Document
United States Patent [191 t111 4,373,184
t451 Feb. 8,1983
Lambregts
3,681,580 8/1972 Gwathmey et al. ................ 364/434
VERTICAL FLIGHT PATH STEERING
3,688,175 8/1972 Rauschelbach ................. 318/584 X
SYSTEM FOR AIRCRAFT
3,705,306 12/1972 Lydon et al. ....................... 364/430
Inventor: Antonius A. Lambregts, Renton, 3,752,967 8/1973 Victor ............................. 364/430 X
Wash. 3,781,628 12/1973 Rauschelbach ..................... 318/584
3,807,666 4/1974 Devlin ............................ 244/196 X
Assignee: The Boeing Company, Seattle, Wash.
3,848,833 11/1974 Rauschelbach ..................... 364/110
3,920,966 11/1975 Knemeyer ........................... 318/584
Appl. NO.: 107,464
3,927,306 12/1975 Miller .............................. 244/186 X
3,955,071 5/1976 Lambregts ...................... 244/182 X
Filed: Dec. 26, 1979
3,967,799 7/1976 Muller ............................. 364/430 X
Int. ( 3 . 3 ....................... G06F 15/50; G06G 7/78; 3,970,829 7/1976 Melvin ................................ 364/434 G05D 1/08 Primary Examiner-Joseph F. Ruggiero U . S . c1. .................................... 364/434; 244/196; Attorney, Agent, or Firm-James P. Hamley; Bernard A.
340/27 AT; 364/430 Donahue Field of Search ............... 364/434, 430, 428, 443; 244/181, 185, 186, 187, 188, 194, 195, 196, 197; 1571 ABSTRACT 340/25, 27 R, 27 NA, 27 AT; 318/583, 584 Disclosed is a vertical flight path angle steering system References Cited for aircraft, utilizing a digital flight control computer which processes pilot control inputs and aircraft re- U.S. PATENT DOCUMENTS sDonse Darameters into suitable elevator commands and 2,553,983 5/1951 Saxman, Jr. .................... 3181583 X cbntrofinformation for display to the pilot on a cathode 2,998,945 9/1961 Meyers ................................ 244/196 ray tube. The system yields desirable airplane control 3,096,955 7/1963 Priestly ............................... 244/188 handling qualities and responses as well as improve-
3,106,903 10/1963 Bentkowsky et d. .......... 3181584 x
oper- ments in pilot workload and safety during 3,386,689 6/1968 Parker et al. ................... 244/184 X Hendrick ........................ 244/187 X ation in the area and under 'Ondi- 3,399,849 9/1968 3,422,418 1/1969 Simoneau ........................ 244/197 X
3,521,227 7/1970 Congleton et al. ........... 34/27 NA
3,521,839 7/1970 Diani ............................... 244/184 X 10 Claims, 14 Drawing Figures
U.S. Patent Feb. 8, 1983 Sheet 1 of 6 4,373,184
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U.S. Patent Feb. 8, 1983 Sheet 3 of 6 4,373,684
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over ground track in the horizontal plane. This system VERTICAL FLIGHT PATH STEERING SYSTEM is called a Velocity (Vector) Control Wheel Steering FOR AIRCRAm System because it provides control of the total velocity vector of the aircraft relative to the earth. The pilot’s BACKGROUND O F THE INVENTION 5 column wheel inputs are processed in a computer to The invention described herein was made in the per- Provide a ProPortional rate of change to the control formance of work under a NASA contract No. NASI- parameters. The pilot closes his control loop by evduat- 14880 and is subject to the provisions of Section 305 of ink2 the changes in flight Path angle andlor ground track the National Aeronautics and Space Act of 1948, Public resulting from his control input on electronic cathode Law 85-568 (72 Stat. 435; 42 USC 2457). 10 ray tube (CRT) displays. With no pilot control input the This invention pertains to the aircraft art and, more last established values of these control parameters are particularly, to an improved control wheel steering maintained by automatic control. In concert with the apparatus for the same. control variables also other pertinent information is The recent economic pressures to increase operating displayed on these CRTs, so that the pilot can evaluate efficiencies of commercial transports have resulted in l 5 the complete vertical and horizontal airplane situation.
more stringent requirements for airplane control, more The main objectives of such a system are to provide sophisticated air traffic control procedures and closer the airplane with superior control handling characteris- adherence to published schedules, regardless of occa- tics, simplify manual airplane maneuvering on complex sional overcongested terminal areas and marginal flight paths, increase control precision and provide weather conditions.
2o airplane path tracking stability unaffected by speed These developments have had a tendency to increase changes, configuration changes, turbulence or varying Pilot workload and, consequently, to degrade flight winds. FIGS. 1 and 2 illustrate the nature of the prob- safety. Automation of functions or operations can re- lem.
duce the pilot workload associated with certain tasks. FIG. 1, an aircraft 10 is shown approaching a For example, the current trend in the development of 25 runway 12 on a desired glide slope 14. The inertial automatic flight control systems is toward auto- velocity vector V I of the craft 10 is the resultant of an mated guidance and taking performance air speed velocity vector V A and a wind velocity vector mization and air traffic constraints into consideration.
vw, here indicated as being a headwind. FIG. illus-
Also, a careful simplification and integration of new and trates the aircraft approaching the runway 12 on a existing control and monitoring functions is required to 30 desired glide slope 14 tailwind landing conditions.
prevent the ever increasing number of new tasks from The problem faced by the pilot in either of these increasing the overall pilot workload, and allow the conditions is that of maintaining desired inertial flight pilot to take maximum of the sophisticated on-board equipment during manual control operations. Path Y I changing the attitude and A relatively large percentage of the flight operations 35 power setting*Forces caused by tWhlence and winds- remain dependent on the pilot’s ability to fly the air- hear further add to this problem creating a high pilot plane manually in a sophisticated operating environ- workload during landing conditions. If the pilot can merit. This is so due to the lack of (or inadequate) simply establish the desired flight path using his co- ground guidance facilities (ILS, MLS, D M E ~ , vops, lumnlwheel and corresponding displays, and the auto- etc.), equipment breakdowns, or abnormal air traffic 40 matic System will capture and track this desired flight control situations. These conditions can generate an path, regardless of the above mentioned disturbances, a almost instantaneous diversion of pilot attention and substantial pilot workload reduction will result.
increase the pilot’s workload. Potential encounters with An early conceptual design for the longitudinal Part windshear further increase the risk associated with such of this controlldisplay system is shown in FIG. 3. Here, 45 the pilot of the aircraft sets a desired flight path angle conditions.
The capability for the pilot to assume manual control Yc by displacement of his column control, indicated of the aircraft in such an environment, at any time, generally at 20. The position of the column is sensed by while maintaining the same safety and performance pick-ups which produce a corresponding electrical sig- levels with minimal increase in pilot workload, is nal &OI. The column signal 6coiisintegrated in the stan- needed in commercial transport airplanes to comple- 50 dard integrator 22 thereby producing a signal yc repre- sentative of the commanded flight path angle.
ment the automatic flight modes.
A first contribution toward meeting this need was the The column signal Gcorisalso fed to the first input 24a design of a semiautomatic airplane control mode called of the control law circuit 24. The control law circuit Attitude Control Wheel Steering, which gives the pilot also receives at a second input 24b a signal representa- control over the airplane through his column and 55 tive of the rate of change of aircraft pitch angle. This wheel, and provides improved control responses and signal is available from sensors located on the aircraft.
automatic tracking of the aircraft’s attitudes as estab- Coupled to the third input 24c of the control law circuit lished by the pilot. 24 is the output from a summer circuit 26.
Although systems of this type provide improved In response to its input signals, the control law circuit performance and reduction in pilot workload compared 60 produces an output command signal 6, which controls to basic manual airplane control, it has been recognized airplane dynamics, indicated generally at block 28. The that a more advanced semiautomatic control mode is airplane dynamics respond to the elevator command needed to further reduce pilot workload and minimize signal 6, in such a manner that the craft assumes the the transition problem from fully automatic to com- desired flight path angle. The actual inertial flight path puter augmented manual control.
65 angle is sensed by the aircraft and applied to the input of To meet this need, a control and display system was a cathode ray tube display 30 and to the negative input devised that gives the pilot direct control over the earth 26a of the summer 26. Applied to the positive input 26b referenced flight path angle in the vertical plane and of summer 26 is the reference flight path angle yo Thus,
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summer 26 produces at its output 26c an error signal Briefly, according to the invention, an aircraft con- trol and display system includes a control reference which, when fed back to the control law circuit at its third input 24c. causes proper response of the output means for receiving and processing a signal representa- tive of the pilot’s steering input and developing a con- signal 6e such that the airplane’s flight path angle is corrected to the desired value. trol reference signal proportional to the time integral thereof. Aircraft sensing means produces a signal repre- This prior art design did not fulfill the expectations for a Velocity Vector Control Wheel Steering system. sentative of the aircraft’s controlled state. The control reference signal and the signal produced by the air- Pilots found it difficult to set up the desired flight craft’s sensing means are processed through control path angle and were repeatedly drawn back into the command means to produce a control surface command control loop because the displayed flight path angle 10 symbol would not stay at the target value. The objec- that controls the craft’s dynamics to smoothly bring the tive of this mode-pilot workload reduction-was thus controlled state to track the control reference value. A provided display displays both the control reference not realized.
It was found that major deficiencies of this prior value and the aircraft controlled state to the pilot.
system could be identified from the time response, an l5 BRIEF DESCRIPTION OF THE DRAWINGS example of which is given in FIG. 4.
The lag between the column input and the response FIGS. 1 and 2 illustrate the changes in relative rela- of the flight path angle symbol on the CRT display tionships between the airplane state variables that must was too long. As a result, the pilot could not occur when the airplane maintains a constant inertial readily assess the long term effect of his inputs and flight path angle while the horizontal wind component was unable to control the flight path angle pre- changes amplitude and/or direction; FIG. 3 shows an early conceptual design of a Veloc- cisely.
The transient response of flight path angle to a col- ity Vector Control Wheel Steering Control/Display system; umn input exhibited a relatively large overshoot of 25 FIG. 4 illustrates the deficiencies in the time response the steady state value. This short term response of a prior art system for step column inputs; characteristic resulted in a tendency to reverse the FIG. 5 gives an example of ideal response of the flight column inbut. which in turn made it difficult to path angle to a column input of constant amplitude and establish t6e desired long term value of the flight limited duration; path angle.
The control response was insufficiently damped, con- FIG. 6 provides a conceptual block diagram of the tributing to the unsteadiness of the displayed flight improved flight path angle control wheel steering sys- tem, including display of the flight path angle command path angle symbol.
The unsteadiness of the display, which also resulted signal for closing the short term pilot control loop and from turbulence, made it difficult for the pilot to an improved flight path angle control law utilizing rate assess the long term flight path angle trend and of change of flight path angle; resulted in undesired pilot control intervention. FIG. 7 shows a detailed block diagram of the pre- ferred embodiment of the improved control law of the SUMMARY O F THE INVENTION invention; &, 86 and & are response characteristics illus- Accordingly, it is the object of this invention to pro- FIGS.
trating certain airplane dynamics as they relate to the vide an improved longitudinal Velocity Vector Control Wheel Steering system that overcomes the above men- response of the display signals and the desired charac- tioned problems associated with prior art control wheel teristics of both; steering systems. FIG. 9 illustrates the addition of the flight path angle It is further an object of this invention to provide a command symbol to the vertical airplane situation dis- Velocity Vector Control Wheel Steering system in play; and which the control and display system components are FIGS. 100, 106 and 1Oc illustrate actual system re- sponses for the improved flight path angle Control designed in an integrated way, to provide coordinated control and display responses throughout the flight Wheel Steering system design according to the present envelope with constant column force per unit normal invention.
acceleration and an invariant flight path angle response DESCRIPTION O F THE INVENTION time constant.
Yet a further object of this invention is to provide a The deficiencies of the prior art system, as identified computer augmented manual airplane control system from its response characteristics shown in FIG. 4, and suitable displays, that take advantage of the on- helped formulate the desired flight path angle control board avionics data and computing capability to reduce wheel steering system response characteristics and ulti- mately the system’s design requirements.
pilot workload by providing superior airplane control handling characteristics, simplifying manual airplane Ideally the flight path angle y should respond to a maneuvering capability on complex flight paths, in- column input of finite duration as shown in FIG. 5. For creasing routine control precision and by providing 60 the duration of the column input, the flight path angle airplane path tracking stability unaffected by speed increases with a rate proportional to the magnitude of the column input, then holds constant after the column changes, airplane configuration changes, turbulence input is reduced to zero. Since the airplane embodies a and windshear.
Another object of the invention is to provide the certain mass and inertia and limited elevator control desired display information and control characteristics 65 authority, the flight path angle will by necessity incur for an improved Velocity Vector Control Wheel Steer- some finite lag 765. The object of this invention is to provide a control law design having flight path angle ing by displaying both the actual flight path angle and the commanded flight path angle to the pilot. responses closely resembling that of FIG. 5, exhibiting
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minimal response lag, good response damping and proper rate sensitivity for a given column input. rhis signal GAMMA is earth referenced and may be The preferred arrangement of the control wheel obtained from well known prior art sensing and compu- steering system according to the invention is shown in tation sources. The yEsignal is used to command the 5 elevator through a proportional signal path having a the conceptual block diagram of FIG. 6. Here, the pilot signal amplifier 280, providing signal amplification by a generates the desired flight path angle command yc by displacement of his control column 120. Pickups at the factor KGEP. The ye signal is also processed in and Sc0i representative of integral signal path. For this purpose signal amplifier control column produce a signal column position. This signal is integrated in a conven- 290 provides amplification of the yEsignal by a factor tional integrator 122 to produce the signal ~~representa- 10 KGEI. This amplified yEsignal is summed in device 300 with the amplified GAMD signal. The resulting signal tive of the commanded flight path angle.
is integrated in device 310. The proportional ye signal A flight path angle error signal yEisformed by taking the difference between this commanded flight path path provides the main elevator control command for angle yc and the actual flight path angle y, as supplied reducing the yEsignal to zero. The yEintegral path is by airplane dynamics block 128, in combiner 126. The needed to assure that yEwill be nulled, even when a steady state elevator needs to be carried or to offset null error signal yE is used as the main input to the control errors in other signal sources making up the elevator law block 124. In the improved control law the flight path angle error signal yr is processed along with a command.
signal representative of the rate of change pf flight path The GAMD signal, which is gain weighted in circuit 20 320 by a gain factor KGDI before being summed with
angle q, column position ScOl, pitch rate 8 and ground
the gain weighted yEsignal in summer 300, is represen- speed VGto form the elevator command 8 e c The signals 3, 8 and VG are also produced from airplane dynamics tative of the rate of change of the actual flight path block 128. The proportional column position signal angle. It is produced in circuit 350 as the quotient of commands the eievator directly for the purpose of vertical acceleration and ground speed according to the achieving a direct and smooth initiation of the maneu- 25 relationship j-h/VGS. The integral signal path of the I ver. The elevator command signal 8 , is applied to the GAMD signal is provided to offset the y E signal integra- aircraft’s elevator control which generates the neces- tion after a change in ye Integration of yE by itself sary airplane dynamics to capture and track the com- would cause the ye target to be overshot during the transient response, because the output of integrator 310 manded flight path angle.
FIG. 7 is a detailed schematic diagram illustrating the 3 0 would have built up to the wrong value at the time the preferred construction of the control law circuit. yE signal would cross zero. With the GAMD signal The column position signal 8 c o ~ is first processed input to integrator 310 the integration stops when through a deadzone circuit 210. This is done to assure (KGEI)(ye)+(KGDI)(GAMD)=O, in other words, that the signal input to the control law is zero when the when the appropriate rate of change of gamma is estab- column is in the neutral position. The signal output from 35 lished.
the deadzone circuit 210 is next multiplied by a signal There is also a proportional signal path of GAMD, KC in multiplier circuit 220 producing a signal including gainflag circuit 330 with lag 765. This propor- DCOLV. A signal KC is produced from the ground tional GAMD signal path provides the main damping speed signal VGS in function generating circuit 230, term to the control of the flight path angle dynamics.
according to the relationship KC=VO/VGS. VO rep- 40 The lag filter cuts down high frequency noise generated resents a suitable normalization constant, here 120 kts. by the vertical acceleration sensor.
The speed programmed column signal DCOLV is The gain weighted yE signal from circuit 280, the summed with a signal GAE in summer 240 before being output from integrator 310 and the gain weighted and used in two signal paths. This GAE signal is normally filtered GAMD signal output from circuit 330 are zero and its function is described in connection with the 45 summed in circuit 400 to form the outer loop flight path GO AROUND feature of the control law. In the first angle control signal.
signal path the output of summer 240 is amplified and Inner loop pitch damping is provided by the pitch lagged in a circuit 250 and then integrated in integrator rate signal input (Q), which is first processed through a 260 to produce the signal yc representative of the com- washout circuit 450, to remove undesirable steady state manded flight path angle. A pilot control column input 50 signal components, then amplified by a factor KQ in thus results in a rate of change of the signal yc which is circuit 460 before being summed in summer 470 with proportional to the column position and inversely pro- the gain weighted and filtered column signal COLP.
portional to speed. The inverse speed relationship is This COLP signal effectively provides the pitch rate provided to achieve proper column sensitivity over the command for direct and smooth initiation of the tran- entire aircraft speed range. It is desirable to maintain 55 sient maneuver for a column input. It is derived from constant normal acceleration response capability, re- the speed programmed DCOLV signal to maintain gardless of speed. Normal acceleration is the product of coordination of the initial and steady state pitch rate for the speed and the rate of change of flight path angle. It a given column input, for the entire speed range. The follows then that in order to maintain the nornal accel- jc/8cor is inversely proportional to VGS and the steady eration authority constant, the column input signal must 60 pitch rate must be equal to the rate of change of y. Thus vary inversely proportional with speed. the pitch rate command per unit column must also be The purpose of the lag circuit 250 will be explained in inversely proportional to VGS. The COLP signal pro- connection with the problem of providing the required cessing contains a small lag provided by circuit 430 for display information to achieve satisfactory handling signal noise suppression and a gain circuit 440 for pro- qualities for the short term pilot control loop. 65 viding the desired signal amplitude. Finally the inner The flight path angle error signal -ye is formed in loop signal output from summer 470 is gain pro- summer 270 by taking the difference between the flight grammed in circuit 480 to provide uniform dynamics in path angle command and the signal GAMMA, repre- the inner loop which is affected by the elevator aerody-
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namic effectiveness. Since the elevator effectiveness Display of the ycsignal further has the advantage that (pitching moment per unit deflection) is a function of the pilot always knows the reference command that the speed CAS the KV signal is programmed in circuit 490 automatic system is controlling to. This helps the pilot to compensate for this speed effect. The output of multi- avoid entering the control loop to correct small flight plier 480 is summed with the outer loop elevator com- 5 path angle perturbations in turbulent conditions, which 400 in summer 500 to form are best left to the automatic system.
mand signal from summer the total elevator command. During pilot evaluations of the system so far de- It should be understood that for proper functioning of scribed, it was found that the lag free yc response ap- the control law all signal gains in each of the signal peared unnatural to the pilots. The yc response would paths must be determined in relationship to the other 10 lead the pitch attitude response (see FIGS. Sa, 86) and gains. In general, the overall gain levels were selected force the pilot to abandon his feel for the dynamics of a to minimize flight path angle response lag, commensu- normal well-responding airplane. Both these problems rate with a high level of response damping and accept- were solved by the addition of a small (-0.2 seconds) able pitch attitude-response. For example, it was found first order lag in the yc signal loop, which is shown as that the flight path angle response lag 765 relative to yc 15 gainflag circuit 250 in FIG. 7 . The responses for the could be reduced to approximately 0.7 sec by increasing system including the yc lag circuit and optimized gains FIG. 8c. Note that the resulting flight path the gain levels to the highest level while still providing are shown in adequate damping. This is shown in FIG. 8u The prob- angle response lag ( T ~ ) is - 1 second, that the initial and lem, however, is that the pitch rate response becomes steady state rate of change of flight path angle are well ierkv and the Ditch attitude exhibits a relatively large 20 coordinated and that the commanded flight Dath anale - .
overshoot of the steady state value. The pitch attituie is captured in a well damped, overshoot-fred way. - response could be smoothed and the overshoot largely Finally, FIGS. loa-c show examples of the responses be avoided by lowering the overall gain levels, see FIG. at various speeds, using identical control column inputs.
8b, <This results however in an unacceptably sluggish Note that the control coordination and response damp- flight path angle response. Therefore, a gain level was 25 ing is maintained throughout the flight envelope and selected that compromises between a small y response that the rate of change of yc decreases with increased lag and a small pitch attitude overshoot, resulting in a speed. Note also that the flight path angle response lag T ~ O ~ sec. stays constant for all speeds. As a result, the airplane's The small flight path angle response lag is desired control characteristics remain virtually the same when using the flight path angle display to close the 30 throughout the flight envelope, while the normal accel- short term pilot control loop and provide satisfactory eration response capability for full control column in- control handling characteristics. (See the discussion puts stays constant.
with respect to FIG. 6) Even the smallest achievable Referring again to FIG. 7 , the remaining portion of response lag ( ~ ~ ~ 0 . 7 ) was found to be too large. For the control law relates to the automatic GO AROUND this reason a yc display was added. It basically responds 35 feature. When the airplane descends to critical altitude lag free. The pilot can use the yc display to close the and all conditions of runway vision and airplane posi- short term control loop and divorce himself from the tion relative to the runway or operation of all critical short term dynamics of the actual flight path angle. systems have not been met, the pilot must make a go FIG. 9 depicts the preferred embodiment of the pilot around. This can be a critical maneuver at low altitude, display used in association with the instant invention. 40 demanding minimum altitude loss after the decision has Here, as in the conventional manner, roll attitude is been made. To assist the pilot in making this maneuver indicated via a roll pointer 600 relative to a suitable roll the automatic go around feature has been designed into attitude scale, indicated generally at 652. this Velocity Vector Control Wheel Steering system.
A landing system glide slope indicator 706 is pro- For this purpose the commanded flight path angle ye is vided with an appropriate scale, indicated generally at 45 quickly but smoothly changed from whatever previous value existed before the activation of the go around h e r indicator 710 is provided logic, to a +2" climb-out value. This is done by taking with an appropriate scale indicated generally at 712. the yc and forming an error signal GAMEG relative to Various pitch lines 720-722 are indicated on the a + 2" bias in summer 800. The error signal GAMEG is screen with a reference airplane symbol 230. 50 appropriately amplified in circuit 810 by a gain factor The principal improvement in the display is that si- KGAE and then fed back through switch 820 to sum- multaneous indications 740, 750 represent both the mer 240 and ultimately to the command integrator 260.
flight path command yc and the actual flight path y~ Activation of the GO AROUND ENGAGE logic respectively. (GAE) results in a temporary closure of switch 820 and The yc symbol has the same wedged shape as the 55 therefore in a quick but smooth synchronization of the actual flight path angle symbol, however, the lines are yc signal with the + 2" climb bias. The airplane responds drawn in lighter and broken up. During tracking condi- with a very positive and well controlled pull-up maneu- tions both symbols overlay. The value of the flight path ver and establishes the 2" climb angle without over- angle is read against the Ditch scale. (The pitch attitude shoot, just as in the case of a pilot controlled maneuver.
is determined 6y the airpiane symbol positibn relative to 60 For this purpose the go around engage error signal GAMEG is fed back such that it not only changes the this scale, however 5" nose up bias is applied to unclut- ter the symbology.) For example, when flying level, the yc value, but also results in a proportional pitch up flight path angle wedges overlay and point at the 0 " command which is summed with the pitch rate signal in (horizon) line. During the column inputs the ye sepa- summer 470.
y symbol in proportion to the amplitude 65 In summary, with the described design features, the rates from the of the column input (and thus yc). When the column control law and display system satisfies all requirements input is returned to neutral the y closes in on yc in a of responsiveness, damping, control sensitivity and flight path angle tracking in turbulence. Tests have quick and well damped way.
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seventh means for processing said ycsignal and said y shown that this system significantly reduces pilot work- signal to produce a signal ys, representative of the load over the system known to the prior art.
flight path angle error signal, according to the While a preferred embodiment of the invention has relationship ye= yc- y ; been described in detail, many modifications and varia-
eighth means for .processing said yE signal, said p
tions thereto are possible, all of which fall within the 5 signal and said 8 signal for developing an elevator true spirit and scope of the invention.
command to control the craft’s dynamics in such a I claim: way as to reduce the ye signal to zero; and 1 . An aircraft control and display system providing computer augmented manual steering capability corn- ninth means for displaying said inertial flight path angle signal y and said flight path angle command prising: control reference means for receiving and processing signal yc to the pilot in proper relationship with the a signal representative of the pilot’s steering input aircraft’s pitch attitude.
to develop a control reference signal which is pro- 4. The vertical flight Path angle steering and display system of claim 3 in which said first means includes: portional to the tirne integral of said pilot$ssteering input; 15 means for amplifying said longitudinal pilot’s control aircraft sensing means for sensing and supplying a input by a gain factor which is inversely propor- signal representative of the controlled state of the tional to speed before integrating said amplified
aircraft: longitudinal . . control input signal to produce said yc
signal.
control command means for processing said control reference signal and said signal representative of 20 5* The flight path and system of claim 3 in which said signal representative of the state Of the aircraft to produce a the pilot’s longitudinal control input of said first means control surface command that automatically con- represents the craft’s control column position.
trols the craft’s dynamics to bring the 6. The system of claim 4, wherein said speed for pro- reference state to track the gramming the integration rate of said pilot’s longitudi- value; and 25 nal control input represents the groundspeed of said display means for displaying said control reference craft.
value and said controlled state of said craft to the 7 . The svstem of claim 4, wherein said lonnitudinal p11ot.
pilot’s conirol input signal which has been amgified by 2. An aircraft longitudinal control and system a gain factor inversely proportional to speed is further providing computer augmented manual flight Path 30 processed through a small lag circuit before being inte- angle steering capability, comprising: grated producing said yc signal to provide the desired first mZtnS for receiving and processing a signal rep- dynamic response characteristic of said yc signal for resentative of the pilot’s steering input to produce a display to the pilot.
flight path angle command signal Y c Which is Pro- 8. The system of claim 7 , wherein the processing of portional to the time integral of said pilot’s steering 35 said eighth consists of a proportiona~signal path input; of said y E signal, a proportional signal path of said signal second means for supplying a signal 7 including a low pass filter, and an integrator signal path of the craft’s inertial flight path angle; having both ye and y signal inputs, a summer circuit for third means for supplying a signal 8 representative Of combining said proportional yE signal path, said propor- the of the craft’s pitch rate; tional i , signal path and the output of said integrator fourth means for Processing said Y c signal, said 7 ’
having both yE and 3 signal inputs, for producing a
signal and said 8 signal to develop a control surface signal representative of the outer loop elevator com- command that automatically controls the craft’s mand.
dynamics to smoothly bring the value of said Y 9. The system of claim 8, further including: signal to track the value of said Y c Signal; and means for processing a signal CAS representative of fifth means for displaying said y signal and said yc the craft’s speed to produce a speed related gain signal permitting the pilot to monitor and control programming factor KV; their values.
means for high pass filtering and gain amplifying said 3. An aircraft vertical flight path angle steering and 8 signal; display system for controlling and monitoring the iner- 50 means for low pass fdtering said pilot control input tial flight path angle y of said craft comprising: signal which includes a gain factor inversely pro- first means for receiving and processing a signal rep- portional to speed to provide a signal COLP repre- resentative of the pilot’s longitudinal control input sentative of the pitch rate command; YC means for co.mbining said high pass filtered and gain to produce a flight path angle command signal which is proportional to the time integral of said 55 amplified 8 signal and said COLP signal to produce longitudinal control input signal; a signal representative of the pitch rate error; second means for supplying a signal y representative means for multiplying said pitch rate error signal with of the craft’s inertial flight path angle; said speed related gain factor KV, producing an third means for supplying a signal 8 representative of inner loop elevator command signal; and the craft’s pitch rate; 60 means for combining said outer loop elevator com- fourth means for supplying a signal representative mand and said inner loop elevator command to of the craft’s vertical acceleration; form a total elevator command signal.
fifth means for supplying a signal representative of 10. The system of claim 9, further including a go the speed VGS of the aircrac; around circuit, for synchronizing the value of said yc sixth means for processing said h signal and said VGS 65 upon activation of a go around discrete logic automati- signal to produce a signal 9 representative of the cally to a preferred go around bias signal, for executing rate of change offlight path angle according to the a go-around maneuver.