Document
NASA Technical Memorandum 108866
Simulation Model of the Integrated
Flight/PropulsionControlSystem,
Displays, and Propulsion System for
an ASTOVLLift-Fan Aircraft
William W. Y. Chung, Paul F. Borchers, and James A. Franklin Ames Research Center, Moffett Field, California April 1995 National Aeronautics and Space Administration Ames Research Center Moffett Field, California 94035-1000 NOMENCLATURE APP approach control mode ATM automatic transition control sub-mode CTO cruise/takeoff control mode FTM full thrust command control sub-mode HUD head-up display IAS indicated airspeed IGV inlet guide vane LFI lift-fan inlet MTV manual thrust vector control mode RCS reaction control system SCAS stability and command augmentation system STOVL short takeoff and vertical landing TRC translational rate command control mode V/STOL vertical or short takeoff and landing WOW weight on wheels commanded angle of attack, radians (Xc AM aerodynamic pitching moment that propulsion system counteracts, ft-lb AX body x-axis force that SCAS commands, lb AZ body z-axis force that SCAS commands, lb quickened flightpath angle for HUD, deg 8A aileron deflection, deg canard deflection, deg 5C iii PRECEEHNG PAGE BLANK NOT FILMED 8CN cruise-nozzle deflection angle, deg longitudinal deflection angle of the lift-fan nozzle, deg 8LF x lateral deflection angle of the lift-fan nozzle, deg 5LFy 8LN lift-nozzle deflection angle prior to yaw axis command additions, deg 6LLN left lift-nozzle deflection angle, deg longitudinal stick deflection, inches 8long _neu cruise-nozzle deflection angle that produces no pitching moment
8PRCS reaction control system pitch thrust, lb
5R rudder deflection, deg right lift-nozzle deflection angle, deg reaction control system roll thrust, lb 8p,_cs ST, 81h throttle position, deg 8yRCS reaction control system yaw thrust, lb
¢ bank angle, radians
cruise-nozzle efficiency factor rlCN lift-fan nozzle efficiency factor rlLF lift-nozzle efficiency factor rlLN net thrust vector angle, deg Os,Oc commanded pitch attitude, radians core (lift/cruise) engine bandwidth, rad/sec _core mean aerodynamic chord, ft pitching moment coefficient due to ground effect
Cmm
.pitching moment coefficient at zero angle of attack
%0
iv pitching moment coefficient due to angle of attack, 1/rad Clnt_ D aerodynamic drag, lb auxiliary inlet force in body axis x-direction, lb FAI x auxiliary inlet force in body axis z-direction, Ib FAI z cruise-nozzle force in body axis x-direction, lb FCN x cruise-nozzle force in body axis z-direction, Ib FCN z FEX total propulsion system force in body axis x-direction, lb FEy total propulsion system force in body axis y-direction,, lb FEZ total propulsion system force in body axis z-direction, Ib lift-fan force in body axis, x-direction, lb FLF x left lift-nozzle force in body axis, x-direction, lb FLLNx lift-nozzle force in body axis x-direction, lb FLNx lift-nozzle force in body axis z-direction, lb FLN z primary inlet force in body axis x-direction, lb FpI x primary inlet force in body axis z-direction, lb FpI z right lift-nozzle force in body axis, x-direction, lb FRLN x acceleration due to gravity, ft/sec 2 g
fi
vertical velocity, ft/sec K control system gain KAUG augmentation ratio of the lift fan 1CN distance from cruise nozzle to center of gravity, ft L aerodynamic lift, lb cruise-nozzle moment arm, ft LCN V lift-fan moment arm, ft
LLF
lift-nozzle moment arm, ft LLN factor to modify jet-induced pitching moment, scaled to lift-fan thrust MGELF factor to modify jet-induced pitching moment, scaled to lift-nozzle thrust MGELN m aircraft mass, slugs mAI mass flow rate, auxiliary inlet, lbm/sec me engine mass flow rate, slugs/sec mLF mass flow rate, lift fan, lbrn/sec mPI mass flow rate, primary inlet, lbm/sec PB aircraft roll rate in body axis, rad/sec QB aircraft pitch rate in body axis, rad/sec dynamic pressure, lb/ft 2 RB aircraft yaw rate in body axis, rad/sec S reference wing area, ft 2 lower command scaling slope for the throttle, authority/percent of throttle travel Slope L upper command scaling slope for the throttle, authority/percent of throttle travel Slope u cruise-nozzle thrust, lb TCN cruise-nozzle thrust command, lb TCNcM D cruise-nozzle thrust prior to modification for axial commands, lb TCN1 total propulsion system rolling moment, ft-lb TEL total propulsion system pitching moment, ft-lb TEM total propulsion system yawing moment, ft-lb TEN factor to modify lift-fan thrust in ground effect TGELF vi TGELN factor to modify lift-nozzle thrust in ground effect total lift-fan thrust, lb TLF lift-fan thrust due to inlet guide vane authority, lb TLF F lift-fan thrust due to fan inertia response, lb TLF s lift-fan thrust prior to modification for axial commands, lb TLF1 lift-fan thrust command, lb TLF CMD left lift-nozzle thrust, lb TLLN left lift-nozzle thrust command, lb TLLNcM D lift-nozzle thrust commanded prior to axial command additions TLN1 fight lift-nozzle thrust, lb TRI.N right lift-nozzle thrust command, lb TRLN CMD total core (lift/cruise) engine thrust commanded, lb TTcM D total core (lift/cruise) engine thrust, lb TTOT referenced lift/cruise engine steady state thrust, lb TTRe f aircraft velocity in body axis, x-direction, ft/sec UB Vaf filtered airspeed, ft/sec aircraft velocity in body axis, y-direction, ft/sec VB x-axis groundspeed, ft/sec Vx VT true airspeed, ft/sec W aircraft weight, lb aircraft velocity in body axis, z-direction, ft/sec WB longitudinal distance from cruise nozzle to aircraft center of gravity, ft xCN longitudinal distance from lift-fan nozzle to aircraft center of gravity, ft XLF vii -i longitudinal distance from lift nozzle to aircraft center of gravity, ft XLN lateral distance from lift nozzle to aircraft center of gravity, ft YLN vertical distance from cruise nozzle to aircraft center of gravity, ft zCN vertical distance from lift-fan nozzle to aircraft center of gravity, ft
ZLv
vertical distance from lift nozzle to aircraft center of gravity, ft ZLN °°° Vlll Simulation Model of the Integrated Flight/Propulsion Control System, Displays, and Propulsion System for an ASTOVL Lift-Fan Aircraft William W. Y. Chung, Paul F. Borchers, and James A. Franklin Ames Research Center SUMMARY A simulation model has been developed for use in piloted evaluations of takeoff, transition, hover, and landing characteristics of an advanced, short takeoff, vertical landing lift-fan fighter air- craft. The flight/propulsion control system includes modes for several response types which are coupled to the aircraft's aerodynamic and propulsion system effectors through a control selector tai- lored to the lift-fan propulsion system. Head-up display modes for approach and hover, tailored to their corresponding control modes are provided in the simulation. Propulsion system components modeled include a remote lift fan and a lift/cruise engine. Their static performance and dynamic response are represented by the model. A separate report describes the subsonic, power-off aerody- namics and jet induced aerodynamics in hover and forward flight, including ground effects.
INTRODUC_ON NASA Ames Research Center is participating in technology development for advanced short takeoff vertical landing (ASTOVL) fighter aircraft as a member of the Advanced Research Projects Agency (ARPA) ASTOVL program. Integration of flight and propulsion controls is one of the criti- cal technologies being pursued in that program. NASA's role in the program is to participate in developing design guidelines for integrated flight/propulsion controls, support ARPA technology development for ASTOVL demonstrator aircraft, and provide consultation on integrated control design to ARPA contractors. This work will be accomplished in a joint program with ARPA, Department of Defense agencies, US and UK industry, and the UK Ministry of Defence. Specifi- cally, NASA will carry out design guideline analyses for the control system and conduct piloted simulations on the Ames Research Center Vertical Motion Simulator (VMS) to evaluate design guidelines and to assess the merits of contending design approaches.
The initial effort in this program has been to develop a mathematical model for simulation of a representative ASTOVL aircraft concept. This simulation development has followed the approach employed in previous STOVL aircraft simulation models, an example of which is provided in refer- ence 1. Representation of the aircraft's aerodynamic characteristics are presented in reference 2. This report describes the aircraft's integrated flight/propulsion control system and head-up display and the propulsion system performance and dynamic response. This simulation model has been used in an experiment on the VMS to gain initial experience with control system behavior and flying qualities for this aircraft concept.
Description of the Lift-Fan STOVL Aircraft The lift-fan STOVL aircraft is a single-place, single-engine fighter/attack aircraft, shown in figure 1, featuring a wing-canard arrangement with twin vertical tails. Geometric characteristics of the configuration are summarized in table 1; mass properties are specified in table 2.
The propulsion system concept is presented in figure 2. It consists of a remote lift fan coupled to a lift/cruise turbofan engine to permit continuous transfer of energy from the lift/cruise engine to the lift fan. Further, the lift/cruise engine exhaust is either ducted aft to a thrust deflecting cruise nozzle in conventional flight or diverted to two deflecting lift nozzles in vertical flight. Throughout transi- tion, flow can be continuously transferred between the cruise and lift nozzles. Lift fan and lift-nozzle thrust can be deflected downwards from 45 to 100 deg relative to the aircraft waterline. The cruise nozzle can be deflected +_20 deg from the vertical.
The basic flight-control system consists of the canard, ailerons and twin rudders for aerodynamic effectors during forward flight. For powered-lift operation, control is provided by differential thrust transfer between the lift fan and lift nozzles, deflection of lift-fan and lift-nozzle thrust, and deflec- tion of cruise-nozzle thrust. Pitch control is achieved by a combination of canard deflection, thrust transfer between the lift fan and lift nozzles, and deflection of the cruise nozzle. Roll control is pro- duced by the ailerons and differential thrust transfer between the lift nozzles. Yaw control is derived from the combination of rudder deflection, differential lift-nozzle deflection, and lateral lift-fan thrust deflection. As an option, reaction control, powered by engine compressor bleed air, can pro- vide additional control moments through nozzles located in the wing extremities and in the tail.
Longitudinal acceleration is achieved through thrust transfer between the lift fan, lift nozzles and cruise nozzles and by deflection of the lift-fan and lift-nozzle thrust. The integrated flight/propulsion control system must coordinate all of these aerodynamic and propulsive control effectors so that, from the pilot's perspective, the aircraft consistently responds to cockpit inceptor inputs, whether in airbome or jetbome flight.
Integrated Flight/Propulsion Control System The flight/propulsion control system model follows the structure shown in figure 3. It includes the pilot's command inputs, the regulator that acts on the pilot's commands and sensed aircraft and propulsion system state variables, and the control selector that couples these commands to the appropriate aerodynamic or propulsion control effectors. The pilot's commands and regulator are subdivided corresponding to pitch, roll, and yaw attitude controls, and vertical, longitudinal, and lat- eral translational velocity controls. Within each of these, the different command modes shown in table 3 may be obtained depending on the phase of flight and the pilot's task. Descriptions of these modes follow, presented in order of increasing system complexity.
In the cruise/takeoff (CTO) mode, the pilot has direct control of the magnitude of the lift/cruise engine thrust. The propulsive lift system is not in use, and the pilot has no direct control of the thrust vector angle.Ratedampingaugmentation is providedfor pitch androll control alongwith dutchroll dampingandturn coordinationfor theyaw axis.
Themanualthrustvector(MTV) modeallowsthepilot to controlthe magnitudeof the propul-
sionsystemthrust(lift-fan plus lift/cruise enginethrust)aswell asthe directionof the net thrust vector.No feedbackcontrolloopsareusedfor eitherspeed or flightpathcontrol. Pitch androll are
controlledthroughratecommand/attitude hold augmentation in transition,blendingto attitude
command/attitude hold at low speed.Yaw controlis the sameasfor CTO athigherairspeeds during transition,andthenblendsto yaw ratecommand at low speed.
The approach(APP)modeactivatesa longitudinalacceleration command/velocity hold system,
with the net thrustvectorangleasthe speed controleffector.In decelerating from wingborneto
poweredlift flight, a flightpath controlsystemis activatedasthe netthrustvectorangleexceeds
70 degandthe commanded coreenginethrustexceeds 70 percentof its maximumvalue.This flight-
pathsystemremainsactivateduntil thenet thrustvectorangledecreases below 50 deg.The portion
of the APPmodethatdoesnot usea flightpath controlleris calledfull thrustmode(FTM), as the
pilot still hasdirectcontrol of thecruiseenginethrust,while thatportionthatusesa flightpath con- troller is calledautomatictransitionmode(ATM). Pitch,roll, andyaw control areidenticalto that for MTV.
Lastly, the translationalratecommand (TRC) modeis reserved exclusivelyfor the low-speed
poweredlift andhoverflight regime.Propulsionsystemcontrolin this modeconsistsof a vertical
velocity anda longitudinal velocity controlsystem. The controlsystemusesroll attitudeto provide the pilot with direct commandof lateralvelocity.The yaw axiscontrolremainsthesameasthat for MTV.
Control modeavailability is subjectto therestrictionsshownin figure 4. Assumingthe lift fan andits inlet/exhaustmechanical components havea 250knot airspeed operational limit, CTO is the only mode available for this and higher airspeeds. The pilot engages the lift fan by selecting MTV or APP when the airspeed is between 250 and 150 knots; if the lift fan is not engaged by the time the airspeed drops below 150 knots, then a landing in CTO mode is required, or the pilot must wave off and start the approach again. TRC cannot be selected until the airspeed has dropped below 60 knots, nor can TRC be used to command an airspeed greater than this value. Upon a landing following jet- borne flight, the flight control mode automatically reverts to MTV, disengaging the vertical velocity control system. It follows that MTV is the only fan-engaged mode used for vertical and short takeoff as it provides the pilot with direct control of the thrust vector. The CTO mode can also be selected while the aircraft is on the ground, allowing for conventional takeoff. The lift fan will be shut down automatically in flight if the airspeed exceeds 250 knots; the pilot can manually disengage the lift fan by selecting CTO mode, but this option is only available if the net thrust vector angle is directed fully aft. This feature prevents the pilot from disengaging the lift fan while the fan is still providing significant propulsive lift for the aircraft. Lastly, the MTV mode can be selected any time the lift fan is engaged by depressing a button on the nozzle lever. This button disengages the clutch that back- drives the thrust vector handle and allows the pilot to manually direct the thrust vector, providing a rapid acceleration or waveoff capability.
J_
The materialwhich follows first describes thepilot commandinputsandregulatorlawsfor each
of the six control axes,andthenpresents the controlselector.
Command and Regulator Models Pitch control- The pitch attitude control modes for transition and hover are a direct application of the state-rate feedback implicit-model following system that has been evaluated on a number of V/STOL aircraft simulations at Ames Research Center, including the AV-8 Harrier and the E-7A and Mixed Flow Vectored Thrust STOVL aircraft. Its structure is shown in figure 5. The pilot's inputs come through the control stick and trim switch. Either attitude-command/attitude hold or rate- command/attitude hold modes can be implemented, depending on whether the gain K 111 is zero or not. Rate-command/attitude hold is blended to attitude-command/attitude hold below 70 knots by ramping K 111 between its nominal value and zero over the speed range from 70 to 60 knots. Gain K222 establishes control sensitivity for attitude command; K 111 serves the same purpose for the rate command mode while K222 is used to adjust overshoot in pitch response. Proportional-plus-integral control in the forward path is obtained from the positive feedback of the pitch command through the first order lag defined by x9. The value of the lag time constant was set to approximate the lag of the pitch control actuator. Gain KFB blends feedback variables between body angular rate and Euler axis sensors depending on flap position (KFB = 0 flaps up; 1 flaps down).
The combination of angular acceleration feedback and proportional plus integral control pro- duces a very robust system to variations in aircraft configuration and flight condition. Angular accel- eration provides sufficient lead compensation to restore stability margins that are eroded by the integral control, while still maintaining the closed-loop gain to achieve the specified control band- width as well as good command tracking and disturbance rejection. As noted in table 4 only the final forward loop gain (K300) is changed in proportion to variations in sensitivity of the control effectors over the transition and hover envelope. Otherwise, the gains are virtually the same as those used for the Harrier. Integrator limiting is imposed when the pitch control effectors reach the limit of their effectiveness or are rate saturated. Attitude stabilization is synchronized and integral control is removed when the aircraft is on the ground (wow on).
Roll control- Roll control modes are presented in figure 6. Rate-command/attitude hold or attitude-command/attitude hold modes are produced with the same form of control laws as for pitch.
The control stick and trim switch provide the pilot's inputs, and the control gains and limiting logic are completely analogous to those for pitch. As was the case for pitch control, rate-command/attitude -hold is blended to attitude-command/attitude-hold, in this case by ramping K333 from its nominal value to zero over the speed range from 50 to 40 knots. Gain K 1 establishes control sensitivity for attitude command; K333 serves this purpose for the rate command mode while K1 is used to adjust overshoot in pitch response. Lateral velocity control is accomplished through appropriate commands to the roll control. Switch C in figure 6 is activated by mode control to select either roll command or lateral velocity command through the control stick for hover operations. Gain K6 sets the control sensitivity, K_2, K9, and the equalization transfer function provide the desired control bandwidth and associated roll response to the pilot's velocity command inputs. The forward loop gain K3 is adjusted in proportion to variation in control sensitivity. Proportional-plus-integral control were achieved as for the pitch control with the positive feedback of the lagged roll control command, where x8 approximates the roll control actuator lag. Again, KFB blends body angular rate and Euler axis feedbacks depending on flap position.Attitude stabilizationandintegralcontrolareremovedon the ground.
Yaw control- Yaw control modes (fig. 7) provide sideslip command during transition and blend to yaw rate command at low speed and in hover. The gain KB provides a linear blend between the two modes over the speed range from 50 to 40 knots. The yaw control command is produced by the difference between the commanded yaw acceleration and the yaw acceleration produced by the basic aircraft's yaw response characteristics. In forward flight the commanded acceleration is produced by a second order dutch roll model derived from measured sideslip and sideslip rate, where the sideslip feedback gain is chosen to set dutch roll frequency and the sideslip rate gain establishes dutch roll damping. During low speed and hover, the same state-rate feedback structure as employed for pitch and roll control is used for yaw rate command, including angular acceleration feedback and propor- tional plus integral feed forward. Feedback of sideslip, roll rate, yaw rate, and lateral control posi- tion, in conjunction with the model of the known nonlinear directional aerodynamic characteristics, define the contribution of the basic aircraft to yaw acceleration.
Flightpath and velocity control- The flightpath and velocity control modes use a combination of the state-rate feedback concept along with the nonlinear-inverse system design which has been investigated in flight on the NASA Quiet Short Haul Research Aircraft (QSRA) (ref. 3) and several STOVL flight simulations which include the E-7A (ref. 4) and Mixed Flow Vectored Thrust STOVL aircraft (ref. 5). The same design method also has been investigated in the E-7D flight simulation (ref. 6) which included a detailed component level propulsion system. The state-rate feedback por- tion of the flightpath and velocity control is used to generate acceleration commands to the nonlinear inverse system when an advanced augmentation mode, i.e., APP or TRC, is engaged.
The flightpath angle or vertical velocity command (fig. 8) is controlled by the throttle. This command input, in combination with the vertical velocity and normal acceleration feedbacks, pro- duces the vertical acceleration command, _.cmd- The implicit proportional-plus-integral feed forward is employed in the forward loop to achieve good model following response and noise rejection. None of the gains vary throughout transition or in hover. The velocity Vx is ground speed along track and is used to convert the pilot's flightpath angle command to an equivalent vertical velocity command.
For groundspeeds below 60 knots, this velocity is frozen at the 60-knot value to convert the pilot's command from flightpath angle to vertical velocity as is appropriate for hover and low speed flight.
Switch G is closed when either APP or TRC is engaged. ATM in the acceleration feedback path is a second order transfer function with a damping ratio of 0.707 and a natural frequency of 5 rad/sec to provide a smooth transition when either flightpath command loop or vertical speed command loop is engaged. Switch F is closed in TRC to allow an altitude hold response for precision hover. When it is engaged, the reference altitude hprim e is synchronized with the measured radar altitude until the pilot's vertical velocity command is centered, at which time the reference altitude is held at the past value.
The longitudinal velocity control shown in figure 9 produces axial acceleration commands, Xcm d, with a control structure similar to that for vertical velocity. During transition between forward flight and hover, the pilot commands the longitudinal acceleration using a thumbwheel on the control stick.
This mode is selected by engaging APP mode, Switch E. Longitudinal acceleration and groundspeed combine with the pilot's commands and proportional-plus-integral feed forward to complete the ,) implicit staterate feedbackmodelfollowing design.All the gainsin this controlloop remaincon-
stant.For precisionhover,thethumbwheelis switchedout in TRC mode,SwitchF, andthecontrol
stickprovideslongitudinal velocity commands. In this mode,the stickis disconnected from thepitch
attitudecommandsystem.While mostof the hovermaneuvering is performedata constantpitch
attitude,attitudechangescanbe madethroughuseof thetrim switchshownin
figure 5.
In CTO mode,the throttle directly controlsthecoreengine'sRPM andtotal thrustmagnitude, figure 8. In MTV, a nozzlelever controller is used to control the thrust vector angle. The total thrust command, which is set by the throttle, and the thrust vector angle command, which is set by the noz- zle lever, are sent to the thrust management system to calculate each nozzle thrust command and deflection.
When either APP or TRC is engaged, the commands for thrust and deflection of each nozzle are derived from longitudinal acceleration command, Xcmd, vertical acceleration command, Z'cmd, (both in flightpath inertia axes), and pitch acceleration command, dlcmd. To convert these acceleration commands to each nozzle's thrust and deflection commands, the axial and vertical components of the required total thrust from the propulsion system are calculated. These forces, AX and AZ, are cal- culated based on the aerodynamic lift and drag forces, aircraft attitudes, inlet forces, weight of the aircraft, and translational acceleration commands (eqs. 1 and 2).
zSX=mXcmdCOS 0s-m_.cmdsin0s +Dcos_c-Lsino_cCOS _+Wsin0s-FpIx-FAIX (1) m_cm d sin 0s + mZcmd cos 0s + D sin O_c- W cos 0s
(2)
AZ = cos _) + L cos (Xc -FpIz -FAIZ APP has two sub-modes, the full thrust mode (FI'M) and the automatic thrust mode (ATM). In APP/FTM, throttle commands total thrust and thumb-wheel commands flightpath acceleration com- mand, Xcm d. In APP/ATM, the throttle controller is converted to flightpath command, ]tcmd, at speeds higher than 60 knots, and to vertical velocity command, lacmd, at speeds lower than 60 knots.
The switch from APP/FTM to APP/ATM occurs when the thrust vector angle is greater than 70 deg and the throttle setting is greater than 60 percent. Reversion APP/ATM to APP/FTM occurs when the thrust vector angle is less than 50 deg and is independent to the throttle setting.
Since in APP/FTM, pilot controls Xcm d with the thumbwheel controller, and altitude with pitch attitude, the vertical acceleration command, _ cmd, anc_ the output of the first order filter, with the time constant of XCNT in figure 8, in the vertical acceleration command loop, are all set to zero. 2 cmd and the output of the first order filter are initialized to the flightpath angle command, _/cmd, at the time when transitioning from APP/FTM to APP/ATM as shown in equation 3.
AX - m Xcmd COS 0s - D cos _ + L sin Otc cos (_- W sin 0s + FpIX + FAIX (3) 7' cmd = - m sin 0 s Os, in equation 1, is the pitch attitude command from the attitude command loop in flightpath command mode, and is controlled by the longitudinal stick, 81ong, and the pitch trim switch. In flightpath commandmode,pitch response is decoupled from the flightpath command andlongitudi- nal speed command.To providethe capabilityof controlling longitudinalspeed with pitch attitudein
APPmode,the flightpath commandandspeed controlloopshaveto be disabled.This is doneby dis-
connecting (Is from pitch attitude control, 8long, when the longitudinal speed control, i.e., thumb- wheel, is in detent and ground speed is less than 60 knots. This effectively converts Rcm d and Y'cmd from the inertia coordinate system to an axes system following the body axis with an angle defined by 0s. As a consequence, the pitch and speed coupling is introduced. The input from the pitch trim rate switch to 0s remains active to give pilot control over this new axes system with respect to the body axes as well as the pitch attitude in this control configuration. The flightpath command and speed control mode is restored whenever pilot roils the thumb-wheel out of detent.
The throttle controller commands total thrust magnitude in CTO, MTV, and APP/FTM. In APP/ATM, the throttle commands _/cmd when ground speed is greater than 60 knots, and l_cmd when speed is less than 60 knots. In TRC, the throttle controls ficmd" At control mode switching, the new control sensitivity is bounded by an upper scaling slope and a lower scaling slope, figure 10, to pre- vent any unreasonable control sensitivity change due to the mode switch. This bounded envelope reflects a reasonable match between the physical throttle position and the thrust command, 7cmd, and l_cm d. To minimize any transient response during the throttle control's reconfiguration, the throttle control command is scaled to match the new control command state at the mode switch. Thus, if mode switching occurs within the bounded slopes at a steady flight condition, no transient response is introduced. The scaling logic is performed as follows: At the mode switch, if the throttle position is greater than (Sth)ref, an upper throttle command scaling slope, Slopeu, is calculated based on the current throttle position, 8th, and the maximum throttle travel, (Sth)max, the maximum command authority, (_cmd)max, and the current state of the new throttle control parameter, 8, which is the thrust magnitude for the thrust command mode, flightpath angle for 7cmd mode, or the vertical velocity for the hcm d mode.
(_cmd) max - 8
(4)
Slopeu = (_ilh)max - _th where 8cmd is: Tcm d in CTO, MTV, and APP/FTM ]tcm d in APP/ATM and Vx -> 60 knots flcm d in Vx < 60 knots This slope is bounded by the shaded area as shown in figure 10. The lower slope for 8th < (Sth)ref, SlopeL, is defined by equation 5 and 6.
(5)
(_cind)ref = (Scmd)max - Slopeu x[ (_h)max - (_lh) ref ] (Scrod) ref - (Scrod) min
(6)
SlOpeL = (81h)ref The _cmd is then calculated as follows: _cmd = (_cmd)re f ÷ Sl°peu × [(_th)ma x - (Sth)ref] _ith -> (_th)ref
(7)
_cmd = (Scmd)min + Sl°peL × 8th _th < (_th)re f The throttle controller is scaled from 0 to 100 percent travel, i.e., (_th)max = 100.0, and the maximum and minimum _cmd, and the slopes are listed in table 5. The maximum and minimum slopes are selected to generate the reasonable mode switching transient response and logical throttle position for this experiment.
Thrust Management System The flight control mode selected by the pilot determines the logic that allocates the core engine thrust to the nozzles and determines the engine nozzle deflection angles. The method used for the CTO mode, MTV mode, and the FTM portion of the APP mode will be presented first, followed by the method for the ATM portion of the APP mode and the TRC mode. The solutions for the propul- sion effectors will satisfy either net thrust and thrust vector angle commands or vertical velocity and longitudinal acceleration commands. The control selector superimposes additional commands on those thrusts and deflections that the propulsion control system determines; these commands satisfy the pitch, roll and yaw SCAS, and will be discussed in detail in a later section of the paper. The response of the propulsion system to the final thrust commands will be discussed after the control selector.
In the CTO, MTV, and FTM portion of the APP mode, the pilot has direct control of the core engine thrust. In the CTO and MTV modes, the pilot directly controls the net thrust vector angle, while a longitudinal acceleration command/velocity hold loop commands this effector in the APP mode. With these control features in mind, the logic for determining the allocation of thrust to the cruise nozzle, lift nozzles and lift fan can be broken down into three basic cases once the lift fan is operative. The first propulsion system solutions that follow apply to those instances when the lift fan is inoperative, or when the commanded net thrust vector angle is less than or equal to zero deg, where the positive net thrust deflection angle is measured downward from the waterline of the air- craft. Figure 11 shows the convention for the nozzle deflection angles and the measurement of noz- zle distances from the center of gravity. The net lift-fan thrust (TLF 1 ) and the net lift-nozzle thrust (TLN1) are set equal to zero, and the longitudinal deflection angles of the lift fan and the lift nozzles are set to their maximum deflection values, that is, the deflection angle that directs the thrust of these nozzles as far aft as possible. The core engine thrust commanded by the pilot is TTCMD. The basic thrust equation for the aircraft, regardless of control mode, is: TLF1 TCN1 TLN1 KAUG TILF + _ + TILN - TTCMD (8) In this equation, rlLF, TILN, and rlcN are the efficiency factors of the lift fan, the lift nozzles and the cruise nozzle, respectively, and KAU G is the lift-fan augmentation ratio. For the case in which the pilot hasmanualcontrol of the core engine thrust, the net cruise-nozzle thrust TCN 1 is simply TTCMD minus the core engine thrust that provides minimal lift-fan and lift-nozzle thrust, multiplied by rlCN. The cruise-nozzle deflection angle, _SCr,i,is determined following the engine nozzle moment arm calculations. The lift-fan pitch moment arm and the lift-nozzle moment arm are calculated from the respective nozzle deflections, _SLFx and _LN, and the distances of these nozzles from the center of gravity of the aircraft: (9) LLF = XLF cos 8LF x + ZLF sin 8LF x (10) LLN = - XLN cos 5LN - ZLN sin _LN For this aircraft, equations 9 and 10 will always produce positive values for LLF and LLN. The cruise-nozzle deflection angle for which cruise-nozzle thrust does not produce a pitching moment on the aircraft, 8neu, is calculated: 8neu = -tan-l(zC------N-N/ (11) \XCN J The distance from the cruise nozzle to the aircraft center of gravity (1CN)is _(zCN)2 + (xCN)2.
Longitudinal cruise-nozzle deflection is used to counteract the aerodynamic pitching moment, AM, as shown in equation 12: sin-l[, -AM - (TLF1)LLF + (TLN1)LLN 8CN + 5ne u (12)
L (TcNI)ICN
where: AM = (Cma (z + Cm(3E KGE + Cmo ) _S_- Cma is the variation in pitching moment due to angle of attack CraG E is the variation in pitching moment due to ground effect KGE varies ground effect as a function of altitude Cmo is the pitching moment coefficient at zero angle of attack The next case is applicable when the lift fan is engaged and the commanded net thrust vector angle is less than 45 deg. Table 6 shows the status of the propulsion system parameters for each group of solutions that follow for the MTV mode and the APP/FTM submode. As before, the lift-fan and lift-nozzle deflection angle are assumed to be at their maximum (aft) deflection, and the lift- nozzle thrust is set to zero. The propulsion system pitching moment equation can be written in terms of the nozzle thrusts and the vertical thrust component of the cruise nozzle, FCNz: 9 5_ (TLF 1) LLF - (FCNz) xCN = - AM + (TLN1) LLN (13) Using TCN 1 as an approximation for the longitudinal component of the cruise-nozzle thrust, the net thrust deflection angle ON can be expressed in terms of the ratio of vertical thrust to longitudinal thrust: tan_l( TLF 1 COS 8LF x - FCN z ] ON (14) _ si--n_ + TCN------ _
A
Equation 14 can be rewritten as: (15) TLF 1 [cos 5LF x - (tan ON) sin 8LFx] - TCI,,I1(tan ON) - FCN z = 0 As TLN 1 is set to zero and TTCMD, ON, and 5LF x are known quantities, equations 13, 14, and 16 can be solved for TLF 1, TCN 1 and FCN z. FCN z and TCN 1 determine the cruise-nozzle deflection angle:
-I(-FCNz
(16) _SCN=sin _,_) Equations 8, 13, 15, and 16 are used to determine TLF1, TCN1, and FCN until the cruise-nozzle Z deflection angle reaches its maximum (thrust directed downward) value. At this point, the hft-fan and lift-nozzle deflection angles are still set to their maximum value, and a different set of equations are used to determine the cruise-nozzle, lift-fan, and lift-nozzle thrusts. The thrust equation (eq. 8) is still applicable; none of the component thrusts are set to zero. With the cruise nozzle at its maximum deflection, the moment arm for the cruise nozzle is determined by: (17) LCN =- XcN sin 8CN -zCN cos 3CN The moment arms for the lift fan and the lift nozzles are determined with equations 9 and 10, with the nozzle deflection angles set to their maximum values. The pitching moment contributions of each nozzle appear in equation 18: (18) (TLF 1) LLF - (TLN1) LLN + (TcN 1 ) LCN = - AM As in equation 15, the net thrust deflection angle can be expressed in terms of the nozzle thrusts.
This equation incorporates the last known propulsion system parameter, the pilot's commanded value of ON: TLF 1 cos 5LF x + TCN 1 sin _SCN + TLN 1 cos 5LN (19) TLF 1 sin _iLF x + TCN 1 cos _CN + TLN1 sin 6LN = tan ON Equation 19 can be rewritten as: TLF 1 [cos _iLFx - sin 8LF x tan ON] + TLNl[COS 5LN - sin 8LN tan ON]
+ TCN 1[sin 5CN - cos 5CN tan ON] = 0 (20)
Equations 8, 18, and 20 can be solved for the three engine nozzle thrusts. If the lift-nozzle thrust or the lift-fan thrust resulting from these equations is below zero or above its maximum allowable value, then the equations for TTCMD and pitching moment are solved again, with the lift-fan or lift- nozzle thrust set to the appropriate limit. If this solution exceeds another thrust limit, then the thrust allocation is determined with TTCMD as the remaining constraint.
If the cruise-nozzle thrust calculated by equations 8, 18, and 20 is less than zero, or if the com- manded net thrust deflection angle is greater than or equal to 45 deg, then the cruise-nozzle thrust is set to zero, and the cruise-nozzle deflection angle is set to its maximum (downward) value. This is the third propulsion control case for manual thrust control. The total thrust equation is essentially the same as equation 8, although TLF 1 and TLN 1 are the unknowns: TLF1 TLN1 KAUG rlLF + _ = TTCMD (21) The lift-fan and lift-nozzle deflection angles are equal and determined by the pilot's commanded thrust vector angle. Pitch moment arms are calculated for the lift fan and lift nozzle from equations 9 and 10. The relationship between the lift-fan thrust, the lift-nozzle thrust, and AM to balance the air- craft in pitch is: (TLF 1) LLF - (TLN1) LLN = - AM (22) Equations 21 and 22 are solved simultaneously to find TLF 1 and TLN 1. If the calculated TLF 1 or TLN 1 exceeds the nozzle limits, that thrust is limited accordingly and TTCMD is used as the constraint in determining the thrust for other nozzle.
A different propulsion system control scheme is used for the APP/ATM submode and for the TRC mode. In these cases, the SCAS determines the forces (AX and AZ) and the pitching moment (AM) that the propulsion system must provide, and the propulsion system controller determines the appropriate nozzle deflection angles and nozzle thrusts.
The logic used in determining the propulsion system commands in the APP/ATM and TRC modes is similar to that used in the manual thrust modes. Once the lift fan is engaged, cruise-nozzle and lift-fan thrust alone are used for flightpath and speed control, with the fan nozzle set at its maxi- mum (rearward) deflection angle. Once the cruise nozzle reaches its maximum (downward) deflection, lift-nozzle thrust is increased to satisfy the SCAS commands. The cruise-nozzle thrust is decreased as the vertical thrust needed for flightpath control increases, until all of the engine thrust is being directed to the lift fan and the lift nozzles. At this point, the lift fan and lift nozzles, which have been held at their aft deflection limits, are free to deflect as needed; the engine thrust is allo- cated between these nozzles to satisfy powered lift and speed requirements in conjunction with the system's commanded nozzle defection angles. This final propulsion system arrangement is used throughout the low-speed semi-jetborne flight regime (typically < 120 knots IAS) and hover. Table 7 summarizes the thrust management solutions for these control modes.
For the first case with the lift fan engaged, the lift fan and lift nozzles are considered to be deflected fully aft, and the lift-nozzle thrust is set to zero. The vertical thrust equation is: (23) - TLF 1 cos _iLFx + TGELF + FCN z = AZ With the flightpath loop closed, the propulsion control system must compensate for thrust variations due to ground effects; the ground effect due to the lift-fan thrust is TGELF, and its value is deter- mined from the actual lift-fan thrust that was commanded in the previous computer cycle. The thrust components that will provide the commanded longitudinal force, AX, are given by:
(24)
TLF 1 sin 8LF x + FCN x = AX The moment arms of the lift fan and the lift nozzles are determined by equations 9 and 10, the same method that was used in the manual thrust cases. With the pitching moment increment due to ground effect represented as MGELF, the thrust contribution to pitching moment is: (25) TLF 1 LLF + MGELF- (FCNz)XCN + (FCNx)ZCN =- AM Using the SCAS values for desired longitudinal force, vertical force, and pitching moment, equations 23-25 are solved simultaneously for TLF 1, FCN x and FCN z. The vector sum of FCN x and FCN z determines the net cruise-nozzle thrust, TCN 1, and the ratio of FCN z to TCN 1 determines the deflection angle of the cruise nozzle. If the lift-fan thrust calculated in this manner exceeds the maximum allowable lift-fan thrust, the fan thrust is set to that maximum and FCN z is recalculated to provide the pitching moment for aircraft trim.
Once the aircraft has slowed to the point that a significant portion of the lift-fan thrust is required to maintain the commanded flightpath angle, the cruise nozzle reaches its maximum downward deflection angle because the pitch trim of the aircraft must be maintained. Once this occurs, the cruise-nozzle deflection is fixed at its maximum value and a different set of equations is used to determine the lift-fan, lift-nozzle and cruise-nozzle thrust. The lift fan and lift nozzle are still deflected fully aft. The vertical, longitudinal and pitching moment thrust equations that determine TLF1, TCN 1 and TLN 1 are: (26) - TLF 1 cos 8LF x + TGELF - TLN 1 cos 5LN + TGELN - TCN 1 sin 8CN = AZ (27) TLF 1 sin 5LF x + TLN 1 sin _LN + TCN1 cos _CN = zSX (28) TLF1 LLF + MGELF - TLN 1LLN + MGELN + (TcN 1)LcN = _ AM If one of the solutions to these three equations violates a thrust limit, then that nozzle's thrust is set to its limit and the zXX and zkM equations are solved to find the remaining nozzle thrusts. If a sec- ond thrust limit is encountered, the last thrust value is determined with the moment equation and the other thrusts set to the appropriate limits. In this way, as propulsion limits are reached, the control of the vertical axis is given up first, followed by control of the longitudinal axis. This axis priority pre- vents thrust commands that could send the aircraft into a violent pitch excursion.
Equations 26-28 are used for thrust allocation until the cruise-nozzle thrust solution is less than or equal to zero. At this point, new solutions involving not only nozzle thrusts but also lift-fan and lift-nozzle deflection angles are used. In the original design of the propulsion system, the longitudi- nal lift-fan deflection limit and the lift-nozzle deflection limits were not the same, and therefore it could not be assumed that these nozzles would ever be at the same deflection angle. To solve for the four unknowns with the two forces and one moment that the SCAS requires, the ratio of AX to AZ was used to determine the deflection angle of the lift fan. The longitudinal deflection angle of the lift fan is expressed as:
(29)
The lift-nozzle deflection angle, the tift-fan thrust and the lift-nozzle thrust are found by simultane- ous solution of the following vertical, longitudinal and pitching moment thrust equations: (30) -TLF 1 cos 5LF x + TGELF + FLN z + TGELN = AZ (31) TLF 1 sin 5LF x + FLN x = AX (32) TLF 1 LLF + MGELF - (FLNz) XLN + MGELN + (FLNx)ZLN = -zXlVl The lift-nozzle thrust and deflection angle are determined by the thrust vector components FLN x and FLN z. As in the case of equations 26-28, control over the longitudinal axis is relinquished first, followed by control over the vertical axis, as nozzle deflection limits and nozzle thrust limits are reached. This change in axis control priority reflects the low-speed flight conditions for which equa- tions 30-32 are used, as the engine must provide most or all of the aircraft's lift.
Control Selector A diagram of the control selector appears in figure 12. The control selector combines preliminary propulsion system and aircraft attitude commands to generate control surface deflections, final propulsion system nozzle deflections, reaction control system deflections, and final propulsion sys- tem thrust commands. As the response of the control surface and engine nozzle actuators was considered to be fast, the computational delay of the computer provided an adequate lag to duplicate the response of these components. No explicit model of these actuators was created or used.
13 -: Commands from the pitch SCAS drive the canard directly. The canard deflection limits are 30 deg trailing edge up and 10 deg trailing edge down. When the lift fan is in use, pitch commands are multiplied by the gain KLF and added to the lift-fan thrust from the propulsion controller, TLF 1 , to produce the final lift-fan thrust command. The value of KLF depends upon the current lift-fan thrust and the total core thrust, TTO T. If the core thrust currently diverted to the lift fan exceeds the core engine thrust diverted to the remaining engine nozzles, then: (33) TILF KAUG Otherwise, KLF is set to 1/30 of the core thrust that is currently being used to drive the lift fan.
The maximum lift-fan thrust is 17,400 lb KPLN, which is used to modify the thrust of each lift noz- zle, is set equal to half of KLF; thus, pitch commands transfer the same amount of core engine thrust between the lift nozzles and the lift fan. If the pitch RCS option has been selected, the switch IRCSP is set equal to one and the pitch commands, multiplied by KPRCS (41.67 lb/deg of canard deflec- tion), cause deflection of the pitch reaction control nozzles; the maximum pitch RCS thrust is 1,250 lb, directed either upward or downward depending on the sign of the pitch command.
Roll control is generated by the ailerons in wingborne flight and through thrust transfer between the lift nozzles in jetbome flight, scaled per degree of aileron deflection. The deflection limits for the ailerons are +30 deg. After taking into account lift-fan/lift-nozzle thrust transfer due to pitch com- mands, the amount of thrust to be transferred differentially between the lift nozzles is determined.
KRLN is determined by the current value of lift-nozzle thrust and TTO T. If the thrust of both lift nozzles exceeds the remaining core engine thrust, then:
(34)
KRLN = 1 (TTo T _ TLN) If the combined lift-nozzle thrust is less than the core engine thrust, then KRLN is set to 1/60 of TLN. The roll commands are summed with the lift-nozzle thrust determined by the propulsion control system to produce the fight and left lift-nozzle thrust commands. Each lift-nozzle thrust is limited to 12,000 lb. When the lift-nozzle deflection angle does not coincide with the vertical axis of the aircraft, differential lift-nozzle thrust produces uncommanded yawing moments on the aircraft. To counteract these moments, the system uses lateral deflection of the lift-fan nozzle. The interconnect gain between differential lift-nozzle thrust and lateral lift-fan nozzle deflection is KRY1, which is determined by: KRY1 = (YLN/sin (SLN) (35) XLF J A yaw RCS can be used instead of the lift-fan nozzle to produce yawing moments. If this option is selected, thenKRY1 is multiplied by (XLF / XRCS). Usually the roll RCS is not employed; however, when IRCSR is set equal to 1 to enable this system, the gain KRRCS is set equal to 83.33 lb/deg, with reaction control thrust limit of 2500 lb.
4 : Rudder deflection, differential li_nozzle deflection, and lateral lift-fan deflection all provide yaw control for the aircraft. Commands from the yaw SCAS directly drive the rudder, which is lim- ited to +30 deg of deflection. Five deg of differential lift-nozzle deflection are available for yaw axis control in the powered lift regime. It follows that KYLN is 1/30 of 5 deg; the product of the yaw command and KYLN is added to the lift-nozzle deflection determined by the propulsion control system to produce the final left and right lift-nozzle deflection angles. The yaw command also pro- duces lateral lift-fan deflection in the powered lift regime. KYLF is set to the same value as KYLN, and after KYLF is multiplied by the yaw command, the differential lift-nozzle thrust interconnect signal is added to it, resulting in the lateral deflection command for the lift-fan nozzle. The deflection of this nozzle is limited to +15 deg, positive deflection producing positive side force. If the IRCSY flag is set to one, the yaw RCS is used instead of lateral lift-fan nozzle deflection. KYRCS is set to 41.67 lb of RCS thrust per degree of yaw command. After the differential lift-nozzle thrust intercon- nect is added to this signal, the yaw RCS command is limited to +1,250 lb.
Propulsion System The propulsion system model developed for this experiment is a representation of a lift fan cou- pled to a lift/cruise engine as shown in figure 2. The 2-dimensional convergent-divergent (2D-CD) nozzle of the lift/cruise engine provides axial thrust in cruise and vectoring thrust during transition.
Thrust can be transferred from the 2D-CD nozzle to a pair of lift nozzles and to the lift fan to pro- duce longitudinal and vertical thrust during transition and hover.
The lift/cruise engine responds to total thrust command, TTCMD, which is calculated as a sum of the individual nozzle thrust commands that are generated from the control selector. The closed loop core thrust dynamic response is model by a second order transfer function with respect to a refer- enced core engine state, TTRef. TTRef is defined as the trim state of the core thrust response and is reset to the total thrust response, TTO T , whenever the engine reaches a steady state. An incremen tal thrust command, ATTcMD, is calculated from the total thrust command and the referenced thrust as shown in equation 37. This incremental thrust command is then fed through a second order transfer function which represents the closed-loop core engine dynamic characteristics, where the damping ratio is set to 0.707 and the bandwidth, O)core. A rate limit, "_Tmax ' and a lift/cruise engine thrust magnitude limit, TTmax, are included in the second order transfer function model to represent physical thermal limits of the propulsion system. Each individual nozzle's referenced thrust and the response due to perturbed lift/cruise engine are distributed proportionally by the ratio of each nozzle thrust command over the total thrust command, equations 41 and 42.
(36) TTCMD = TCNcM D + TLLNcM D + TRLNcM D + TLFcM D / KAU G where KAU G is the lift-fan augmentation ratio (37) ATTcMD = TTCMD - TTRef
ATT = (s 2 0)c°re2 )ATTcMD (38)
+ 2_0)core s + 0)core 2 TTO T = TTRef + AT T (39) TiCM/) R i - TTCM D (40) where i = CN, LLN, RLN, and LF TiRef = TTRef R i (41) (42) ATi CORE = AT T Ri To model the thrust transfer rate limit effect, the thrust transfer command for each nozzle is defined as,
(43)
ATi.IF = TiCM D - TiRef - ATTcMD Ri • ° The thrust transfer rate for each nozzle, TiT V is limited to TTFma x.
The thrust response for each nozzle is calculated as shown in equation 44, in which rl is the thrust recovery factor for each specified nozzle.
(44)
TCN = ( TCNRe f + ATCNcoRE + ATCNTF ) x TICN TLL N = ( TLLNRef + ATLLNcoR E + ATLLNT F ) X TILN TRI2, l = ( TRLNRef + ATRI3qcoRE + ATRLNT z ) X TILN TLF0 = TLFRef + ATLFcoRE + ATLFTe For this experiment, a generic energy transfer between the lift/cruise engine and the lift fan is assum- ed. Dynamically, the lift-fan thrust at the lift-fan nozzle exit, TLF, due to TLF0 from equation 44 is dependent on lift-fan rotational speed and inlet guide vane setting. The dynamic response is modeled by a first order transfer function with a time constant of reEF, plus a first order washout transfer function, as shown in figure 13. A rate limit, "i'LFmax, and a maximum lift-fan thrust limit, TLFmax, are included to represent physical constraints of the lift-fan dynamics. The upper and lower IGV authorities are defined as shown in equation 45, in which KIGVu and KIGVL are in percent of lift-fan inertia thrust.
TIGVU = KIGVuTLFs, TIGVL = KIGVL TLF s (45)
The total li_fan thrust, TLF, is then calculated from the fan inertia thrust and IGV thrust as shown in equation 46, where rlLF is a generic thermal/mechanical efficiency constant.
TLF = ( TLF s + TLF F ) x rlLF (46) The mass flow rates of the primary inlets and auxiliary inlets are functions of the total thrust, maximum lift/cruise engine thrust, and respective maximum mass flow rate as shown in equations 47 and 48. A maximum mass flow rate of 11.2 slug/sec was used in each case. Primary inlets are opened in CTO and APP, and are closed in TRC. Auxiliary inlets are only opened in TRC to reduce the amount of hot gas reingestion. The mass flow rate of the lift-fan inlet is modeled as a function of the lift-fan thrust, maximum lift-fan thrust, and the maximum lift-fan mass flow rate (eq. 49) which was set to 14.0 slug/sec.
Iilpi = TT_OT 1 rilPima x
(47)
ITMAX)
( TTOT llilAima x
(48)
rnAi
=
)
N (49) rnLFI _. TLFMAX _ The resulting propulsion forces and moments are resolved into the aircraft body axes from each indi- vidual nozzle thrust and deflection, inlets, and reaction control system (RCS) when it is activated.
17 _ Axial force: (50) FEX = FCN x + FLLNx + FRLNx + FLF x + Fpi x + FAI x +FLH x where FCN x = TCN cos _SCN FLLNx = TLL N sin _iLN FRLN x = TRLN sin _iLN FLF x = TLF sin 8LF x cos _SLFy FHx = -riaH UB FAI x = -riaAi UB FLFI x = mLFI ( -UB - QB zpI ) Side force: (51) FEy = FLFy+ Fpiy+ FAIy+ FLFIy + FRCSy where FLFy = TLF sin 8LFy ' FpIx = nan ( -VB + PB zH - RB xpI ) FAI x = fnAI ( -VB + PB ZAI - RB XAI ) FI.FI x = fnLFl ( -VB + PB ZLFI - RB XLFI ) Normal force: (52) FEZ = FCN z + FLLN z + FRLNz + FLFz + FpI z + FAI z + FLHz where FCN z = TCN sin t3CN FLLNz = TLL N cos 5LN FRLNz = TRLN cos t_LN FLF z = TLF cos 5LF cos 5LFy FHz = -riaH WB FAI z = -rhAI WB FLFIz = riaLFI (-WB + QB xpi ) Rolling moment: (53) TEL =- FLLN z YLLN + FRLN z YRLN + FLFYz ZLF + FRCSY z ZRCS + FRCSy ZRCS FRCS z YRCS - Fpiy zpi - FAIy ZAI - FLNy ZLFI Pitching moment: (54) TEM = - FCN z xCN - FLLN z XLLN.+ FRLN z XRLN - FLF z XLF + FCN x ZCN + FLLNx ZLLN + FRLNx ZRLN - FLFx ZLF - Fpi z xpI - FAI z XAI - FLFI z XLFI + Fpi x zpI - FAI x ZAI - FLFI x ZLFI Yawing moment: (55) TEN = FLLNx YLLN- FRLN x YRLN + FLFYz ZLF + FRCSY z XRCS + FLFy XLF + Fply xp! + FAIy XAI + FLFIy XLFI Head Up Display A head-up display (HUD), that has been employed by NASA in several previous V/STOL simu- lations, provided the primary flight display for this experiment. The display is described here in gen- eral terms. The reader should consult reference 7 for a complete description of symbology and drive laws.
Separate presentations were adopted for transition and hover. For the transition phase, shown in detail in figure 14, the display was a flightpath centered, pursuit presentation that enhanced the external visual cues, centered them on the aircraft's flightpath, and presented the pilot with a pursuit tracking task for following the intended transition and approach guidance to a final hover point.
Course and glideslope guidance was provided in the form of a lead (ghost) aircraft that followed the desired flight profile. The pilot's task was to track the ghost aircraft with the flightpath symbol. As indicated in reference 7, the flightpath symbol was quickened to compensate for lags in the airframe and propulsion system response. For the MTV or FTM control modes, where thrust was controlled manually using the throttle, the flightpath compensation included lagged pitch rate and washed out throttle commands in combination with the true flightpath in accord with the following equation:
Jt.S+GwJ
and _w -.ZWhove r + Vaf Z'w ZWhover = g(l_dW) Z'w = 1.58 pgS/W K6T = A_3 T sin 0j / _w The pitch rate term was blended out for speeds below 55 knots and true airspeed was frozen at 100 ft/sec for speeds less than 100 ft/sec. Gains and washout frequencies were Kq = 1, A 5_ = 0.3. For • • T ATM, the flightpath was complemented with its commanded value in the short term according to s tan_l (1:1 / VT ) (57) Tq = tan-1 _-_-- + The washout frequency coy was 1 rad/sec. True lateral flightpath was represented by the flightpath symbol. Deceleration guidance was presented by an acceleration error ribbon on the right side of the flightpath symbol which the pilot nulled to follow the deceleration schedule. Situation information that accompanied the flightpath and ghost aircraft symbology included aircraft attitude, speed, alti- tude, engine rpm, thrust Vector angle, longitudinal acceleration, heading, and distance to the hover point.
During the latter stages of the deceleration as the aircraft approaches the intended point of hover, selective changes are made to the approach display to provide guidance for the hover point capture.
Specifically, the longitudinal velocity vector, predicted longitudinal velocity, and station keeping cross appear referenced to the vertical velocity diamond symbol as shown in figure 15. The drive law for predicted longitudinal velocity is shown below in the discussion of the hover display (eq. 58).
The pilot controls the predicted velocity toward the station keeping cross position and adjusts veloc- ity to bring the cross to rest at the reference hover point indicated by the cross being adjacent to the vertical velocity diamond. Once the aircraft is stabilized in this condition, the pilot is ready to per- form the vertical landing.
For the vertical landing, including recovery to the ship, the HUD format superimposed vertical and plan views and provided command and situation information in a pursuit tracking presentation (fig. 16). In the horizontal situation, the aircraft velocity vector was represented by a line emanating from the aircraft symbol. A pad symbol represented the landing area. Horizontal and vertical velocity predictor symbols, whose displacement and orientation from the aircraft symbol indicated magnitude and direction provided the pilot lead information for hover maneuvering. Predicted horizontal and vertical velocity presentations were compensated for aircraft and propulsion system lags as they were for transition. In this case, for MTV mode, true velocities were complemented with transla- tional accelerations and washed out control commands according to the relationship: 2O
T9 s - K0c0c (58)
Vxc= Vx+ -g
s+I/T 6 s s+l/T 2 where CYu = X Uhover + Vaf X' u XUhove r = g(riae/W) X' u = CD aero p gS/W T_ = 1.11 T6 = 10 Koc = 3.45 T2 = 10 For TRC mode, the commanded horizontal velocities were displayed directly. The vertical situa- tion was displayed by a diamond referenced to the left leg of the aircraft symbol whose displace- ment, t w , was proportional to complemented vertical speed. For MTV mode vertical velocity is c complemented with vertical acceleration so that (59)
t.: wrh s
L T6s+I L K6TSTT6 +_w where Kw = 0.2 and the other terms have the same values as for the flightpath symbol drive. When TRC is engaged vertical velocity was complemented with washed out vertical velocity command to provide the command
(6O)
twc: wEh+ hC/s 1]
and my = 1. A horizontal bar indicated the altitude remaining to touchdown. Attitude, air velocity, engine rpm, thrust vector angle, heading, vertical velocity limits, and wind direction are provided as situation information.
CONCLUSIONS A simulation model of an integrated _ght/propulsion control system, propulsion system, and head-up display has been developed for an advanced, short takeoff, vertical landing lift-fan fighter aircraft. These models, combined with a model of the aircraft' s aerodynamic characteristics have been employed in piloted evaluations of takeoff, transition, hover, and landing characteristics of this class of aircraft. The flight/propulsion control system includes modes for several response types which are coupled to the aircraft's aerodynamic and propulsion system effectors through a control selector tailored to the lift-fan propulsion system. Design of the system has drawn on experience gained from simulation of earlier STOVL configurations, including the Mixed Flow Vectored Thrust and Augmentor Ejector concepts as well as from the VSTOL Systems Research Aircraft, which is a modification of the YAV-8B Harrier prototype. Propulsion system components modeled include a remote lift-fan and a lift/cruise engine. Their static performance and dynamic response are repre- sented by the model. A separate report describes the subsonic, power-off aerodynamics and jet induced aerodynamics in hover and forward flight, including ground effects.
J_ REFERENCES 1. Engelland, S. A.; Franklin, J. A.; and McNeill, W. E.: Simulation Model of a Mixed-Flow Remote-Lift STOVL Aircraft. NASA TM- 102262, Feb. 1990.
2. Birckelbaw, L. G.; McNeill, W. E.; and Wardwell, D. A.: Aerodynamics Model for a Generic ASTOVL Lift-Fan Aircraft. NASA TM-110347, Apr. 1995.
.
Franklin, J. A.; Hynes, C. S.; Hardy, G. H.; Martin, J. L.; and Innis, R. C.: Flight Evaluation of Augmented Controls for Approach and Landing of Powered-Lift Aircraft. AIAA Journal of Guidance, Control, and Dynamics, vol. 9, no. 5, Sept.-Oct. 1986.
o Franklin, J. A.; Stortz, M. W.; Gerdes, R. M.; Hardy, G. H.; Martin, J. L.; and Engelland, S. A.: Simulation Evaluation of Transition and Hover Flying Qualities of the E-7A STOVL Aircraft. NASA TM-101015, Aug. 1988.
° Franklin, J. A.; Stortz, M. W.; Engelland, S. A.; Hardy, G. H.; and Martin, J. L.: Moving Base Simulation of Control System Concepts and Design Criteria for STOVL Aircraft. NASA TM-103843, June 1991.
° Chung, W. W.; McNeill, W. E.; and Stortz, M. W.: A Direct Application of the Non-Linear Inverse Transformation Flight Control System Design on a STOVL Aircraft.
NASA TM-108808, May 1993.
7. Merrick, V. K.; Farris, G. G.; and Vanags, A. A.: A Head Up Display for Application to WSTOL Aircraft Approach and Landing. NASA TM-102216, Jan. 1990.
22 •
Table 1.Aircraft geometry
55.4 ft
Overalllength
14.16 ft
Overallheight
Area 523.3 ft 2
Wing
36.17 ft Span 18.42 ft Mean aerodynamic chord 2.50 Aspect ratio Leading edge sweep 40.0 deg Trailing edge sweep 30.0 deg Airfoil NACA 64A005 243.1 ft 2
Canard Area
24.65 ft Span 12.55 ft Mean aerodynamic chord 2.50 Aspect ratio Leading edge sweep 40.0 deg Trailing edge sweep 30.0 deg Airfoil NACA 64A004.5 Area 39.0 ft 2
Vertical tail (each)
6.98 ft Span 7.11ft Mean aerodynamic chord 1.25 Aspect ratio Leading edge sweep 40.0 deg Trailing edge sweep 30.0 deg Airfoil NACA 64A004.5 Table 2. Mass properties 30,000 lb Weight 373.3 in.
x cg location 0.0 in.
y cg location 96.0 in.
z cg location Pitch moment of inertia 91,200 slug-ft 2 Roll moment of inertia 14,300 slug-ft 2 Yaw moment of inertia 101,000 slug-ft 2 Product of inertia 0 slug-ft 2
Table3. Flight controlmodes
Controlaxis
Control modedesignations (applicableflight phases)
TRC
CTO (wingbome MTV (transition,hover) APP:ATM
(hover)
flight) APP:FTM (transition) (transition,hover)
Pitch/roll Rate command - Rate command - attitude Rate command -
attitude hold hold, blend to attitude attitude hold, command blend to attitude command Yaw rate Yaw Sideslip command Sideslip command, blend Sideslip command, command to yaw rate command blend to yaw rate command Vertical Aerodynamic lift Thrust magnitude Flightpath command, Velocity command blend to velocity command Acceleration Thrust magnitude Thrust vector angle Velocity Longitudinal command - command (MTV), acceleration velocity hold command - velocity hold (APP) Lateral Velocity command
Table4. Controlmodegains
Attitudecontrol
Pitch
Roll Control limits = _+5.65 in.
Control limits = +4.2 in.
Forcegradient= 1.0lb/in.
Force gradient = 0.7 lb/in.
Breakout= 0.15 in.
Breakout = 0.05 in.
K111= 0.28rad/sec/in.
K1 = 1.2 rad/in.
K222= 0.3 rad/in.
K333 = 1.6 rad/sec/in.
K0 = 4.0 rad/rad K¢ = 9.0 rad/rad K0 = 4.0 sec KS = 6.0 sec K300 = 20.0/(1 + 0.0047_1) deg/rad K3 = 15.0 (100/KRLN)/(1 + 0.029q) deg/rad 7:9 = 0.05 sec 7:8 = 0.05 sec Yaw - transition Yaw - Hover Control limits = +2.12 in.
Same Same Force gradient = 14.0 lb/in.
Breakout = 0.1 in.
Same K10 = 0.3 tad/in. K10H = 2.3 K_ = 2.8 sec K_= 4.0 sec KI3 = 40 rad/rad K30 = 40.0 deg/rad 7:[3= 0.25 sec KB = 0 (V < 40 knots); 1.0 (V > 50 knots) 7:7 = 0.05 sec Velocity control Longitudinal velocity Vertical velocity Control limits K 7 = 0.00545 rad/deg Stick = +2.25 in.
Kw = 0.71 sec -1 Thumbwheel = +100 deg K3W = 0.14 Breakout 7:CNT = 0.1 sec Stick = 0.225 in.
Thumbwheel = 5.5 deg Lateral velocity K2U = 0.1 K6 = 10.0 ft/sec/deg KVI = 20.0 sec -1 PR = 5.25 KV = 14.0 flYsec2/in.
K_2 = 0.58 rad/ft/sec KU = 0.69 sec -1 K9 = 0.285 rad/rad K3u = 1.0 7:U = 0.35 sec 25 ;_ Table 5. Maximum and minimum throttle control commands and control sensitivity slopes _cmd Maximum Minimum Max slope u (1/Sth) Min slope u (1/Sth) 1/% 1/% Thrust magnitude, % 100 3 1.0 0.5 Flightpath angle, ?, 10 -20 0.3 0.287 deg Vertical velocity, la, 20 -40 0.6 0.5 ft/sec Table 6. Propulsion management summary for MTV and APP:FTM modes Equations in Case TCN 1 _CN TLF1 8LFx (=SLN) TLN1 text 1. Lift fan Unknown Unknown Unknown Maximum 0 8, 13, 15, 16 engaged and O N < 45 deg 2. Case 1 Unknown Maximum Unknown Maximum Unknown 8, 18, 20 8CN > max.
3. Case 2 21, 22 0 Maximum Unknown ON Unknown TCN 1 < 0 or ON > 45 deg
Table7. Propulsionmanagement summaryfor APP:ATM andTRC modes
Case
Equations
TCN1 [iCN TLF1 5LF x TLN1 [iLN
in text 1. Lift fan Unknown Unknown Unknown Maximum 0 Maximum 23,24,25 engaged 2. Case 1 Unknown Maximum Unknown Maximum Unknown Maximum 26, 27, 28 SEN > max.
3. Case 2 0 Maximum Unknown Unknown Unknown Unknown 29,30,31, TCN 1 < 32 ":';_i_..'.'_i._:i:i::;:::; ....
iiii_: •_..=_, _,.'_..
•_.'._ _ • " _._ iiii_._s_ • • :. ._:_
_i _iiiiiiiiiiiiiiiiiiiiiiiiiiiiii_iiiiiii
i!!i!i!!!ili!!i!iiii_iii?!ii!ii
.... _iiiiiiiiiiiiiiiiiiiiiiiiii_i_iii?:: %iii25iiii2ii:isi:i:_ :_iiiiiii!ii: Figure 1. Views of the ASTOVL Lift Fan Aircraft.
Lift Fan I I
I I
I I Lift Fan Nozzle Figure 2. Propulsion system configuration.
Sensors Aero controls LI Pilot Regulator = Control Aircraft v I selector commands I Propulsion_ I controls Figure 3. Flight-and-propulsion control system structure.
4' Wingbome/Jetbome WOW Any < 250 kts, > 150 kts < 60 kts Airspeed: > ]250 kts < 150 kts, > 60 kts Ai_peed Lift Fan Off CTO CTO CTO CTO CTO Allowable Mode Switching
11 \\ I
MTV _ APP MTV _ APP MTV Lift Fan On • When airbome/jetbome, the pilot cannot manually disengage the lift fan until the total thrust vector is deflected fully aft • Disengagement of the nozzle lever clutch in APP or TRC mode causes reversion to the MTV mode ° Switching between FTM and ATM submodes within APP not shown as this change does not depend on airspeed Figure 4. Control mode selection logic.
3O e I i+ qB y--__
[Z -I 1-
I If wow on hold past value r bo I;9S+ 1
TR, M --_--I_
SWITCH If wow on or flaps up initialize to K e e +I STICK _._PITCH CM D I limit logic t 4- Figure 5. Pitch stabilization and command augmentation system.
+ P If wow on B hold past value TRIM SWITCH 1;8S+ 1 If wow on or flaps up initialize to K_ + limit STICK D -_ ROLLcM logic + +_
Tc
r ate limiter
I
S (1_ -PR)+K_ $2+ _ S + K( h +
Vy
+
Vy
Figure 6. Roll stabilization and command augmentation system.
1 -KB I P
ICn0/b'2v/K
r+
IOnr/b'2v/l_
-i- t_O L_ t Cn(5 a)
E
E
PEDALS_
t_ _ t_,ot •
o.I
+ ) logic __[limit m Figure 7. Yaw stabilization and command augmentation system.
THROTTLE T cmd IT max
LEVER \"
G hprime
F
1Jb F 4_ cmd ) limit logic +
Vx
F
1 I
I;cNTS + I h oo h K WWD Figure 8. Vertical velocity stabilization and command augmentation system.
NOZZLE @N LEVER Vx
Vx
F 0 • + -I
ST,OK
S÷Kvl _-_ KU I__+
Kvt
t_
°'_ 1
I "l;uS + 1 I limit " logic I
I
E_ Oe E +I STICK logic Xcmd THUMBWHEEL 1 + Figure 9. Long#udinal velocity stabilization and command augmentation system.
(5 cmd) max ( 6 ) @mode switch Slope u (5 cmd) ref I
I
I Slope L [ (5 cmd) min I I (_th) ref (_th) @mode 100 switch Throttle Control Input in % Travel Figure 10. Throttle control/command scale.
-.,, xCN XLF zCN Lift Fan Body-Fixed Lift-Cruise Engine X Axis ZLN CG ZL / k --- XLN Body-Fixed Z Axis Note: Distances measured forward of the center of gravity are considered positive.
Distances mesured below the center of gravity are considered positive.
xCN = - 19.72 fl 8CN range = +20 °, -20 ° zCN = 0 ft _LF range = +45 °, -10 ° XLF = 11.70 ft 8LN range = +45 °, -10 ° ZLF = 1.93 ft XLN = -8.93 ft ZLN = 1.41 fl Figure 11. Propulsion system dimensions and sign conventions.
, IRC_/0
' IKPRCSl --'/ 1 _SPRCS
_--_ T_-c_ m r_ h KRLN ROLLclvlD _LN
B
--_SRLN
I I _/10 _8LFy
[i_,_ UN]
YAW CMD
D
[KY
[
XLF / xRCS [
+l
-_KYRCS __----.f 1 SYRCS +___"_ 8LIN Figure 12. Control selector.
TIGV u
-I
Figure 13. Lift fan inertia and IGV modeL (heading tape)
,10
(wind direction I and magnitude) 05 06 07 08 E (pitch ladder)
I I I I I I
20 (airspeed in knots) A
100 0 (altitude in feet)
200 1.TM
(range in mile)
-1 80 (dirnb rate
in ft/min)
A
! (control mode indicator)
2_
(landing deck) V ( _ bracket) (horizon) (acceleration error ribbon) I A ateral acceleration ball) (-3 °glide slope) - - _ _j_ &_ ___host a/c)
>
.6.__1 • .__45 (thrust vector (RPM %)_ m_,..,. (longitudinal ,,_.., .._ angle in degrees) acceleration in ft/sec 2 ) Figure 14. Head-up display approach mode.
, 39 I
,10
05 06 07 08 E 10 I I I I I I
A
5 0 (range in feet)
900 200
-120
(station keeping point, Sl_) I I (longitudinal velocity 0 predict or ball) (Iongit udinal velocity vect or) Figure 15. Head-up display approach mode for station keeping point capture.
4O
,10
05 06 07 08 E 10 I I I I I I
A
20 100
-1 20
(ground velocity 18
H
in kts) (control mode indicator) C) (velocity predictor ball) (v elocit y vect or) =__(a/c trident) (wind direction) _//_ (direction of Sl_ .I.
if SI_ not in sight) velocity) ( SKP moves to the TD point on the deck _'! _(vert ical when a/c is closing in to the deck) (deck and allowable sink I_1 rate ribbon are displayed .(_8_.=,,,_ lu-&"'""'_hen a/c is over the dec"k'_'=w Figure 16. Head-up display hover mode.
Form Approved REPORT DOCUMENTATION PAGE oMBNoo7o4-o188 Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington. VA 22202-4302, and to the Office of Management and Budget. Paperwork Reduction Project (0704-0188), Washington. DC 20503.
1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED April 1995 Technical Memorandum 4. TITLE AND SUBTITLE 5. FUNDING NUMBERS Simulation Model of the Integrated Flight/Propulsion Control System, Displays, and Propulsion System for an ASTOVL Lift-Fan Aircraft 505-68-32 6. AUTHOR(S) William W. Y. Chung, Paul E Borchers, and James A. Franklin 18. PERFORMING ORGANIZATION 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) REPORT NUMBER Ames Research Center Moffett Field, CA 94035-1000 A-950048 10. SPONSORING/MONITORING 9. SPONSORING/MONITORING AGENCY NAME(S) ANDADDRESS(ES) AGENCY REPORT NUMBER National Aeronautics and Space Administration NASA TM-108866 Washington, DC 20546-0001 '11. SUPPLEMENTARY NOTES Point of Contact: Paul F. Borchers, Ames Research Center, MS 211-2, Moffett Field, CA 94035-1000; (415) 604-6115 12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE Unclassified -- Unlimited Subject Category 05, 07, 08 13. ABSTRACT (Maximum 200 words) A simulation model has been developed for use in piloted evaluations of takeoff, transition, hover, and landing characteristics of an advanced, short takeoff, vertical landing lift fan fighter aircraft. The flight/ propulsion control system includes modes for several response types which are coupled to the aircraft's aerodynamic and propulsion system effectors through a control selector tailored to the lift fan propulsion system.
Head-up display modes for approach and hover, tailored to their corresponding control modes are provided in the simulation. Propulsion system components modeled include a remote lift and a lift/cruise engine. Their static performance and dynamic response are represented by the model. A separate report describes the subsonic, power-off aerodynamics and jet induced aerodynamics in hover and forward flight, including ground effects.
14. SUBJECT TERMS 15. NUMBER OF PAGES STOVL, IFPCS (integrated flight/propulsion control system) 16. PRICE CODE A04 17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT OF THIS PAGE OF REPORT OF ABSTRACT Unclassified Unclassified NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89) Prescribed by ANSI Std. Z39-18