i I. ^ !^) . illlff^ifR I I fl i ; i .
{t^wrw _ ,ono (I h r ioll I^i III O 4u 30 Of- 1 A^4s^eza^
eF
f Arme g ,..^^ .... .-ni111V^^ I -
a kir ^ Is m,Ihnllm nMitmry
I ERrr
A l
4D Trim Qnid;LIon av 40,0a; 11-.;. 81 1 .?p 10 &1 Vo 60 ktr = 7160 Ibs ^p17^M I ^'''" 'M ,^ ^lllf' - i1i11fflflll(I^f(ill^ ^^^
e
— /D ar Attitude S'.,10 t::a K Q 2.0 d h b ^O L e, deg G sec 3# /•S dc ^liY/9/LiQ ^tEVgIn2^ d O ' 1 '^ a ;. K K ^y^s O ^ ,
L4
1°^"` ♦ /o A000 1.fpr- .Sm^81L 1^ el^Lj T __ f4 liiiiiiiiiiiiiiiijuliiulilLULLb Tllim &4 lbo Thor pr _ O
/^08tii n[ ,p ► 1
IQM^LI •. ^ _ rpS^Tlcn! _ do? ^ ^7 4d)4D
to 4
6uog (GCA te*_ swot --^ E^eot .....
-^'^^^ - iuuiu` ta^, .u;u^tE` day dif l7 -4 Figure 15. Flight Path Control with Thrust, Airspeed Control with Attitude - Pitch Kate Command/Attitude Hold System l,l
^or ► 9 ^iud ^na l
tt
as^s5 W'
^•tin^^tS
AIuuafor
A
_A IK,4 Figure 16. Block Diagram of Airspeed Stabilization System 1! T"r AtrirvDf t.; O F t i0ft ^O A06LS Asa spczz I . ^, : Trim C,-%,1 i Wit -del GW 110,030 11,S• Pf-ldel MOM.
V (0 kto ZS Ir-O r Attitude 6tr•blLfzntion t * dj Isec K K , , LeG.
dot', I der, de - F./sce K.
K6 -.2 . 0 2S dt F, K 410 -00 , oil W.71"kMhff MMM, 77-A ft -/ 'r. c KO p/Qp deC 2 0-0 K u f t7 i, - jc Aor rm&ar dor- ro,,wjr (oe,6,# r ) OCOPH S, lbs /*S rt- 4 P ^20t'AJ A- deg day
_s
Apr 4 d code Longitudinal Response to a Figure 17.
Step Column Input - Pitch Rate Command and Airspeed Stabilization Systems J.— ................
rw.• — .-^^1' YYYi' M low h Ft E^to^.
E9' 041 wf -4 • ► TIIfYfT N6itf^.ti T Mt^J ZuCfA1^ 2 1400 PAM- -- -
room BOA$ -o-
^mr+ _ Fe AMD 9s O Trim Conditint: tn<. = 81.-1 d'9
ou - 40,000 b„
fZD &. _ -1140 vo - 6o kte Ibs
14"" 20
Ale-3PEED kts - 40 v v Attitude Stabi I I ,:,t un d^ K w d Ka ^ Zp deg 6 deg fo /b I S deg sec Ko Kb ° -2.0 d^/c,., ARrI ^TCN __ ^/ B Speed Stabilizat ,t.
e do do -fp ^f0 KO RU 0.0 , yd f^t:t• d.
K^ • /O 0.0 ftai•r Fs?^rve &evpne ^^ s I -rc b nee roxe moo Abe
CN
___ S Sc .-^. Tit&s1S_ g.
Af 1 rw (et. EAt) 4000 TNOr ar lba fh! 'ewer ar
-4
/f0 , naJ ^ v o'
d °9
.o fG
S
S171BiujeiC 1O0 Acorn
A.AdhMjM
p IrYN^7 ^__ ^l ^f0 .10 ,, , iunrtmme^M`t11MM ^,,, Figure 18. Longitudinal Response to a Step Thrust Change - Pitch Rate Command and Airspeed Stabilization Systems
^ _/Ow
A COLWAIJ AcnrwF '^ '1 - e 0* LdD ^ , /AM h -low '^. ^ rl O Trlm Ci.
ou 40, 000 1, : 87. 7 del ANbc a _ 2O vo 60 *.qtr, llw ^ 1b5 Jr o -O 'i71t .
K kis 1U p _ ak .-- Attitude Sta'. • 11 - ----r— dc' ^^ KA :2.0 deb b de; IO de C acc KA 1 0 Kb 1'S d/ar Dr 7e F + ATE* ti de! sPC L $Labi'i%ati on dy.w ^ ` _ do do K 9.0 U O
r
M /O 10 etc.
D^ aSTJ?BILILt TH07 60 QT '47 -,o 4ZL&4E
r r^ ''P
gs,rloN
80 w
da^ de9 ,0
`° ♦ Ems_
LCALI fCk ( ^ 06 `-ys day dAj ^- ♦ LT-4 Figure 19. Landing Approach Time History - Pitch Rate Command and Airspeed Stabilization Systems a. -1
s x^i
X&,
_, 101
-aSr
Murndexs i-,d tc +&
-+es+ coN-bi ura+1oKs
© 0f Table 4
Q
-. 05
.01 ® . 00 5
M ST' root
ft
zsT
O O O
.OS
-15
X6T 17 6T
MODE RATE TO
SMaLl.
-.10
SPEED
LA IM E 5PM C H AW WeF
CNAM""S
W
frw Tft IzU3T
MDDERAW T'L7
WITM
-IUQL)tr
LA" E 1El^N ;
-. 05
CNAN WtTW
7146441T
.01
.005
rad
jK'S'r Figure 20. Axial-Vertical Force-Pitching a k Moment Test Configurations for Thrust Control t.
6u or
44.77 rune 't Ea^eoe- 1 i A14 PATE LOO °F -0- ^Fi^C70^ - 9 s - x.100 Trim Cunditiuu I ZD 0w 40, Ol0 1 t .. b y = TV 4 dr1 1,0 AN4xe I L! Vo 60 kt i ^^ a pg40 tf» S^a —_
of C _ 15 Af, —^
_T.S de b ^ - p ro asfAtaC kt-3 V - 20 Thrust Control Co ,3 /o /V x6„ M d 7, rar .out - - R*w Rr<M xe c ^p b EL&VAMe E.ceVATLe I^VtLI.Q ^^ S —_a &SA1 ddl - -ro 10 sec M Saco Cbw.^r ^' &- 7, - ,hW4L _ sir ► gV (at. EAt) 4000 -^ rnj /bc ~•, R, -4 13D .
L t,, .
- — — - -- --
^ n01LI ^ oil d°9 b ro !^ ^71Bt ujt^ p 1O0 C , wr per » CuM I - 4p 41, -ro ao Longitudinal Response to Figure 21.
Thrust Change - Configuration r 6.0 or As 1 nos y^ w f t — i
-4
Alpo of " 9 ^ - Ague Trim C.,nditLoe 0W _ 40,000 lbs.
IZD "c) Alcsp"D v = 60 kts 16;Zu Ib" ar _ - -7. 5 des b5 da.j a+ ro V all^ O Thrust Control Co pli" to p x6T M b T rad a 7 ^.ro ^1, 1-- rt A.t,^ RrcK 6 T S T F a / xe IO ^ ^LiVA1trE • EtEVATD2 4y&" bG S: sew - -ro io sa.
H 4 ^l00 OXJAiW be t _ i fl k (et. Enj) 4oeo -row wr - ^ns .4 Iba 13D to AketS r-AP ^no1J 40
d °9
40 dy A q by h ► B/uge[ ?tars i60 Aldo" o aaff Ic C i e, M del -ro ^o Figure 22. Longitudinal Response to Thrust Change - Configuration 1 .
Yom
I a: 41-n rLve of -4 41 1 A70o eATQ L4no of -0- n &JOAO 9s - Avoo K Trim Condition 4 = 40 de., GW 40,W, Lbs.
IZD Iby v o 5to AN 60 ktc 61 6 dr s 1 or ! ^'^'!.- r -T.5 deg dF s bS ci..` Ir A ouc del Mrust Control Colplfng - ZO X6 /O A "L -,OOpft o A Ank il1LM Z6TT.
_P AW -_ 41mve c IO ^ Gu-^me &--vArte —
Av^
—e std s asks doq 10 Sw War ^c +N (R1• CI) TMor e r IOU /ba -4 /70 so
t
Itl^ti FcAp del ro o /0 !1^3 loo B/LfjC^ Mi t r jo t I Longitudinal Response to Figure 23.
Change - Configuration 2 Thrust 4.&r not —
lip
de7
I I-Ov -0-
RV.,m _L
9 S hADM Trim Condit1w.
IZD OW - 40,030 lbs. &I) 21D ANA 6r 1520 At e VO - 60 kto _3p-EED 9 7 C 1 w l r,; -0- ft Iro - -7.5 dee k,t3 4 0 A ID to 16T S T iad ..053 ^,-6 T,— T fAW `T e 06 J/ ML Cie, -to 3 cote b FA.ev.+ne, :5A S, - 01 47 % (0 EAt) Inj -4 13D FZ- , qp OD dj day
Ass 'so
6kWA TM MON 0 Ole, del ..........
Figure 2 4. Longitudinal Response to Thrust Change Configuration 4 Cowwv Acrrrws ^ ^^ y r o - 2-0 tow AV all Trim Ce_r.tition 4J,OOn 1r d = 4v /xp cw ""9 tsw 'es Qijue 60 kts br = A V pj _ a^ drt OF --_ -- °` _ _/.y ecg r o l^lr'/71CK ^ ^^ ^, Tt.rvst Ccrti^^ -10 M x ,Ar ad .T ".3 s Ptak 6T ATE* 4 dayl&A CWY
-io ao `
^o
x
=o Cis-mrue s^ do g °l°9 tt] ► 8/L1^ ^T ^^ ^STV `, -^o -• -- /Its - .
4aL46
; TAP 80
b^ AV61'e QS,r?w!
(129 er 4 LO ♦ Bee oe- nrAL.I 6-j Eq,
s<p
all dAf LT
-4
Figure 25. Landing Approach Time History - Configuration 2
ti
COC."a) 4t77nAw
2-0 wr
&lkD O Trim Conditi, n Gw % 60 kt• 6r '75 W ib" OF 64. 65 del lr o -7.5 due -0 04"Mic It* 1 lvVI 41, "brunt Cont Ml Om!k).Ln M X6T S rad ..033 SCSI ru- V4 79.0 ^r r RTC* ddy /sea "I &'Lsv, q rtke Qov*mp - ses* s =o
— Se
-—r-
O (Mq — 11 C 4 , — —
to SIX to &MO O LIteg 44or- c747 4&Etz ;:ZAP
.44V & J,& W qS, ntoj
da?
d29 LI D UX AL I 6 COL 6W&W
311"
dAf LT -4 Figure 26.
Landing Approach Time History Configuration 4
e
-Cou,wy Acne G — ?• ,.,^ ors - -e- _ `^fy^ G!u.^B i ^ • ,P1 YY 9p T_!a -- Cundi 4l — ^^ un sL> A>va^ ^' Ow 40,0Oo 11,.
dy, z 76.4 d`, Vo • 60 k6n 1y
ll S 6440
bT = kis RAG ° ^ -7.5 deg - 15 a y ey _ L 0o"_tm, ^.: 1t to Ttr'' -- ^nrRlf x6T ^TCI^ ► +a _ 61.
a(& / Ada
d o
^e
-lo
oe
O at yu ' ^..e _ d•f 1 o see — - . a^'718fLf^ +—+ ^T ^ / r,^ a, IJobtcb C:4p daq 40 dy
b+ -
^4
Ow L WAL+ ^rQDR .
vtq μ-+ Lr-4 Figure 27.
Landing Approach Time History - Configuration
n^
.5
z_bv
XSV
.03" .04 . U .UL
b^
L ITRE
AT woc
FXWSSI ve
M
TA( E
tr^rx At 6AT rp- t M G*A*)
OR. X
w1r4 Tk s7
k TW C4M.4E ^v
•51
PMTUOE '
V^iro Q ^ N ^ Cc^ rc Ttt^¢a^r C'NMU6E r iU6 e_ 1/^ TD
W 1 7W
Toiwsr
^ p .^
UL _. •U3
• ^ Figure 28. Axial-Vertical Force - Pitching Moment Test Configurations for Thrust Vector Control
I
I ^
t
_^
Z- 1 4 Guar ^0i.nnfafc ^ ,__ ^' '^, iFll 3U_wG 0 Eeeoe.
del rf -4
1 A000
RAYN LW
-o-
r F,^ctoft CL O-48 CLO-48 ,Ft /-.
r1 9s Trim C^n^Altfnn 01: dy s W dol 40,0JO 1 t :.
ANbt,.E V o 60 kts
A,zsP^o
a MAO tbs Sr - deg (e5 dr.)
r'o 7.5 b f - a1r*cd .0- rr y kt3 ntl .ast Vector Control Cvj1,11- del - M e v .
br . o.o o „^
10 to x6 ) fti
R"j. Aft,,, z _Arm d7^set -10
b
^D &evorce S.
_. ---tea - -- e9 to ^ M doo0 War e c TWA&gr- -- As ( nqN (Rt. L C At) 4400 MOT ar -4 12D li 9^ — r-.qp cl y de l io loo ^7 ► Bturc^ _, Mors tCc J_— K i -!o fr .^G Longitudinal Response to a Step Figure 29.
Thrust Vector Change - Configuration 11 o A 4rnAx Coeww ^ 1 k --- ^°' I"` rA* ,add (061 flaw - .._....
,000 R*m to+o
M , --
M - a .- Gam ---0.
low Trf n Cun-W ion O cw - 40,030 It... ay z 40,t" /^
i{^J"0 10 bo Icts Sr a 75'2A lbs
Vo .
^ f = 65 deg
^ I'o -7.5 a o E
/OsrllGil' - o CO%trul C oupling,
k V Thrust Vecto r
dol M to f r e O.0 X bJ Odd Zed .
OfRrr P _ d•^, l day -20 - IO °49 -1 -20 sO Stt Ip °°°°
y ee
:.S^+eI^I^ Nor -^o e % 446 LO qs, r?oo/
° a 9
4 °9 b4 Or4 1ocAu tsvit 6u oar E t,$3 Figure 30. Landing Approach Time History
Configuration 13
4L ri rwov Co 4,4" a 6C k 9 q T E
lwo
ri
Aju x
'Os dal Trim Cot.' 1 1 . 1 Ow 40,0,)o )L..
to v., 6o kLs ST lbs Di 1G1^1 ATW C" -7.5 dv -10 lftsvqme A"A,Z- d" tow t o FAI L sm smlem ' 414ALie 'D —JL "R C 1 4 7 O lab I rIAP 4ZLLZ 11199 UX4LI tC* dif LT -4 Time History of an Engine Figure 31.
Failure During Landing Approach Basic Augmentor Wing Aircraft
V—
, 4LnnAw --&-
C-
2.0 T". 'i 77.
Ij y C.
ISO ' Gx) 11, = 91 1 del .5t, Av6LC
1100 1 b5
A g es farzv toF OL -Y.5 dv,^ b5 del v ell W Attit -2 . 0 K,, 2 0 dcC, KO 20 Yi '.6; RM4 dd y /sec do t' do K u (I " C) j d,, do AO K. 0-0 ' - - U ^L7r.vl.
ELsvArn,e Se -ab .
#0 -W- 1 L
I &;-t 8l L 60-
S., ..
4abb L , T -ftr P-r o r 14 P ^nr6lE 4v ^^ ^1 r1o^/ de9 40 4 0 11
+
II x, 06 ib,- 4 LI 45e*
de y IJ Af $A LT -4 Figure 32.
Time History of an Engine Failure During Landing Approach- Attitude and Speed Stabilization Systems Kngaged