STABILIZATION CONTROL CIRCUIT FOR VERTICAL POSITION IN AN INERTIAL NAVIGATOR

Abstract: An inertial control system for providing vertical velocity and vertical position information for an aircraft is described. Vertical velocity is determined as a function of vertical acceleration of the aircraft and a correction signal. The vertical acceleration correction signal is derived from a function of the barometric altitude and an altitude error signal. The altitude error signal is derived from the difference between the inertially determined vertical position and the barometric altitude. The vertical acceleration correction signal is derived from the altitude error signal and its integral multiplied by variable values which are a function of the time differentiation of the barometric altitude. Vertical position of the aircraft is determined as a function of the derived vertical velocity corrected by a vertical velocity correction signal.

Description: STABILIZATION CONTROL CIRCUIT FOR VERTICAL POSITION IN AN INERTIAL NAVIGATOR The present invention generally relates to inertial navigational systems. More specifically, the invention relates to a control system circuit for providing output signals representative of vertical velocity and vertical position of an aircraft. BACKGROUND OF THE INVENTION Inertial navigational systems for aircraft usually employ a triad of accelero eters in combinatipn with a triad of angular rate sensors for providing velocity and position information to a navigation computer. Specifically, .from continuous knowledge of the vertical acceleration, i.e., perpendicular to the earth's surface, one can determine the vertical velocity and vertical position by time integration of the vertical acceleration as is well known. Simple integration of vertical acceleration to provide vertical velocity, and integration of the vertical velocity to provide vertical position will result in the vertical axis information of an inertial navigator to have unbounded error do to two causes. First, error in the measurement of vertical acceleration (vertical acceleration being determined as a function of the three accelerometers) is directly integrated to cause a vertical velocity error and a vertical position error. Second, the erroneous vertical position error results in an error in the computer derived vertical-gravity component which, in turn, causes an even faster growth in vertical position error in the overall inertial navigation system. To avoid the altitude diversion problem as just presented , it is common practice to blend the inertial altitude measurement from the primary control system, i.e. accelerometers and gyros, with the altitude derived from a barometric altimeter. The barometric altimeter being a well known, device for providing altitude information as a direct function of the jvalue of the barometric pressure. In the prior art, the barometric altitude and the vertical position signal from the inertial system are compared to generate an altitude error signal. In turn, an acceleration correction or stabilization signal is derived as a function of the altitude error signal and used for modification of the inertial system vertical acceleration output before integration to determine vertical velocity. Further, a velocity correction or stabilization signal as a function of the altitude error signal is combined with the vertical velocity signal before integration to obtain the inertial vertical position. Although prior art control loops, constructed in a manner as just indicated, have been successful, they lack accuracy needed in some applications where the aircraft has to perform accurate flight patterns including rapid dive and climb operations. BRIEF DESCRIPTION OF THE INVENTION An object of the present invention is to provide a stabilization control scheme for providing accurate vertical velocity and vertical position information for an aircraft in the presence of rapid dive and climb rates of the aircraft. The present invention provides an inertial control system for providing vertical velocity and vertical position information for an aircraft. In the pres.ent invention vertical velocity is determined as a function of vertical acceleration of the aircraft and a correction signal. The vertical acceleration correction signal is a function of the barometric altitude and an altitude error signal. The altitude error signal is the difference between the inertially determined vertical position and the barometric altitude. The vertical acceleration correction signal is the altitude error signal and its integral multiplied by variable values which are a function of the time differentiation of the barometric altitude. The correction signal function includes variable gains determined by time differentiation of the barometric altitude. Vertical position of the aircraft is. determined as a function of the derived vertical velocity corrected by a vertical velocity correction signal. In the present invention, the vertical velocity correction signal is a time independent function of the altitude error signal multiplied by a variable gain which is a function of the time differentiation of the barometric altitude. BRIEF DESCRIPTION OF THE DRAWINGS Figure 1 is a block diagram of a typical inertial navigation control system. Figure 2 is a block diagram of an inertial control system of the prior art for providing vertical 'velocity and vertical position information for an aircraft. Figure 3 is a block diagram of a vertical stabilization control system in accordance with the present invention. DETAILED DESCRIPTION OF THE INVENTION Shown in Figure l is a simplified block diagram of a inertial navigation control system. Only that portion of the inertial navigation system which is related to the present invention is shown. There shown in Figure 1 is an aircraft 10 which includes a triad of accelerometers having outputs A-^, A 2 and A 3 presented as inputs to an inertial navigation system computer 20. Typically, the triad of accelerometers are orthogonally positioned relative to each other in order to provide orthogonal components of acceleration. After initialization of the inertial system computer, the system computer can calculate the orthogonal components of acceleration AX, AY and AZ, where the acceleration AZ is identified as the acceleration in the vertical direction, i.e., perpendicular to the earth's surface. Because gravitation and the earth's radius are part of the systems equation used by the inertial system computer 20, vertical position information is also provided as input to system computer 20. The system of Figure l as just briefly described is well known to those skilled in the art of an inertial navigation control systems. It should be understood that any errors or bias in the acceleration inputs k , A , and A 3 and the vertical position input PZ will particularly impact the value of the vertical acceleration output AZ. Of course, any errors in AZ and PZ will ultimately affect the determination of the vertical position of the aircraft and its flight operations. Figure 2 illustrates a control system for determining the vertical position PZ as a function the inertially sensed vertical acceleration AZ and the barometric altitude signal BZ derived from a barometric altimeter. The vertical velocity output signal VZ is the integration 215 of the signal difference between signals AZ and the acceleration correction signal S7. The vertical position signal PZ, i.e., the altitude, is the integration.225 of the- signal difference between the vertical velocity signal VZ and a velocity correction signal S5. Circuit means 210 shown as a "summing means" performs the signal difference-operation between signals AZ and S7, and circuit means 220 performs the signal difference operation between signals VZ and S5. Circuit means 230 provides an error signal "E" as the difference between the vertical position signal PZ minus the barometric altitude signal BZ. The velocity correction signal S5 is shown as the error signal E multiplied by a fixed gain Gl. The acceleration correction signal S7 is a sum 250 of signals SI and S3. S3 being the error signal E multiplied by a fixed gain G2, and signal SI is the integration 240 of the error signal E multiplied by fixed gain G3. The value of G2 is in the order of twice the gravitational force divided by the earth's radius, and the value of G3 is in the order of 1.0E-06. The integral feedback provided by the circuitry of integrator block 240 and gain block 242 provides a delay time in the order of 50 to 100 seconds. This latter mentioned circuit .provides a feedback dampening loop to filter measurement noise contained in the barometric altitude signal BZ. Operation of the control system of Figure 2 will now be described. Calculation-of- the -vertical position PZ using the acceleration measurement AZ. - derived from inertial sensors is an inherently unstable process. This is so, since PZ is used by computer 20 to calculate the value of gravity. The value of gravity is taken into consideration for the determination of the output AZ as substantially illustrated in Figure 1. In affect, this is positive feedback around a loop from PZ to AZ and back through the two integrators 215 and 225. This feedback is small — approximately two times the value of gravity divided by the earth's radius. However, even though the feedback is small, the loop does diverge causing unacceptably large errors, after several minutes, if undamped. As illustrated in the prior art of Figure 2, the barometric altitude BZ is used for damping. Velocity correction single S5 provides proportional feedback dependent upon the error signal E. Acceleration correction signal S7 provides integral plus proportional feedback, SI and S3 respectably, also dependent upon the error signal E. The purpose of the SI and S3 feedback signals is to zero out any offset error in the vertical acceleration input AZ " . The combination of correction signals S5 and S7 provide control loop stability.- During dives and climbs, the barometric altitude signal BZ can have large errors and thereby introduce large transient system errors in the vertical position PZ, and vertical velocity VZ, and vertical acceleration AZ. In accordance with present invention, the system of Figure 2 may be provided with added stabilization as will be now described with reference to Figure 3. Referring now to Figure 3, there shown is a control system similar to that shown in Figure 2 with similar functioning and arranged components in Figure 3 having the same numeral designations as in the previously described Figure 2. Added to the circuit of Figure 2, as illustrated in Figure 3, is a time differentiation circuit 347 for determining the time differentiation value BZD of BZ with time lag. The output of circuit 347 is presented to a gain control block 380 having output signals for controlling variable gain blocks or multipliers as will be subsequently described. The circuit of Figure 3 includes variable gain blocks 325, 345, and 342 substituted for fixed gain blocks 235, 245, and 242, respectively. Blocks 325, 345, and 342 provide variable gains of Kl, K2^ and K3, respectively. The variable gains Kl, K2, and K3 are under control of gain control block 380. The values of the variable gains Kl, K2, and K3 will be further described. In the present invention as illustrated in Figure 3, the time differentiated output BZD is presented to gain control block 380 for determining values of Kl, K2 and K3 as follows: Kl * = .03 + .1575 * X K2 - 4 Y 2 + 2g/R K3 - 2 Y 3 where X = ____________ BZD 2 + 50 2 Y - .01 * (1 - X) g = earth's gravitational force R = earth's radius BZD * = Filtered derivative barometric altitude In turn, gain control block 380 is coupled to gain blocks 325, 345, and 342 for providing appropriate output signals to variably control the gain values of Kl, K2, and K3, respectively, of gain blocks 335, 345, -and 342. Gain control block 380 may be constructed from a wide variety of analog and digital circuits well known in the art. The signal outputs from gain control block 380 to control variable gain blocks 325, 342, and 345 depend upon the implementation of these blocks. -In the preferred embodiment of the present invention, it is intended that gain control block 380, as well as the variable gain blocks 325, 342, and 345 are all part of the inertial navigation computer. Time differentiation circuit 347 includes means 350 for providing a difference between the input signal BZ and the integral 352 of the time differentiated output BZD. The output of means 350 is presented through gain block 358 having a gain value of G7. The combination of circuit blocks 350, 352, and 358 provide the time differentiated value BZD of the barometric altitude BZ with high frequency noise roll off and a time delay or lag. Further shown in Figure 3 is a time delay circuit 390 between the output signal PZ and summing means 230. Circuit 390 is shown comprised of summing means 290, gain block 292 having gain G6, and integrator 294 having an output signal S6 which is also fed back to summing means 290. Summing means 290 performs the difference function between the vertical altitude PZ and the integrator 294 output signal. Circuit 347 as just described is intended to provide a delay of the presentation of the vertical velocity signal PZ to summing means 230 for determination of the altitude error signal E. The amount of delay is intended to be substantially that delay which is inherent in the barometric altimeter.- The intended delay is comparatively short, and is in the order 3 seconds for values of G6 equal to .333. The control system as -described by Figure 3 will now be described. During dives and climbs there can exist barometric altitude errors which can introduce large transient errors in AZ, VZ, and PZ. The variable loop gains Kl, K2, and K3 take advantage of the short term (5 minutes) stability of the inertial calculation by computer 20 for determining AZ and the fact that the dive and climb durations are short, for example 1 minute. During dives and climbs the rate of change of barometric altitude output BZD is sufficiently large such that the variable gains of K2 and K3 are reduced, and approach zero. This essentially opens the acceleration correction loop and thereby minimizes the effect of barometric altitude signal errors into the input of integrator 215 and the resultant vertical velocity VZ. At the same time, the gain Kl is increased so that the control loop 305 # operates in a manner that PZ more closely tracks signal BZ such that error signal E approaches zero. At the end of the climb or dive, when level flight resumes, the previous gain values of Kl, K2, and K3 are not immediately changed but rather maintain there dive/climb values for a short time after level flight is resumed. This time delay is predominately controlled by the circuit of block 347, and specifically the value of gain G7 of block 358. This short time delay in maintaining the variable gains Kl, • K2, and K3 further minimizes errors. While the invention has ~ been described in its preferred embodiment, it is to be understood that the words which have been used are words of description rather than a limitation and that changes may be made without departing from the true spirit and intent of the present application. Specifically, it should be understood, that the control system described by the block diagram of Figure 3 may be implemented entirely by software in a computer. Referring to Figure 1, it should be recognized by those skilled in the art .that the control system block 30 may be entirely incorporated into the inertial system computer 20. On the • otherhand, it should also be recognized by those skilled in the art that the circuit of Figure 3 may be implemented by a wide variety of analog and digital circuits. Specifically, any of the time delays may be incorporated by a variety of techniques including digital time delay circuits.

Claims: THE CLAIMS 1. An apparatus for providing signals representative of vertical velocity and vertical position of an aircraft comprising: first means for generating a first signal representative of vertical acceleration of said aircraft; second means for generating a second signal representative of the altitude of said aircraft derived from a barometric altimeter; third means for integrating " the sum of said first signal and an acceleration correction signal to provide a vertical velocity output signal representative of the vertical velocity of said aircraft; fourth means for integrating the sum of said vertical velocity output signal and a velocity correction signal to provide a vertical position output signal representative of the vertical position of said aircraft; fifth means for time differentiating said second signal and providing a third signal representative of said differentiation; sixth means for generating an altitude error signal representative of the difference between said vertical position signal and said second signal; seventh means for generating said velocity correction signal as a time independent function of said altitude error signal multiplied by a first variable value which is a function of said third signal, and in which said first variable value increases in value with increasing values of said third signal; and eighth means for generating said acceleration correction signal as a second function of said altitude error signal such that said acceleration correction signal is inversely related to said third signal. 2. The apparatus of claim 1 wherein said sixth means includes means operative to perform said difference of function on the instantaneous value of said second signal and said vertical position signal occurring a selected time before said instantaneous value. 3. The apparatus of claim 1 wherein said acceleration correction signal includes the sum of: (i) said altitude error signal multiplied by second variable value which is a function of said third signal, and in which said second variable value decreases in value with increasing values of said third signal, and (ii) the time integrated value of said altitude error signal multiplied by a . third variable value which is a function of said third signal, and in which said third variable value decreases in value with increasing values of said second signal. 4. The apparatus of claim 3 wherein said first, second, and third variable values are related to said third signal by: First value -» C_ + c 2 * X Second value ■ C 3 * Y 2 + C 4 , and Third value - C= * Y 3 where BZ *■ said second signal BZD ■ » first time derivative of BZ (BZD) 2 BZD" Y =- C 7 [1 - X], and *-"l * * -2' ""3 , ^4' ^5' ***6' anc; -- C « - are selected constant coefficients. 5. The apparatus of claim 1 wherein said fifth means includes means for filtering high frequency noise of said second signal. 6. The apparatus of claim 1 wherein said fifth means includes time delay means for time delaying said third signal response to change in said second signal. 7. The apparatus of claim 1 wherein said fifth means includes: integrating means for integrating said third signal and providing a fourth signal indicative of said integration; summing means for running said second signal and said fourth signal and providing a fifth signal indicative of said sum; and gain means for multiplying said sum signal and providing said third signal.

Inventors: ROSS HUGH C, ROSS, HUGH, C.
Applicants: HONEYWELL INC, HONEYWELL INC.
Application year: 1988
Main classification: G01C21
IPC codes: G01C5, G01C21