Conformal Head-Up Display

1. Field of the Invention

This invention relates to aircraft control systems and, more particularly, to systems which present to the pilot position-controlled, generated image, cue indications on a head-up display (HUD) superimposed with visible external cues.

2. Description of the Prior Art

A head-up display (HUD) is one in which a generated image is projected onto a transparent screen in the pilot's normal line of sight so that the pilot can observe the projected image while continuing his normal observation of outside cues.

Prior HUD systems have been developed and are well known, such as a lead-computing gunsight for military aircraft and approach-to-landing path displays. The latter type of HUD designs have made use of a selectable fixed approach path depression angle and/or the instantaneous descent angle of the aircraft (velocity vector) either separately or combined in one display symbol.

One HUD system, of particular interest for commercial aircraft, is the system disclosed in my prior U.S. Pat. No. 4,104,612 entitled HEAD-UP DISPLAY COMMAND BAR GENERATOR. In that disclosure, a feedback-compensated control system makes use of the fixed depression angle combined with computed rate information to derive the HUD command signal. The pilot, by maneuvering to keep the symbol on an aimpoint (the desired touchdown zone), closes the loop of a feedback-compensated control system. The system is referred to herein as the "Compensated Control" HUD.

Arrangements in accordance with the present invention permit the pilot to fly the aircraft head-up during approach and landing for virtually all conditions which may be encountered by the aircraft. In accordance with one aspect of the invention, the HUD is arranged so that the pilot is able to use the display in the same manner in both IFR (Instrument Flight Rules) and VFR (Visual Flight Rules) conditions, thus avoiding a shift from one type of cue to another as the pilot proceeds from IFR conditions to VFR conditions. The system also provides a smooth transition from an instrumented approach to a mode independent of the ILS (Instrument Landing System) and then to a flare mode. Particular arrangements in accordance with the present invention provide magnification of the vertical control cue of the Compensated Control HUD in accordance with glideslope and glideslope rate information, derived from an ILS beam, during approach for landing under such control.

The Compensated Control HUD does not require ground-base instrumentation to provide guidance to a runway that is in sight. An arrangement in accordance with the present invention provides lateral guidance that is cued from a visible runway and requires no ground-based instrumentation. In this arrangement the aircraft symbol display is controlled to remain astride the runway which is being approached, even though the aircraft may be crabbed into a crosswind.

One particular arrangement in accordance with the present invention serves to control the aircraft (at the pilot's option) at a selected fixed altitude prior to pitching over to proceed on the selected descent path. In this arrangement, the location of a reference aim symbol relative to the aircraft symbol in the display provides an altitude hold command. The pilot by maneuvering his aircraft to hold the HUD aircraft symbol on the reference symbol will cause the aircraft to maintain the selected altitude. When the aircraft symbol intercepts the runway aimpoint (visible from the ground through the display), the pilot knows he is at the point where he should begin his selected descent path and he then pitches over to proceed toward the runway aimpoint.

A better understanding of the present invention may be had from a consideration of the following detailed description, taken in conjunction with the accompanying drawings in which:

FIG. 1 is an illustration of one particular HUD display in accordance with the invention, showing runway parameters and other generated flight cues projected on the display;

FIG. 2 is a block diagram of circuitry for incorporating glideslope error signals to amplify the aircraft symbol vertical deflection signal in an ILS approach;

FIG. 3 is an illustration of the HUD parameters related to an instrumented approach;

FIGS. 4A and 4B illustrate the geometric parameters required in the computation of the display parameters of FIG. 3;

FIG. 5 is a block diagram illustrating circuitry for determining lateral deflection of the aircraft symbol as shown in FIG. 3 in accordance with the geometric parameters of FIGS. 4A and 4B;

FIGS. 6A and 6B illustrate in the display and the vertical plane the parameters related to the preflare and flare modes;

FIG. 7 is a block diagram of circuitry for controlling the aircraft symbol vertical deflection in the final phases of landing approach;

FIG. 8 is a block diagram of circuitry for controlling the aim symbol vertical deflection in the final phases of landing approach;

FIG. 9 is a block diagram illustrating particular filter circuitry to provide lateral command with lead for a non-instrumented approach;

FIGS. 10A and 10B illustrate aircraft locations and corresponding displays during altitude hold and transition to the approach;

FIG. 11 is a block diagram of circuitry for controlling vertical deflection of the aim symbol in the altitude hold mode;

FIG. 12 is a block diagram illustrating particular lead circuitry for controlling vertical deflection of the aircraft symbol to provide altitude hold damping and transference to the descent path; and

FIG. 13 is a schematic diagram illustrating a system for generating a head-up display in accordance with the present invention.

The Compensated Control HUD of my prior patent provided vertical guidance for the pilot during descent prior to landing. Preferred embodiments of the present invention utilize my patented Compensated Control HUD and in addition provide HUD lateral and vertical control guidance in a manner which minimizes the transition from IFR to VFR conditions. It also provides command guidance from the minimum ILS control altitude to flare, during a non-instrumented approach and during altitude hold including a smooth transition to a descent path. The ILS is referenced herein, but the invention is not restricted to any particular radio beam landing system.

My Compensated Control HUD is disclosed in detail in my prior U.S. Pat. No. 4,104,612, incorporated by reference herein, and the reader is referred to that patent for a more detailed discussion of that system and a showing of the image generator and projection arrangement which develops the HUD. Summarizing very briefly from that patent, a selected glidepath angle plus aircraft motion terms like pitch, altitude rate and normal acceleration are coupled into the control law circuit module, which generates the deflection signal. This deflection signal is coupled into a symbol generator which generates the image of a command bar (aircraft symbol) on a cathode ray tube (CRT). The deflection signal controls the vertical location of the command bar. The CRT image is projected onto a collimating lens and to a combiner, which is a non-interfering transparent glass located between the pilot's eyes and his forward field of view. The command bar appears to the pilot as superimposed on his view of the runway during a landing approach. The collimated command bar constitutes an aimsight. In steady state, a deflection signal transmitted from the control law module represents the selected glidepath angle, causing the aimsight (collimated command bar) to be held at the selected angle relative to the horizontal. If the aircraft is on the selected glidepath, the pilot will see the bar superimposed on his runway aimpoint. Should the pilot see the bar above or below his aimpoint, he interprets it as control error and corrects with elevator control. An aircraft maneuver produces changes in the aircraft motion terms, which are coupled into the control law module, which produces a change in bar deflection, thus completing the loop of the Compensated Control HUD system.

One principal embodiment of the present invention is an augmentation to my prior U.S. Pat. No. 4,104,612 which utilizes the glideslope signal of a radio beam landing system such as an instrument landing system (ILS) to provide path error magnification. Though my prior patented system is capable of providing vertical guidance in a VFR, non-instrumented approach to the accuracy required of an autopilot by FAA Circular AC 120-29 for an IFR Category II approach, the addition of the ILS augmentation of this embodiment will command closer tracking. According to my prior patent the pilot is guided by the vertical motion of a symbol, herein called aircraft symbol 10 in FIG. 1, relative to a point on the ground, herein called aimpoint 18 in FIG. 1. The vertical location of the aircraft symbol relative to the aimpoint is determined by the compensated control command signal. Where an ILS glideslope beam is available, the aimpoint can be supplemented with a HUD symbol, herein called aim symbol 14 in FIG. 1. Also with the aid of the glideslope error signal the aircraft symbol vertical command signal is amplified. This amplification increases the separation between the aim symbol 14 and the aircraft symbol 10 for a given separation of the aircraft from the glidepath. By increasing the visibility of the error the pilot may fly a tighter control during this instrumented approach.

In this embodiment vertical position of the aircraft symbol in the display is determined by the Compensated Control law of my prior patent referenced above plus the glideslope signal augmentation. A vertical signal of the prior patent (.eta..sub.c -.theta.+.eta..sub.ld) is combined with ILS glideslope error and glideslope error rate as shown in FIG. 2. FIG. 2 shows a Compensated Control circuit 13 from my prior patent combined with glideslope error proportional gain 15 and glideslope error rate derived from the rate taker 17. The rate taker time constant .tau..sub.G should be fast relative to aircraft pitch dynamics. Gains K.sub.G and K.sub.G are adjusted for optimum control. The outputs from stages 15 and 17 are added to the Compensated Control output in a summing stage 19. When no ILS signal is available, the constants K.sub.G and K.sub.G are set to zero.

Another aspect of the invention relates to a method of combining the lateral flight director, vertical guidance, and a perspective display of ground cues so as to provide an instrumented (radio beam) guidance that is conformal with the real world visual cues. The pilot may "fly" the aircraft symbol, which pitches, rolls and yaws relative to the real world, to either a simulated runway aimpoint (a point at the center of the runway where the glideslope intersects the runway) while flying during restricted visibility or the actual runway when it becomes visible.

FIG. 1 shows such a HUD comprising an aircraft symbol 10, a perspective line 12 that superimposes on the actual runway centerline when it is visible, and an aim symbol 14 that appears like a line on the ground that crosses the runway 30, 32, 34 parallel to the threshold 34 at the aimpoint, which is located at the glideslope antenna. The runway aimpoint 18 is indicated by the intersection of the perspective centerline 12 and the aim symbol 14. (The aimpoint can be simulated by other symbologies; for example a perspective display of the total runway 30, 32, 34 with aimpoint, or a small circle located at the aimpoint.) The horizon line 24 has displayed on it a line 50 indicating the aircraft heading and a "V" symbol 40 denoting the runway bearing, a value which is set into the equipment by the pilot in preparation for the approach. Thus, the symbol 40 relative to symbol 50 represents the runway bearing relative to aircraft heading, or Course Error. The aircraft symbol 10 symbolizes an aft view of an aircraft controlled by the pilot in pitch, roll and yaw to correspond to motions of his aircraft. These motions are made evident by a pair of lines or "wings" 16 emanating from a central point, which can be indicated by the circle as shown. During the approach the pilot "flies" the aircraft symbol to keep it centered on the aimpoint 18. The particular symbology described herein is only one example for picturing the space location of aircraft and ground cues in the display. A portion of the embodiment disclosed encompasses the generation of electronics that determine this space location in the conformal display.

FIG. 3 illustrates parameters required in the generation of the conformal display for an instrumented approach. The vertical deflection of the aircraft symbol 10 relative to the horizon 24 is determined by the control law of my prior patent (Compensated Control law) and equals .eta..sub.c +.eta..sub.ld where .eta..sub.c equals the desired angle of descent and n.sub.ld is the feedback compensation lead term. The Compensated Control circuit is designed so that if the pilot maneuvers his aircraft in the vertical plane to hold the aircraft symbol on some line fixed relative to the ground, the aircraft will fly a path that descends at .eta..sub.c degrees and terminates at the fixed point. The vertical deflection of the aircraft symbol relative to the runway aimpoint is .eta..sub.e. If the aircraft descent angle is controlled to hold the center of the aircraft symbol on the aimpoint 18, the command is satisfied and the aircraft will smoothly acquire and track the desired descent path. Aim symbol 14 is computed to stay fixed in the direction of the glideslope antenna on the ground. Therefore, by holding the aircraft symbol 10 on line 14, the pilot will follow an .eta..sub.c descent path that terminates at the glideslope antenna.

The lateral deflection of the aircraft symbol 10 relative to the runway aimpoint 18 is generated from a signal that is like a lateral flight director roll command signal. It contains raw localizer signals plus lateral lead inputs. If the pilot maneuvers his aircraft laterally to hold the center of the aircraft symbol 10 on the aimpoint 18, the flight director command is satisfied and the aircraft will smoothly acquire and track the localizer beam.

The aim symbol 14 is located in the display at the runway aimpoint by depressing the symbol below the horizontal by an angle equal to the depression angle of the aimpoint located at the glideslope transmitter antenna .eta..sub.AIM (FIG. 4A). This is computed from the glideslope angle .eta..sub.c plus the vertical angle deviation of the aircraft from the glideslope .eta..sub.GS as measured by the glideslope error signal at the aircraft.

Referring to FIG. 3, .lambda..sub.ca, the lateral deflection of the aimpoint 18 relative to the display center (aircraft heading) 50, is equal to .lambda..sub.c, the horizontal angle between aircraft heading and selected runway bearing (Course Error), plus .lambda..sub.a, the horizontal angle between runway bearing and aimpoint.

.lambda..sub.a is the product of the depression angle of the aimpoint .eta..sub.AIM and the ratio of the pilot's eye height (HEYE, FIG. 4A) to the lateral distance his eye position is offset from an extension of the runway centerline (OSANT, FIG. 4B).

The perspective line 12 is drawn from the selected runway bearing 40 on the horizon to the aimpoint 18.

The following geometric computations are necessary to the solution of this embodiment. Referring to FIG. 4B, the distance the aircraft localizer receiver antenna is offset from the runway centerline OS is computed from the angle derived from the localizer error signal .tau..sub.LOC and the distance from the localizer transmitter R.sub.LOC. HEYE and OSANT are corrected for eye location relative to radio altimeter and aircraft localizer receiver antenna respectively. Finally the distance to the localizer the distance to the localizer transmitter is computed as the distance to the glideslope transmitter R.sub.AIM, which is a function of altitude and glideslope depression angle (described above), plus an approximate distance between glideslope transmitter and localizer transmitter R.sub.GL.

At R.sub.AIM distances that are less than the minimum limit of glideslope reliability R.sub.AIMIN, R.sub.LOC becomes a function of airspeed TAS and time t.

Runway edgelines, which may be added to the display, are computed in the same manner as the centerline by adding and subtracting to OSANT the runway half width. The depression angle of the threshold .eta..sub.THR and runway far end can be computed as the arctangent of (HEYE/horizontal distance from aircraft to the respective runway ends). The horizontal distances are computed by adding or subtracting the distance between glideslope transmitter and runway ends.

FIG. 5 is a schematic representation in block diagram form of circuitry that generates the signal .lambda..sub.ca, which locates in the display the perspective direction of the aimpoint as seen by the pilot, as described hereinabove by reference to FIGS. 3 and 4A-4B. As is shown in FIG. 5, signals indicating pitch angle and altitude are applied to a stage 100 which corrects altimeter reading to eye height, which is a function of pitch angle, and provides a signal HEYE to divider stages 102 and 104. In the divider stage 102, HEYE is divided by a signal derived from the summation of .eta..sub.c and .eta..sub.GS (see FIG. 4A). The output of divider 102, R.sub.AIM, is added to the signal R.sub.GL to develop the signal R.sub.LOC which is applied to one pole 106 of a switch 108. The armature 110 of the switch 108 is connected to a multiplier stage 112, the other input to which is the localizer error signal .tau..sub.LOC. The output of the multiplier 112 is the signal OS which is applied to a correction stage 114 for combination with a signal .lambda..sub.c which is the difference between aircraft heading and selected runway bearing. The correction stage 114 provides the correction for lateral offset between the aircraft's localizer receiver antenna and the pilot's eyes when the aircraft is not on heading, and provides an output signal OSANT which is applied to the divider 104. The output of divider 104, OSANT/HEYE, is multiplied by .eta..sub.AIM in a multiplier stage 116 and the product, .lambda..sub.a, is summed with the horizontal angle .lambda..sub.c to provide the signal .lambda..sub.ca.

When the distance to the glideslope transmitter, R.sub.AIM, becomes less than the minimum limit of glideslope reliability, R.sub.AIMIN, the armature 110 of switch 108 transfers to the pole 107 and R.sub.LOC is generated as a function of time. Pole 107 receives a signal which is the difference between R.sub.AMIN +R.sub.GL and the integration (provided in integrator stage 118) of true air speed. Under these conditions, the signal at pole 107 is applied through switch 108 via armature 110 and is processed as previously described in place of the previous R.sub.LOC signal.

Where it is considered unsafe to continue to follow the ILS glideslope below a minimum height, the pilot must either discontinue the landing or be dependent upon runway and terrain visual cues. One preferred embodiment allows a continuation of HUD guidance below this minimum height by transitioning from an instrumented approach to a mode independent of the ILS and then to a flare mode. This embodiment also provides smooth transitioning from one mode into the next.

A landing approach can be continued below ILS usability, with ground cues not visible, by maintaining the same angle of descent .eta..sub.c. This is accomplished by converting the aim symbol 14 so that it will indicate the direction that it is desired to descend and converting the aircraft symbol 10 so that it will indicate the direction the aircraft is descending at any moment. The aim symbol, fixed relative to the ground, will become the command aim symbol 64 (FIG. 6A), which is depressed below the horizontal by the fixed angle .eta..sub.c ; and the aircraft symbol will be depressed below the horizontal by the velocity vector .gamma.. Thus by maneuvering to hold the aircraft symbol 10 on the aim symbol 64, the aircraft will be flown in the direction .eta..sub.c. In this mode the aim symbol at .eta..sub.a equals .eta..sub.c and the aircraft symbol at .eta..sub.syn equals .gamma., the velocity vector.

The transference from the ILS dependent mode to the constant direction mode should be done at the lowest altitude that the ILS glideslope is satisfactory (designated "transition height"), because the constant direction is maintained open loop and is dependent upon the accuracy of obtaining .gamma., the velocity vector. In the absence of an Inertial Navigation System, .gamma. is dependent upon air measurements which are affected by wind.

FIGS. 6A and 6B illustrate particular parameters used in preflare and flare modes. In the flare mode aim symbol 64 is converted to command a programmed sink rate reduction, while the aircraft symbol 10 is converted to indicate a sink rate error relative to the aim symbol. Thus in this mode the aim symbol 64 at .eta..sub.a becomes sink rate command and the aircraft symbol at .eta..sub.sym indicates sink rate. The flare embodiment of this invention, though described as associated with an instrumented IFR approach, operates the same in a non-instrumented approach under VFR conditions.

An instantaneous change in mode at a selected altitude would produce a step in the input and a jump in the symbology. Circuitry for avoiding this is depicted in the block diagrams of FIGS. 7 and 8. The circuit of FIG. 7 controls the vertical deflection of the aircraft symbol 10 while the circuit of FIG. 8 controls the vertical deflection of the aim symbol 64.

FIG. 7 depicts the circuitry that transfers aircraft symbol control from Compensated Control to velocity vector by linearly reducing the gain of the Compensated Control input from one at transition height TH to zero at flare height FH, while the gain of the velocity vector input is linearly increased from zero to one. This is depicted respectively at 150 and 152. Then between flare height and touchdown, aircraft symbol control is similarly transferred from velocity vector to altitude rate feedback as shown by 154 and 156. The switching is indicated conceptually by the switch 160. Thus, the switch 160 may be considered to be at the position A for altitudes above TH (transition height), at the position D between transition height and flare, and at the position F for altitudes below the flare height.

FIG. 8 depicts the aim symbol position control that corresponds to the aircraft symbol modes described. Above TH the aim symbol designates the direction of the aimpoint (or the glideslope antenna). At TH the gain of the ILS glideslope error signal input to the aim symbol begins to reduce linearly, reaching zero at flare height as indicated by 162, so that at flare height the aim symbol depression angle equals .eta..sub.c. Then between flare height and touchdown the aim symbol depression angle is reduced to an angle consistent with the desired touchdown descent rate. This is depicted at 164 and 166. Mode switch 170 in FIG. 8 corresponds to 160 in FIG. 7.

Note that the switching altitudes described need not be restrictive. For instance in mode D, designated as between TH and flare, the gain change could have been completed above flare and constant gain maintained until flare height is reached. Also the gain changes need not be linear.

In FIG. 7 the velocity vector input to 152 and 154 is shown computed as vertical speed h divided by horizontal speed V. Horizontal speed may be obtained from airspeed (TAS) or ground speed where available. Lead inputs .theta. and A.sub.n are combined with the velocity vector to make it flyable.

Flare is controlled according to aircraft altitude and descent rate. The aim symbol in FIG. 8 converts from a descent angle reference .eta..sub.c to a descent rate reference TD h CMD, while the aircraft symbol in FIG. 7 converts from a measured descent angle to a measured descent rate. Thus the vertical position of the aircraft symbol relative to the aim symbol cues the pilot that he is above or below a programmed reduction in descent rate as he approaches touchdown.

A further embodiment of the invention provides lateral guidance without the aid of ground-based instrumentation such as the ILS. This, combined with my prior U.S. Pat. No. 4,104,612, which has been referred to herein as the Compensated Control HUD, makes it possible to "fly" an aircraft symbol both laterally and vertically to an aimpoint on a visible runway where no ILS facilities are available. The lateral guidance aids the pilot in acquiring and holding a path that is colinear (lines up) with the runway centerline. Lead information is incorporated in the lateral deflection of the aircraft symbol to aid the pilot in anticipating overshoot. Additionally, because the aircraft symbol is held astride the runway, vertical control to the runway aimpoint is made easier.

FIG. 9 shows the generation of the signal controlling the lateral deflection .lambda..sub.VFR of the aircraft symbol relative to aircraft heading (center of the display). In the short term the aircraft symbol moves laterally with aircraft heading. In the steady state, however, the lateral position is washed out to selected runway heading. In steady state the aircraft symbol lateral position is the angular distance of runway bearing (selected heading) from aircraft heading; this equals the Course Error input. A change in aircraft heading is transmitted through the wash-out filter 182, 184 to cause, in the short term, the aircraft symbol to hold its position relative to the display center, since equal and opposite signals are summed at 180. At the same time the real world runway moves in the direction opposite to the aircraft's turn (thus providing lead relative to the real world). In the long term the signal that is transmitted through the wash-out filter circuitry reduces to zero, so that again the aircraft symbol is deflected from the center of the display by the Course Error, and thus it remains in the direction of the runway bearing. The lateral deflection of the aircraft symbol relative to aircraft heading is ##EQU1## The time constant .tau..sub..PSI. of the wash-out circuit should be long relative to aircraft lateral dynamics (10 seconds for large aircraft). This design, as indicated by FIG. 9, produces lateral lead in a turn by the motion of the aircraft symbol relative to the ground. The lead input to the lateral guidance may be obtained by various techniques. Since the purpose of the wash-out circuit is to provide a deflection of the aircraft symbol in the short term for changes in aircraft heading, aircraft heading may be substituted for the Course Error input to the wash-out circuit. This avoids the transient due to the change in Course Error when the pilot selects a new Selected Heading. ##EQU2## Additional lead may be gained with a roll angle input, with K.sub..PSI. 182 and K.sub..phi. 186 gains adjusted. Also if an Inertial Navigation System is available, the inertial lateral motion signal can provide the best possible control lead.

A further major embodiment comprises an altitude hold mode. In the altitude hold mode, the aim symbol 64 of FIG. 6A provides a command relative to the aircraft symbol 10 that enables the pilot to hold his aircraft at a constant selected altitude. It also enables him to fly a smooth transfer without overshoot to the descent path when the proper point is reached without change in symbology or manner of control. A HUD approach generally involves flying at a constant altitude while watching the HUD aircraft symbol as it moves relative to the ground. When this symbol reaches his runway aimpoint, the pilot knows he is to pitch over and begin his descent along the selected descent path. This situation is represented in FIGS. 10A and 10B which show the altitude hold mode at points a through c. At d, the aircraft symbol 10, located below the horizon by the fixed descent angle .eta..sub.c is superimposed over the aimpoint. Thus at point d, the pilot pitches over and proceeds through the point e along the descent path to touchdown at the aimpoint. In this manner altitude hold guidance and guidance for transitioning to landing approach is provided as the pilot continues to fly head-up (looking through the windshield).

The altitude hold mode comprises the two display symbols shown in FIG. 10B, a guidance symbol, or aircraft symbol 10, and a command symbol, or aim symbol 64. While the aircraft holds the selected altitude the aircraft symbol and aim symbol will be located together at a fixed depression angle .eta..sub.c as shown at points c and d, FIGS. 10A-10B. Should the aircraft be maintaining a constant altitude but below the selected altitude, the aim symbol will appear above the aircraft symbol by the angle .eta..sub.e that is proportional to the altitude error. This is shown at a, FIGS. 10A-10B; this tells the pilot to "fly up". The pilot's task is to maneuver his aircraft to hold the aircraft symbol in line with the aim symbol as at b, c, and d. He "flies" the aircraft symbol to the aim symbol. At b, as the aircraft climbs, the command is satisfied though the selected altitude has not yet been reached. Dynamic lead terms .eta..sub.ld added to the fixed angle .eta..sub.c have deflected the aircraft symbol up. This constitutes feedback damping, which aids the pilot in smoothly acquiring and tracking the selected altitude.

The circuit of FIG. 11 serves to provide a command signal for aim symbol deflection. The selected altitude .eta..sub.SEL is set into the equipment by the pilot and is summed with the radio altimeter reading h in a summing stage 202. The resulting difference is applied through a gain stage 204 which converts an altitude error to an aim symbol deflection angle; the gain is optimized for ease of tracking. When in the altitude hold mode, switch 206 is closed and the negative of the scaled altitude error is applied to a summing stage 203 for combination with the descent path angle .eta..sub.c and the pitch angle .theta.. .theta. provides pitch stabilization (that is, symbology will not pitch with aircraft), a requirement in any HUD system. The signal out of summing stage 203 is applied to the deflection circuit for the aim symbol 64.

The circuit of FIG. 12 serves to provide a command signal for aircraft symbol deflection. The symbol is deflected by dynamic lead signals comprising washed-out pitch, pitch rate, and washed-out altitude rate coupled through optimized gains 209, 211 and 213 combined at summing stage 214. Other lead combinations can be used for altitude hold damping; for example washed-out pitch and pitch rate alone would be adequate. Also any lead .eta..sub.ld disclosed in my prior U.S. Pat. No. 4,104,612 would be satisfactory. With switch 206A in the position shown (Alt. hold/Approach) integrator 212 will wash out the signal at summing stages 207 according to time constant .tau..sub.a..tau..sub.a is made sufficiently long to damp long period altitude transients. Like the aim symbol deflection, the steady state aircraft symbol deflection is .eta..sub.c, which is pitch stabilized by .theta. at summing stage 208. The signal out of 208 is combined with the lead signals at summing stage 214 to produce the signal applied to the deflection circuit for the aircraft symbol.

An important adjunct of the altitude hold function is the transference from the altitude hold mode to the approach mode. FIGS. 10A-10B show the aim symbol 64 and the aircraft symbol 10 maintained in alignment at the fixed deflection angle .eta..sub.c at point c as the aircraft approaches the runway 30, flying at a constant selected altitude. At point d the display shows the aircraft symbol directly in line with the aimpoint on the runway, thus signifying that the aircraft is on the selected glidepath .eta..sub.c where the pilot should be controlling the aircraft to descend. Switch 206 (FIG. 11) should be opened to remove the altitude hold signal from the aim symbol. The pilot may transfer his control reference from the aim symbol to the aimpoint, thus transitioning to the approach mode with negligible eye movement or attention change. By removing the altitude hold error signal from the vertical deflection of the aim symbol, the aim symbol is converted to the fixed depression angle .eta..sub.c. The aim symbol may or may not be removed from the display. In IFR conditions, the aim symbol can be converted to designate the runway aimpoint 34.

Capture of the glidepath without overshoot may be made easier by reducing the time constants of the pitch and altitude rate wash-outs during the period of capture as indicated in FIG. 12. At the instant of pitch-over, 206A is switched to the Capture mode so that the integrator gains will be increased to K.sub.F 210 and the time constant of the wash-outs will be reduced by the factor 1/K.sub.F. This allows the pitch and altitude rate steady state reference to change more quickly from the values acquired in altitude hold to the glideslope values, thus minimizing the potential for overshoot. When the glidepath is established, 206A is switched back to Alt. Hold/Approach so that the time constants will revert to the longer period to damp long period altitude rate transients during the approach. Further details as to variations possible for aircraft symbol (command bar) lead circuitry .eta..sub.ld are described in my prior U.S. Pat. No. 4,104,612 and incorporated herein.

FIG. 13 depicts in block diagram form an overall system in accordance with the present invention. As shown in FIG. 13, a cathode ray tube 220 is used to develop a visual image which is applied through a collimating lens 222 to a combiner 224 which presents the display seen by the pilot. The combiner is a non-interfering transparent glass located between the pilot's eyes and his forward field-of-view. The image of the cathode ray tube 220 is developed by symbol generators 226 which generate the various symbols to be displayed in response to inputs from the control circuitry 230 as described hereinabove. Thus, the pilot is enabled to close the loop between the system signal outputs generated by circuitry 230 and his control of the aircraft by "flying" the aircraft symbol to coincide with the aim symbol. Additional symbology provide reference information to the pilot. Use of the system in this fashion substantially eases the pilot's task, particularly during landing approaches under adverse weather conditions, and may permit landing of the aircraft under IFR conditions.

Although there have been described above specific arrangements of a conformal head-up display in accordance with the invention for the purpose of illustrating the manner in which the invention may be used to advantage, it will be appreciated that the invention is not limited thereto. Accordingly, any and all modifications, variations or equivalent arrangements which may occur to those skilled in the art should be considered to be within the scope of the invention as defined by the annexed claims.