CA1235777A - Speed capture in climb for aircraft - Google Patents

Abstract

ABSTRACT OF THE DISCLOSURE

An improved method for the automatic acceleration of an aircraft to a preselected speed is described. At any point in the climb portion of flight, a capature of a computed altitude that is increased at a specified rate is performed to provide optimal acceleration.

Description

~235777 1 BACKGROUND OF _ IN~ENIION
1. Field of the Invention The present lnvention relates yenerally to aircraft automatic flight control systems and more specifically to speed control and acceleration to a higher speed during the climb portion of flight.

2. Description of the Prior Art Most co~nercial transport aircraft, general aviation aircraft and military aircraft are equipped with an automatic flight control system.
Automatic flight control systems generally prov_de the pilot with the capability of altering the flight path of the aircraft to achieve and maintain a desired speed, measured either in knots or Mach number.
During a climb portion of flight, the pilot may elect to increase the speed of the aircraft through a manually-entered speed com~and or may be mandated by Pir Traffic Control (ATC) to accelerate to a specified speed, Soon after liftoff, the pilot will generally increase the speed of the alrplane ln order to retract the flaps. In addition, in the United States, the Federal Aviation Administration (FAA) requires that aircraft speed be no greater than 250 knots indicated airspeed at altitudes less than 10,000 feet~ Ihus, the pilot ls generally required to accelerate the aircraft in successive steps as the aircraft configuration changes and a e~e/erafe minimum altitudes are reached. Further, it is desirable to ~cccrate to the optlmum climb speed as quickly as possible in order to maximize fuel savings.
In addition, a positive rate of climb should be maintained at all -times while accelerating and it is generally an ATC re~uirement that a minimum climb rate of 500 fpm will be maintained, particularly at lcw altitudes.
In the prior art, these accelerations were generally accomplished by the well kncwn method of decreasing the pitch attitude of the aircraft in a manner proportional to the difference between the new speed c nand and the actual speed of the aircraft, or speed error. While this scheme will ~0 accelerate the aircraft to -~he desired Mach or airspeed, there is no assurance this will be done in a timely or optimal manner or that any minimum altitude rate restrict1ons will be met.

~23S7~77 e present invention overccmes the shortcomings of the prlor art by computing an altitude based upon the present climb rate, or altltude rate of the aircraft, and then increasing the altitude at a rate that will assure a 500 fpm rate of cli~b or less if the plane is not capable of accelerating at a cllmb rate of 500 fpm. This altitude is then used to tend to pitch the aircraft down to achieve the s~ecifled rate with the engine(s) at climb thrust until the difference between the desired speed and the actual speed of the aircraft is within a predetermined amount, at which time conventional speed control proportional to speed error is resumed.
SUMMARY OF TffE INVENIION
Ihe present invention provides means for automatic, optimal acceleration of an aircraft to a higher selected or commanded speed during the climb portion of flight by the capture of a computed altitude, based on the actual climb altitude rate of the aircraft, that is increased at a specified rate. Ihe rate of increase is aenerally 500 fpm to provide optimal accele~tion and still meet minim~m altitude rate restrictions but may be less lf the plane is not capable of accelerating at a climb rate of 500 fpm. Means are further provided to automatically command a higher speed, based on pilot-entered or stored data, such that the commanded speed will be achieved when above specified altitudes. Optimal accelerations may thus be achieved either with specified altitude require-ments or in the absence of such requirements.
BRIEF ÆSC~IPTION ~ THE DRAWINGS
Fig. 1 is a graph of altitude versus distance illustrating the cli~b flight path generated by the present invention to accelerate to a cc~manded speed which is significantly higher than the actual speed of the aircraft.
Fig, 2 is a graph of altit~de versus dtstance illustrating the climb fllght path qenerated by the present invention when the commanded speed is only slightly higher than ~he actual speed of the aircraft.

123S~77'7 1 Fig. 3 is a block diagram of the present invention shcwing the calculation of a computed altitude, the increase of the computed altitude at a selected rate and the switching of pitch control between conventional proportional s~eed error and automatic capture of the computed altitude.
Fig. 4 is a logic diagram illustrating the various parameters-used in the controlling of the tran~sfer of speed control from conventional proportional airspeed-on-pitch to the altitude capture acceleration control o~ the present lnvention and vice versa.
DESCRIPTION OF THE PREFERRED EM~ODIMENTS

The present invention is u.seful in any automatlc flight control system or ln any performance management sys~em (P~S) that is fully coupled to an automatic flight control system, and provides apparatu~s for automatically transitioning the aircraft to a higher commanded speed in the climb portion of the flight. In either system, it is often desirable for the pilot to increase the speed of the aircraft to achieve a particular speed in order to retract flaps or to accelerate to 250 knots or the desired climb speed when specified minimum altitudes are reached.
To illustrate the invention, refer to Figure l. Assume the aircraft is cli~bing at an altitude above l0,000 feet and that no acceleration of the aircraft has been commanded. Hence, the aircraft will be climbing at some altitude ra~e H. Now assume that the aircraft is to be accelerated by either a pilot-entered or an automatic speed command which is significantly higher than the present speed of the aircraft an~ that the speed ccmmand occurs at 60. ~ computed altitude 61 is deter~dned by the relationship HS = H + KH (l) Where HS = the computed altitude in feet.
H = the actual altitude of the aircraft in feet.

~235777 1 K = a characterizing parameter that determines the shape of the capture flight path and may he be of the form K = ha where he = altitude error in feet between the selected altitude and the present altltude and ha = rate of climb.
H = the present altitude rate of the aircraft in feet per mtnute.
It w111 be noted that the computed altltude will always be KH feet above the actual altitude of the aircraft according to equatlon (1) above.
~ n increasing altitude 62 which will be used by the autopilot or Performanoe Management System (PMS) is determined by (2) where HRAMp = the computed ramplng altitude in feet.
HS = the initlal computed altitude in feet.
HRAMp = the predete~mined rate of increase in the computed altltude in feet per minute.

t = time in minutes.
to = ti~e at 60 in minutes.
It will be noted that when t = to~ or at 60, the ramping coTputed altitude will equal the initial cc~puted altitude according to equation (2) ~bove.
Ihe automatic flight control system or PMS will then co~mand the elevator or horizontal stabilizer in such a fashion as to alter the flight path of the aircraft along line segment 63 according to the control law:
~ - KH = 0 (3) 123577'7 1 where ~ = altitude error in feet between the computed ramping altitude and the actual altitude.
K = a characterlzing parameter as deflned above.
H = the actual altitude rate in feet per minute.
The aircraft will thus accelerate to the new speed command while maintainlng the computed ramplng altltude untll the actual aircraft speed ls withln some predeter~tned amount of the commanded speed, for example, 0.01 ~ach. This event ls shown as 64. At 64, the altltude capture control law, equation (3) above, ls no longer used and the new speed ls malntalned uslng conventlonal proportlonal control speed along the line segment 65.
Referrlng now to E'lgure 2, assume that the newly ccmmanded speed ls only sllghtly more than the present speed of the alrcraft, but the difference ls greater than some predetermined amount, for example 0.005 Mach. E~rther assume that the new speed command occurs at 67.
~s ln the above example, the present lnventlon will compute a ramplng altltude 71 and the control law expressed by e~uatlon (3) will be invoked. ~s the aircraft beglns the capture of the computed ramplng altltude along line segment 68, lt wlll also begln to accelerate toward the new comnhnded speed. When the actual speed is within some predetermined amount of the commanded speed, represented by 69, the capture of the computed ramping altitude is abandoned and conventional speed control is resumed along line segment 70.
By way of example, assume the aircraft is being flown at 250 knots wlth a climb rate of 3000 feet per minute and has just clinbed through 10,000 feet. Assume that a speed command of 300 knots is entered automatically or is entered manually by the pilot. Assume for slmpllclty that the value of K ls held constant at .25. lhe automatlc command of the higher speed would result ln the computation of an lni-tial capture al-titude by equation (1) above and the result would be :~L23S777 1 (10,000 + 750) or 10,750 feet. Assume that the predetermined rate of change of this altltude is to be 500 feet per minute. According to equation (2), the capture altitude would equal 10,750 feet at the point where the new speed is commanded, would equal 11,250 feet one minute later, 11,750 feet two minutes later and so forth. The control law equatlon (3) would thereafter be used to capture the ramping altitude and the aircraft would begin to accelerate toward the commanded 300 knots indicated airspeed. Once having achieved the co~manded speed within some predetermined amount, the altitude capture control law descrlbed above would no longer be used and speed control would revert to conventional proportional control.
To c~nply with FAA regulatlons and approved flying procedures, the speed and altitude requirements could either be entered by the pilot or stored within the automatic pilot or PMS. As the aircraft clim~s frcm takeoff to cruising altitude, the climb speeds would automatically be commanded at predetermined altitudes and changes in alrplane conflguratlon (i.e., retraction of flaps and landing gear) and the new speed command would be captured in the same fashlon lllustrated by the above ex~nple.
The present invention may be implemented by using conventional analog clrcuitry and computational techniques or by using conventional wholly digital techni~les or by using conventional hybrid digital/analog techniques. To simplif~v the understanding of the invention, it will be explained by using a aenerally analog format as shown in Figure 3, it being understood that the same analog format may also represent, in block diagram form, the program of a programmable digital computer wherein the ~arious analo~ inputs are converted to digital signais for digital processing and the various digltal outputs are converted to analog signals for driving the control surface servomotors and the like.
Referring to Figure 3, assume the aircraft ls climbing at an altitude hiqher than 10,000 feet and that no acceleration of the aircraft 12357~77 1 has been commanded. Switch blade 20 wlll be in the position shcwn, making contact wlth contact 19. A signal proportional to the actual Mach num~er of the aircraft is supplied by conventional air data computer 1 and appears on lead 21 and junction 22. It is compared with a signal proportional to the com~anded Mach number, ~c, which appears on lead 24 and ls a~plied to conventlonal summation device 25. The actual Mach number from junctlon 22 appears on lead 23 where it is also applied to summation device 25. Ihe output, which represents the difference between the commanded and actual Mach numbers, Mach error, appears on lead 26 and is applied to summation device 31. Simultaneously, the actual Mach number from junction 22 is applled to conventional rate generator 27 whose output appears on lead 28 and is a signal proportional to the time rate of change of actual Mach number or Mach rate. Mach rate is multiplied by an appropriate gate G 29 and the result appears on lead 30 whlch is supplied to conventional summation device 31. Ihe output of summation device 31, appearing on lead 32, wlll be the well-kncwn proportlonal plus rate control of Mach number.
Lead 32 supplies limiter 33, whose characteristics apr~ear on the face, with the proportlonal plus rate signal. Limiter 33 is a conventional limiter whose function is to assure the resultant pitch and pitch rate of the aircraft will be maintained within specified limits. ~he output of limiter 33 appears on lead 34 and at s~itch contact 19. Ihis signal is applied to conventional summation device 35 via switch blade 20.
Signals proportional to the pitch angle and pitch rate of the aircraft are supplied to summation devlce 35 vla lead 36 in the conventional manner. The output signal of summation device 35 is applied to serv~motor 39 which, through the mechanical linkage 40, moves the aircraft's elevator or horizontal stabilizer 41. Mechanical linkage 37 supplies summation device 35 with a slgnal proportional to elevator or horizontal stabilizer position so that the signal on lead 38 ls reduced to null ln the steady state condition.

~2357~7 1 Slmultaneous with the actlon aescrlbed above, alr data computer 1 supplies a signal proportional to the altitude rate of the alrcraft, H, on lead 2 to fllter 3. Filter 3 ls a conventional fllter whose purpose ls to ellminate or minimize atmospherlc or electronlc nolse that may be present on the altitude rate signal. The output slgnal of fllter 3 ls supplled to gain block 4 where it is multlplied by a value K which may elther bc a constant or a varlable number. The output slgnal of galn block 4 appears at junctlon 5 and represents the term K~l. One lead from junctlon 5 appears on lead 6 and thence to conventlonal summatlon device 8.
Air data computer 1 also supplles a slgnal proportlonal to the actual altltude of the aircraft H on lead 13 and at junctlon 14. One lead from junction 14, lead 15, supplies the altltude slgnal to summatlon device 8 where lt ls algebraically added to the KH term explalned above. The output of the summation appears on lead 9 and represents the term H + KH.
Lead 9 ls supplied to latch 10 whlch ln the present case is synchornlzed such that lts output on lead 11 is identical with the signal on lead 9.
A signal proportlonal to the well known relationship (Thrust-Drag)/
Welght ls applled to lead 42. ~he term on lead 42 is multiplled by a slgnal proportional to true airspeed provided on lead 43 by air data computer 1.
The resulting term appears on lead 44 and is proportional to the maximum altltude rate that can be achleved by the aircraft. The term on lead 44 is multlplled by gain 45 whlch ls chosen so as to produce a percentage of the maximum altltude rate on lead 46, such as 25%. ~he term on lead 46 ls then llmited to ~e wlthln certain values, such as between 100 fpm and 500 fpm by limiter 47. Ihe resulting llmited term on lead 53 represents the rate of lncrease ln the computed capture altltude, HRAMp. Ramp generator 48 normally produces a delta H value on lead 54 computed as:
Delta HRAMP = ~ at (t - to) (4) where Delta HRAMp = the delta H produced by the ramp generator in feet.

~2357~7 HRAMp = rate of increase in the computed altitude in feet per minute t = time in minutes to = time at start of ramp generation in minutes.
When no acceleration has been commanded, the value of to is continuously set equal to t thus providing synchronization and producing a zero term on lead 49.
The KH term from junction 5 is supplied to conven-tional summation device 12 via lead 7. Actual altitude signal from junction 14 is also supplied to summation device 12 via lead 16. The ramping delta H signal is provided to summation device 12 via lead 49. As can be seen, the output of the sum~
mation device on lead 17 represents the term (KH + H) ~ (KH) - (H) + HRAMp at (t - to) which will be at a null value. The latter term will be at a null value since to is continuously set equal to t.
Now assume that the pilot elects to accelerate the aircraft to a new commanded speed by entering the speed through either an analog dial or a compute display panel. If the dif-ference between the newly commanded speed and the old commandedspeed exceeds a predetermined value, for example, 0.005 Mach, switch blade 20 will be moved to make contact with contact 18 and latch 10 will be activated. The value appearing on lead 9 which as explained before represents the term H + KH, at the instant switch blade 20 moves to contact 18 will be stored and maintained on lead 11 reqardless of subsequent changes in the value appearing on lead 9. The value of the signal on lead 11 therefore represents a computed altitude Hs, The value of to at the instant switch blade 20 moves to contact 18 will be stored providing a time reference point for ramp generator 48.

12357~7 Ramp generator 48 will then produce a ramping delta H on lead 49 according to equation (4~ above. The values appearing on leads 7 and 16 do vary with time and represent KH and H repre-sentatively. Thus, the output on lead 17 represents the solu-tion to the expression:
[HS + ~1 at ~t-t )]-H-KH
This signal is applied to summation device 35 via contact 18 and switch blade 20 and is thence used to manipulate the position of the aircraft's elevator or horizontal stabili-zer in such a fashion as to reduce the value appearing on lead17 to a null value. This action will cause either an asymp-totic or circular flight path towards the computed ramping altitude, HRAMp, depending on whether the value of K is a constant or a variable dependent on the actual altitude rate of the aircraft.
As the aircraft pitch angle is decreased during the capture maneuver described above, the aircraft's speed will increase toward the newly selected value. When the difference between the actual speed of the aircraft and the commanded speed is less than some predetermined value, for example 0.01 Mach, switch blade 20 will make contact with contact 19/
returning pitch control to the proportional plus rate control of speed described above. In addition, latch 10 will return to its synchronization mode, ramp generator 48 will return to its synchronization mode and the output on lead 17 will be a null value.
I-t will be clear that the aircraft may not actually capture the computed altitude if the difference between the newly commanded Mach and the previous Mach command is small.
During the decrease in pitch angle of the capture maneuver, ~23S777 the aircraft may accelerate to within the predetermined value before it has captured the ramping a]titude. In such cases, a slight decrease in the rate of climb would occur as the air-craft accelerates.
Referring to Fig. 4, assume the aircraft is climbing at an actual speed M that is within some predetermined amount of the commanded speed MCMD such as .02 Mach. The condition of being in climb mode will produce a logic 1 at junctionlll and on leads 112 and 113. A logic 1 on lead 113 causes the out-put on lead 107 of conventional logic OR gate 109 to be the same as that of lead 116. Similarly, a logic 1 on lead 112 will cause the -lOa-$Z3~7~7 1 output on lead 106 of conventlonal loglc AND gate 108 to be the same as that of lead 110. In the condition described above, the speed ls within the predetermined amount and thus a loglc 1 is produced on lead 116. A logic 1 on lead 116 produces a logic 1 as output on lead 107 from conventional logic OR gate 109. A logic 1 on lead 107 causes conventional latch 105 to be reset producing a logic O output on lead 104.
A logic 0 on lead 104 will cause switch blade 102 to make contact with 119 and will thus transfer the Mach error or airspeed err.or based autopllot command to the autopllot via lead 103.
Now assume the commanded speed i5 inereased more than a predetermined amount such as .005 Mach, such that a logic 1 is produced on lead 110.
A logic 1 on lead 110 will cause a logic 1 to be output on lead 106 from conventional logic AND gate 108. A logic 1 on lead 106 will cause latch 105 to be set producing a logic 1 as output on lead 104. A logic 1 on lead 104 will cause switch blade 102 to move to make contact with 118.
Switch blade 102 will then transfer the synthetic altitude capture auto-pilot ccmmand present on lead 101 to the autopilot via lead 103. When the actual .speed is again within the predetermined amount of the commanded speed, a logic 1 will again be produced on lead 116 which, as explained above, will callse switch blade 102 to make contact with 119, returntng autopil.ot control to the speed error based command.
If the mode were not climb, a logic O would be produced at junction 11 and on leads 112 and 113. A logic 0 on lead 112 will cause the output on lead 106 of conventional logie AND gate 108 to always be zero regardless of the logic state on lead 110. A zero on lead 106 will cause the con-ventional latch 105 to not be set. A logic O on lead 113 will cause the output on lead 107 of conventional logic OR gate 109 to always be a logic 1 regardless of the logic state of lead 116. A logic 1 on lead 107 will cause conventional latch 105 to be reset producing a logic 0 on lead 104 as the output of latch 105. As explained above, this will cau~se switch blade 102 to ma~e contact with 119, returning autopilot control to the ~speed error ba.sed co~and.

~Z35777 1 From the foregoing, lt will be appreciated that the present invention provides improved automatic acceleration of the aircraft in the cli~b portion of the flight in the follcwing manner:
1. 'rhe aircraft is controlled by the computation and capture of a ram~ing altitude in such a way as to provide the optimal acceleration to a ccmmanded speed.
2. 'rhe acceleration described above ls consistent with altitude rate restrictions imposed on the acceleration.

3. 'Ihe aircraft will always maintain a positive rate of cli~b while accelerating.

Claims (6)

THE EMBODIMENTS OF THE INVENTION IN WHICH AN EXCLUSIVE
PROPERTY OR PRIVILEGE IS CLAIMED ARE DEFINED AS FOLLOWS:

1. Acceleration control apparatus for an aircraft flight control system for automatic transition of said aircraft to a selected higher speed including servo means for controlling aircraft pitch attitude comprising: first means for providing a first signal representative of proportional plus rate control of said aircraft speed, said first means being coupled to receive signals representative of actual aircraft speed and said selected higher speed, second means for providing a second signal representative of a computed altitude beginning at a predetermined point above said aircraft's actual altitude, said computed altitude being increased at a selected rate to main-tain said aircraft in a predetermined rate of climb, said second means being coupled to receive signals representative of said aircraft's actual altitude, rate of climb, true airspeed and thrust minus drag divided by weight, and means for selecting said first or second signal for coupling to said servo means.

2. Acceleration control apparatus according to claim 1 wherein said first means for providing said first signal in-cludes means responsive to signals proportional to said actual aircraft speed and said selected higher speed for providing a first error signal proportional to the difference between said actual aircraft speed and said selected higher speed, means responsive to signals proportional to said actual aircraft speed, for providing a time rate of change signal proportional to said actual aircraft speed, means responsive to said time rate of change signal and said first error signal for providong a second error sig-nal proportional to the difference between said first error sig-nal and said time rate of change signal, and means responsive to said second error signal for limiting said second error signal within a predetermined range, and said limited second error signal represents said first signal for coupling to said servo means.

3. Acceleration control apparatus according to claim 1 wherein said second means for providing said second signal includes means responsive to a first algebraic sum of signals proportional to actual altitude and rate of climb of said air-craft for providing a latched signal, means for providing a ramp signal proportional to a signal representative a maximum rate of climb achievable for said aircraft wherein said maximum achievable rate of climb signal is limited to a predetermined range, multiplication means responsive to signals proportional to said aircraft's true air speed and thrust minus drag divided by weight for providing said signal representative of maximum achievable rate of climb, and means for providing a second algebraic sum of signals comprising said ramp signal, said latched signal, and said signals proportional to actual altitude and climb rate of said aircraft, said second algebraic sum being representative of said second signal.

4. Acceleration control apparatus as recited in claim 3 wherein said latched signal is of the form HS = H + K?
where HS = computed altitude in feet H = actual altitude of the aircraft in feet K = a parameter that determines the shape of the flight path and may be of the form he where he = altitude error in feet between the selected altitude and the present altitude and ?a = rate of climb ? = present rate of climb of the aircraft in feet per minute.

5. Acceleration control apparatus as recited in claim 3 wherein said ramp signal is of the form HRAMP = HS + HRAMP at (t - t0) where HRAMP = computed ramping altitude in feet HS = initial computed altitude in feet HRAMP = predetermined rate of increase in the com-puted altitude in feet per minute t = time in minutes t0 = time at initial computed altitude.

-14a-

6. Acceleration control apparatus as recited in claim 3 wherein said second algebraic sum signal is of the form HE - KH = O
where HE = altitude error in feet between the computed ramping altitude and the actual altitude K = a parameter that determines the shape of the flight path and may be of the form where he = altitude error in feet between the selected altitude and the present altitude and ha = rate of climb H = said actual rate of climb in feet per minute.