CN1230721C

CN1230721C - Aircraft pitch-axis stability and command augmentation system

Info

Publication number: CN1230721C
Application number: CN 96111536
Authority: CN
Inventors: 齐奥玛斯纳玛巴第; 蒙蒂·R·埃文斯; 爱德华·E·科尔曼; 罗伯特·J·布利格; 理查德·S·布鲁豪斯; 多尔·M·安德森; 蒂莫西·A·尼尔森
Original assignee: Boeing Co
Current assignee: Boeing Co
Priority date: 1996-08-22
Filing date: 1996-08-22
Publication date: 2005-12-07
Anticipated expiration: 2016-08-22
Also published as: CN1175017A

Abstract

The present invention relates to a pitching stability augmentation and command control augmentation system. A feedback system which is configured to receive a steering column input signal and convert the input signal into an elevator command signal is made use of. The feedback part processing of the system stands for a signal of current airplane data, and the signal is formed according to the previous elevator command. The stability augmentation of the system is finished by converting the steering column input signal into a C<*>U command signal which is required by a pilot and comparing the command signal and a C<*>U command signal calculated on the basis of the current state of the airplane. An error signal stands for the difference of the command signal required by the pilot and a response signal of the airplane, and the error signal and the elevator command signal are added after the error signal is integrated. Therefore, the pitching stability augmentation and command control augmentation system of the present invention realizes zero deviation between a pilot command and an airplane execution command.

Description

Aircraft pitch increases steady and the instruction control augmentation system

Technical field

The present invention relates generally to aircraft flight control system, particularly utilize pitch-axis stability augmentation and instruction control augmentation system to carry out the instrument and the method for aircraft evelvator instruction control.

Background technology

Some new aircrafts have on purpose relaxed the requirement to the aircraft static stability, comprise that aircraft has the exemplary aircraft of neutral static stability.This system is improving fuel efficiency, weight reduction and is reducing can obtain remarkable benefit aspect the resistance owing to can reduce required empennage size.Though the aircraft with the static stability of relaxing may have some undesirable and unacceptable handling quality characteristics.

A characteristic is exactly the flying speed of indifferent equilibrium aircraft before can't getting back to automatically after the pitch attitude change.On the positive static stability aircraft of a trim, pitch attitude with the driver to stick-forword movement or pull back jociey stick and change.When the pilot unclamps jociey stick, aircraft will continue in the short time along new angle of pitch flight.Along with the variation of aircraft flight speed, aircraft has the natural tendency that comes back or bow, and gets back to former trim speed.This trend is a kind of expected handling quality of conventional commercial aircraft, naturally and understandably can make a response during the pilot driver aircraft.

The indifferent equilibrium aircraft is not got back to the trend of trim position and flying speed after angle of pitch adjustment.The also corresponding change of flying speed when keeping the angle of pitch to change of these aircrafts.When the angle of pitch became big, flying speed reduced; When the angle of pitch diminished, flying speed increased.In order to make aircraft get back to trim position and trim speed, the pilot must use jociey stick to reorientate elevating rudder, makes aircraft get back to the trim angle of pitch.

Also have one to the aircraft that has or do not have the static stability of relaxing all very at large characteristic be that it is very sensitive with respect to the variation of mean aerodynamic center to weight change and centre of gravity place to finish the required elevating rudder input quantity of a certain angle of pitch.In light weight and center of gravity is near the aircraft of mean aerodynamic center, only needs the little change of the elevating rudder angle of attack just to produce the big change at aircraft pitch angle.Otherwise, a weight and have the aircraft of preposition center of gravity will produce the same angle of pitch to change, just need the bigger elevating rudder degree of bias.If control does not strengthen, for obtaining the angle of pitch response same with lightweight aircraft, the pilot must mobile biglyyer jociey stick when driving the aircraft of Heavy Weight.Because it serves as that the jociey stick input quantity is adjusted on the basis with aircraft weight and centre of gravity place that this situation needs the pilot, and these all are the information that can't obtain very soon, thereby this is a kind of undesirable characteristic.

Though the third characteristic is irrelevant with the aircraft that relaxes stability, it relates to wishes the mode that aircraft is made a response when the pilot moves jociey stick.As everyone knows, the pilot wishes that mobile jociey stick can produce the change of rate of pitch when hanging down flying speed.The pilot wishes that then mobile jociey stick can produce the change of normal acceleration when high flying speed.This situation has just been recognized as far back as the sixties, and has been developed the C of optimum aircraft response as Consideration ^*Criterion.To go through C below ^*Criterion.

Like this, just need the good pitch attitude control system that to improve aircraft control characteristic and handling characteristic.From optimal angle, in order to reduce the variable quantity of required airmanship, this pitch attitude control system should be aircraft provides the handling characteristic of the desirable conventional airplane that is similar to positive static stability (according to pilot's viewpoint).This desirable control system should not considered to the expectation of the short term response of elevating rudder instruction and long-term response and by in an identical manner pilot's pitch demand being made a response that flying weight or center of gravity move and reaches this requirement by satisfying the relevant aircraft of pilot.This ideal control system should further adapt to the expectation of pilot's control rate of pitch when low flying speed is handled and the expectation of controlling normal acceleration when high flying speed is handled.By reading following explanation, can find out that the present invention will directly provide this good pitch attitude control system.

Summary of the invention

The invention provides the pitch-axis stability augmentation and the order set that increase control aircraft evelvator instruction usefulness.The present invention uses and is called C herein ^*The criterion of U criterion is to provide desirable pitching response characteristic of aircraft and pitch axis long period speed stability.This C ^*The U criterion is by optimizing the response of aircraft pitch angular velocity and normal acceleration and providing the good speed stability characteristic of conventional airplane to finish the above-mentioned purpose of the present invention.

The feedback system that utilization of the present invention is configured to receive the jociey stick input signal and input signal is converted into the elevating rudder command signal.The feedback fraction of system is handled the signal of the current airplane data of representative, and this signal formed according to former elevating rudder instruction.By the jociey stick input signal being changed into C by pilot's requirement ^*U command signal, and the C that contrasts command signal and on aircraft current state basis, calculate ^*The U command signal is finished increasing surely of system.Error signal is represented the poor of pilot's command signal and aircraft response signal, behind the error signal integration with the addition of elevating rudder command signal.So pitch-axis stability augmentation of the present invention and instruction control augmentation system have been realized the zero-deviation between pilot's instruction and the aircraft execution command.

According to the present invention, pitch-axis stability augmentation and instruction control augmentation system comprise that one changes the jociey stick input signal pitch demand processor of feedforward command signal into, and this signal is one of two signals that produce the elevating rudder pitch demand.The secondary signal component is discussed below.Instruction processing unit can also be instruction C ^*The U processor provides the jociey stick position signalling of correction, instruction C ^*The U processor is correction jociey stick position transition the C that jociey stick that the representative requirement angle of pitch changes moves ^*U pitch demand signal.

According to a further aspect in the invention, calculate C ^*The U processor forms on the basis of aircraft current state and calculates C ^*U signal.The calculating C of currently preferred embodiment ^*U signal comprises three components at least: vertical acceleration signal, rate of pitch signal and speed error signal.In a preferred embodiment, vertical acceleration signal and rate of pitch signal all are to do reference with the earth, and are provided by normal acceleration complementary filter and rate of pitch complementary filter respectively.

Speed error signal is provided by the speed stability processor.The speed stability processor comprises the undulatory motion damping when forming speed error signal.The speed stability processor has further realized allowing the pilot to utilize the method for balancing device setting with reference to flying speed.

According to other aspects of the invention, provide three C ^*U compensation and guard signal further strengthen the handling characteristic and the response characteristic of aircraft.Owing the speed protection signal is provided by the stall protection processor; this processor comprises sets up minimum attachment device with reference to flying speed; the pilot can't trim when being lower than this speed, and this processor can also strengthen aircraft response characteristic and handling characteristic during aircraft is owed speed operation.Evening up compensating signal provides by evening up compensation processor, so that aircraft produces the action when running into ground during the similar landing and influencing.Supervelocity protection processor provides the supervelocity guard signal to instruct elevating rudder that aircraft is come back, as long as aircraft is in the supervelocity state of operation.Provide composite set the stall protection signal, even up compensating signal and the supervelocity guard signal is combined to form a C ^*U compensation and guard signal.The pilot just can have precedence over by mobile jociey stick generation counteracting pitching signal and owe speed protection signal and the speed of mistake guard signal control aircraft.Like this, by conscious selection, aircraft just can remain on the speed of owing or supervelocity state.Can reduce pilot work load's configuration compensation system when in addition, also having comprised configuration change.

According to other aspect of the present invention, provide another composite set to make up the C that the pilot requires ^*The subtraction input signal of U command signal, the C that calculates ^*The addition input signal and the C of U signal ^*The subtraction input signal of U compensation guard signal.The error signal that obtains offers integrator and exports an integrated error signal.Another composite set receives integrated error signal as the addition input signal, and it and the signal plus of representing rate of pitch damping instruction.Provide rate of pitch damping command signal to suppress short period response and adjustment cycle frequency.The output signal of composite set is as the addition input signal of the composite set with another feedforward instruction addition input signal.The signal that obtains like this is C ^*U increases steady elevating rudder command signal.

According to a further aspect of the invention, provide the structure modal filter with the frequency content in the elimination elevating rudder command signal.Provide stable offloading functions to adjust stabilizator so that in whole flight course, keep elevating rudder control.

The invention provides a kind of aircraft pitch and increase steady and command control system, have the response of representative instruction elevating rudder, jociey stick position, normal acceleration, and the input signal of aircraft pitch angular velocity; This aircraft pitch that is used to change the aircraft pitch angle increases steady and command control system presents stable speed responsive characteristic; This aircraft pitch increases steady and command control system comprises:

(a) one first signal combiner has first and second input ends and an output terminal;

(b) angle of pitch instruction processing unit is used to receive the jociey stick position signalling and is used to provide the first input end of feedforward elevating rudder command signal to first signal combiner;

(c) an instruction C ^*The U processor is used to receive the input that comprises the signal of representing the jociey stick position and is used to generate the output that comprises the signal of representing the response of instruction elevating rudder, and instruction elevating rudder response signal embodies required flight characteristics;

(d) one calculate C ^*The U processor is used to receive and comprises every input of representing aircraft normal acceleration, aircraft pitch angular velocity, air speed and these signals of reference velocity, described calculating C ^*The U processor generates according to the current flight state representative and calculates the output signal of elevating rudder response, calculates the elevating rudder response based on aircraft vertical acceleration signal, aircraft pitch angular velocity signal, air speed signal and reference speed signal at least;

(e) a secondary signal combiner is used to receive and calculates elevating rudder response signal and instruction elevating rudder response signal and be used to generate the output signal that difference between the elevating rudder signal and instruction elevating rudder signal is calculated in representative;

(f) integrator is used to receive described difference signal and is used for difference signal is guided into about null value, and integrator generates output signal, receives first signal combiner by second input end.

Also a kind of method that is used to generate the elevating rudder command signal that presents the stabilized speed response characteristic of the present invention comprises the steps:

(a) generate the instruction elevating rudder response signal that the required angle of pitch of representative responds;

(b) generate the representative calculating elevating rudder response signal that the expection angle of pitch responds under current state, comprise the steps:

(i) signal and the signal of the signal of representing air speed with generation representation speed error of reference velocity represented in combination, and

(ii) make up the signal of representing normal acceleration, the signal of representing rate of pitch, and speed error signal, the elevating rudder response signal calculated to form;

(c) combined command elevating rudder response signal and calculating elevating rudder response signal are to generate the output signal of the difference of representing them; And

(d) utilize output signal to revise the elevating rudder command signal.

The present invention provides one in addition for utilizing C ^*Improving one's methods of the aircraft pitch angle control system of criterion comprises:

(a) speed stability signal is used to make aircraft to be returned to its trim speed, and the speed stability signal comprises:

(i) represent the signal of air speed and represent the signal of reference velocity; And

(ii) one first composite set is used for the input of Combined toy plane speed and imports with the formation speed error signal with reference velocity, and speed error signal is used to formation speed stability signal; And

(b) one second composite set is used for group speed stability signal and C ^*Criterion.

The present invention provides a kind of method that is used to generate the elevating rudder command signal with speed stability again, comprises following each step:

(a) from the signal of representing the jociey stick position, generate feedforward elevating rudder command signal;

(b) generate the instruction elevating rudder response that the required elevating rudder of representative respond;

(c) generate the calculating elevating rudder response signal of representative, generate the response of calculating elevating rudder and comprise following steps based on the response of current state aircraft evelvator:

(i) combined reference rate signal and air speed signal are with the formation speed error signal; And

(ii) combination vertical acceleration input signal, rate of pitch signal and speed error signal calculate the elevating rudder response signal to form;

(d) combined command elevating rudder response signal and calculating elevating rudder response signal are to form the output difference signal; And

(e) utilize the output difference signal to revise the elevating rudder command signal.

The present invention provides a kind of sensation order set again, be used to have among the aircraft of jociey stick, the sensation order set is used to provide the power gradient to output signal to the device that is used to regulate the required power of mobile jociey stick, the sensation order set is used for increasing the required power of mobile jociey stick during near stalling speed or stall angle when aircraft, feels that order set comprises:

(a) one first power gradient component signal has a predetermined value that is relevant at least one flying quality input signal;

(b) one the 4th composite set has first and second input ends, and first input is rate signal or angle of attack signal, and second input is with reference to shaking bar angle of attack signal or with reference to the stalling speed signal, the 4th composite set generated error output signal;

(c) scaler is used for the round-off error output signal to form the second power gradient component signal: and

(d) one the 5th composite set receives the first and second power gradient component signals with generative power gradient output signal.

The present invention also provides a kind of bias voltage correction system, be used among the aircraft with jociey stick and angle of pitch control system, the position of jociey stick is by the representative of jociey stick position signalling, and the bias voltage correction system is used for suppressing any jociey stick bias voltage when jociey stick is in neutral position, and the bias voltage correction system comprises:

(a) one the 3rd composite set is used to make up jociey stick position signalling and correction signal;

(b) switch, has the moving state of conciliating that activates, switch is set to the state separated when in fact jociey stick removes its neutral position, the state of separating allows correction signal to be retained in its last numerical value place, and actuating state allows correction signal to equal the jociey stick position signalling haply.

Description of drawings

Can more easily understand above-mentioned various aspects of the present invention and attendant advantages thereof with reference to following detailed description in conjunction with the accompanying drawings, in the accompanying drawing:

Fig. 1 is the control chart that the whole pitch-axis stability augmentation instruction control augmentation system of formation is described according to the present invention:

Fig. 2 a is the control chart of pitch demand processor among Fig. 1;

Fig. 2 b is the control chart of sensation order set among Fig. 2 a;

Fig. 2 c is the control chart of jociey stick zero-deviation rejector among Fig. 2 a;

Fig. 3 is for instructing C among Fig. 1 ^*The control chart of U processor;

Fig. 4 a calculates C among Fig. 1 ^*The control chart of U processor;

Fig. 4 b is the control chart of turning compensation processor;

Fig. 4 c is the control chart of rate of pitch complementary filter;

Fig. 4 d is the control chart of normal acceleration complementary filter;

Fig. 4 e is the control chart of Fig. 4 a medium velocity stabilized treatment device;

Fig. 5 a is the control chart that compensates and protect processor among Fig. 1;

Fig. 5 b is the control chart of stall protection processor among Fig. 5 a;

Fig. 5 c is for evening up the control chart of compensation processor among Fig. 5 a;

Fig. 5 d is the control chart of supervelocity protection processor among Fig. 5 a;

Fig. 6 is the control chart of rate of pitch damping processor among Fig. 1;

Fig. 7 is C among Fig. 1 ^*The control chart of U integrator.

Embodiment

The present invention aims to provide a kind of stable and order set of pitching that any aircraft provides good chain of command response that can be, and has the aircraft that having of good handling characteristic relaxed stability when comprising pilot driver.The present invention proposes a kind of C of being called ^*The new ideas of U (pronouncing C asterisk U) criterion realize these requirements.C defined herein ^*The U criterion is the C with the middle nineteen sixties development ^*Criterion is the basis, will sketch it below.

General C ^*Criterion

Before middle nineteen sixties, the demarcation that the rate of pitch accepted that rate of pitch responds and the pilot obtains that the handling characteristic of the arbitrary specific longitudinal maneuver of aircraft normally produces by contrast jociey stick input signal responds is assessed.If response is in certain envelope curve, just think acceptable.This alanysis is used for assessing the performance of aircraft under high speed and lower-speed state.Yet confirmed that under fast state the most significant motor message prompting is the normal acceleration of aircraft concerning the pilot.Therefore, need find a kind of method to represent to consider simultaneously the acceptable aeroplane performance of rate of pitch and normal acceleration response.

The sixties, C ^*Analytic approach is developed into a kind of vertically aircraft response and the method for having considered the handling characteristic of chief aviation pilot's motor message prompting under low speed and the fast state assessed all.C ^*It is a kind of method that is used for short period flight control quality estimating.C ^*Variable-definition is rate of pitch and normal acceleration sum.

C ^*＝N _z+K _qq (1)

N wherein _zBe body normal acceleration, K _qBe hybrid gain, q is the body rate of pitch.

C ^*Variable is the sign of the vertical dynamic response characteristic quality of assessment aircraft.The reason of selecting rate of pitch and normal acceleration is that they are respectively the most significant motor message indication of pilot under low speed and fast state.Use C ^*During criterion, the analyst should be the C that calculates under any certain high-speed or low speed longitudinal maneuver state ^*The accepted C that value and pilot obtain ^*The demarcation of value contrasts.If response just thinks to have acceptable handling characteristic in the envelope curve scope.

Since having developed C ^*Since the criterion, some commercial and military aircrafts are C ^*Criterion is as the basis of Longitudinal Control System.Be based upon C separately ^*Control system on the conceptual foundation can change pilot's instruction into C ^*Instruction.Therefore, under low flying speed, pilot's instruction is regarded as the requirement of different rate of pitch, and under high flying speed, pilot's instruction is regarded as the requirement of different vertical acceleration.Be based upon C separately ^*Control system on the variable basis provides the neutral velocity-stabilization characteristic (that is, if air speed has changed, aircraft is not just got back to the trend of trim speed) of aircraft.Neutral speed stability is the specific character that many early stage aircrafts all exist.For the aircraft with neutral speed stability, the aircraft employee loads when obtaining reducing, and the pilot feels that by the power on the jociey stick consciousness of aircraft flight speed has weakened.

The present invention uses a new variables C ^*U, C ^*U is defined as above-mentioned C ^*The combination of a criterion and a speed stability item, thereby be called C ^*U (U representation speed stability).C ^*The U variable can be determined as follows:

C ^*U≡C ^*-K _VU _ERROR (2)

C wherein ^*Shown in defining in the equation (1), K _VBe speed stability gain, U _ERRORBe aircraft longitudinal velocity and C ^*Error between the U reference velocity.U _ERRORBe defined as:

U _ERROR≡V _CAS-V _REF (3)

V wherein _CASBe aircraft calibration flight speed, V _REFIt is a reference velocity.Relevant U _ERRORCan be according to equivalent flying speed, practical flight speed, the indication flying speed, with and/or mach indicator calculate K in these cases _VNumerical value will with here the narration difference.

In simplifying item, control law of the present invention has received a jociey stick input signal and has produced a feedforward C ^*The instruction of U elevating rudder.In addition, the C that imports according to the desirable RESPONSE CALCULATION pilot of aircraft ^*The U instruction.On the current aerodynamic characteristic basis of aircraft, calculate C according to every the calculating in the equation (2) ^*U signal.Calculate C ^*The instruction of U signal and pilot's input compares and has just drawn error signal.This error signal of integration, and it and feedforward C ^*The instruction of U elevating rudder and the rate of pitch damping packing of orders have just formed the output signal of elevating rudder pitch demand.This point is explained first below in detail.

As what can understand better from following explanation, though the present invention is designed to use for relaxing stable aircraft control system, and interrelate with a kind of like this system during explanation, be appreciated that, saying if desired, the present invention also can be applicable to the aircraft control system of other type.Secondly, according to convention,, be appreciated that in fact the present invention can finish with various distinct methods though illustration of the present invention and explanation take to comprise the form of the control law that designs the branch stile of finishing specific function.For example, the difference in functionality of described control law can be finished by the digital computer system of suitable programming.Another kind method is that these functions can be finished by digital circuit or mimic channel.

To be divided into seven parts below is elaborated.Overall enhanced pitch control system and its critical piece are discussed by first.Other parts go through each critical piece respectively.

1. the discussion of pitch-axis stability augmentation and instruction control augmentation system

Fig. 1 represents the control chart according to pitch-axis stability augmentation of the present invention and instruction control augmentation system 19.The output signal of pitch-axis stability augmentation and instruction control augmentation system 19 is elevating rudder instruction δ _E.FILT, it finally deliver to elevating rudder servo control mechanism instruction it and adjust elevating rudder.Fig. 1 is a reduced graph, thereby is provided with and comprises all input signals.Remaining figure will provide more information.

Elevating rudder command signal δ _{E, FILT}Before being sent by pitch-axis stability augmentation and instruction control augmentation system 19, earlier by 24 filtering of structural modal wave filter, and the input signal of structural modal wave filter 24 is the unfiltered elevating rudder command signal δ that formed by first composite set 20 _{E, UNF}Structural modal wave filter 24 is attempted the unfiltered elevating rudder command signal of elimination δ _{E, UNF}Structural modal wave filter 24 is attempted the unfiltered elevating rudder command signal of elimination δ _E.UNFThe frequency content of middle possibility countermeasure aircraft structural modal.Do like this and be for fear of bad and/or unstable flight characteristic and avoid when specific control is motor-driven, causing aircraft vibration or resonance.Preferred construction modal filter 24 depends on the ad hoc structure frequency of aircraft.This wave filter is included in the order set 19 a lot of modes can also be arranged.It can be placed on arbitrary place of order set 19, perhaps can be separated into different piece and is placed on the diverse location.

First composite set 20 makes C ^*U feedforward instruction C ^*U _FFCWith increase steady feedback command AFB _COMCombination.Feedforward instruction C ^*U _FFCWith by C ^*The pilot driver bar input signal δ that U pitch demand processor 26 is revised _CBe the basis.C ^*U feedforward instruction C ^*U _FFCThe preferably motion of relevant jociey stick, but the present invention comprises the instruction elevating rudder position input signal that use is come by robot pilot or other Controlling Source, replaces pilot driver bar input signal δ _C

Preferably, operation pitch demand processor 26 comprises " sensation " of management jociey stick to finish multitask, connects the external control system, eliminates the jociey stick zero-deviation, forms C ^*U feedforward instruction C ^*U _FFCAnd machine is opened robot pilot input signal mode in due course.Pitch demand processor 26 also calculates the jociey stick position signalling of a correction _{δ C, COR}Other parts for pitch-axis stability augmentation and instruction control augmentation system 19 are used.Certainly, for concrete an application, be not that all these functions all suit.The main task of pitch demand processor 26 is pilot driver bar input signal _{δ C}Change C into ^*U elevating rudder position (C here ^*U _FFCPreferably be unit with the degree) and the jociey stick position signalling that changes correction into _{δ C, COR}For instruction C ^*U processor 30 uses.Pitch demand processor 26 part 2 below goes through, and shown in Fig. 2 a, Fig. 2 b and Fig. 2 c.Instruction C ^*U processor 30 will go through in the 3rd part, and as shown in Figure 3.

Increase steady feedback command AFBCOM normally by in multiplier 25, revising C with elevating rudder loop gain coefficient 29 ^*U integration instruction IC ^*U _COMWith rate of pitch damping instruction Q _COMCombination and form.C ^*U integration instruction IC ^*U _COMRepresent C ^*The integration of U error signal, C ^*U such as top equation (2) definition.Rate of pitch damping instruction Q _COMFor the adjustment of short period response and short period frequency provides damping.Rate of pitch damping instruction Q _COMTo go through in the 6th part below, and as shown in Figure 6.IC ^*U and Q _COMCombination finish by composite set 23.Combination is revised by multiplier 25 and is handled.Elevating rudder loop gain coefficient 29 is used to flutter mode that enough stability margins are provided.It is the function (being Mach number or calibration flight speed) of flying speed.The concrete condition that will use according to the present invention and embodiment are not that all these class factors discussed herein are all necessary.Similarly, can revise factor to produce elevating rudder command signal δ with the mode of any hope _{E, FILT}The present invention includes all these type of changes.

As viewed from Fig. 1, the preferred embodiments of the present invention are also unloaded borne processor 21 utilizations by stabilizator and are increased steady feedback command AFB _COMProduce a trim erasure signal TNS.Stabilizator unloads borne processor 21 and adjusts stabilizator to guarantee obtaining elevating rudder control in whole flight course.In case increase steady feedback signal instruction AFB _COMExceed a certain threshold position, stabilizator just begins to move.Along with moving of stabilizator, the position that stabilizator position transducer (not shown) perception stabilizator unloads borne processor 21, after there it and gain being multiplied each other again differential signal TNS is offered integrator 28 is added to integration C again with formation trim negative acknowledge character (NAK) TNS ^*In the U instruction.By this way, stabilizator moves and controls with additional to elevating rudder, and the resistance that makes the elevating rudder unloading also help to reduce on the aircraft is born.It is known that suitable stabilizator unloads borne processor 21, can obtain in the present invention using.For example, the trim general autoflight system of ignoring instruction provides.Novelty of the present invention comprises unloads that part that borne processor 21 is positioned at pitch-axis stability augmentation and instruction control augmentation system 19 to stabilizator, and it only represents elevating rudder pitch demand signal δ _{E, FILT}Increase steady part.This point is very important, because do not wish mobile stabilizator under the short period maneuvering condition.

C ^*U integration instruction IC ^*U _COMRepresent the feedback fraction of pitch-axis stability augmentation and instruction control augmentation system 19, when not having pilot's input signal, its instruction elevating rudder is got back to trim condition.Pilot driver bar input signal δ is being arranged _CThe time, C ^*U integration instruction IC ^*U _COMGuarantee that the response of aircraft satisfies pilot's requirement.C ^*U integration instruction IC ^*U _COMBy C ^*U error signal EC ^*U passes through C ^*U integrator 28 integrations form.C ^*U integrator 28 the 7th part below goes through.Error signal EC ^*The C that the integration of U can make pitch-axis stability augmentation and instruction control augmentation system 19 require the pilot ^*U instructs C ^*U _PoilotCmdWith the C that calculates ^*U signal C ^*U _ComputedWith compensation guard signal C ^*U _{C﹠amp; P}And between constantly attempt to eliminate any difference.

The C that the pilot requires ^*U instructs C ^*U _PilotcomdIn basic function is correction jociey stick position signalling δ _C.CORBe converted into C ^*The instruction C of U-shaped formula ^*Form in the U processor 30, because of being called as the C that the pilot requires ^*U instructs C ^*U _PilotcomdInstruction C ^*U processor 30 the 3rd part below goes through.Calculate C ^*U signal C ^*U _ComputdCalculating C according to equation (2) ^*Form in the U processor 32, and with the 4th part discussion below.C ^*U compensation guard signal C ^*U _{C﹠amp; P}In compensation protection processor 34, form, and with the 5th part discussion below.

The variation of the control loop that requires for finishing the work among the present invention may be very widely.Therefore the total of being explained here should be counted as exemplary, rather than restrictive.Though provided preferred embodiment here, yet be appreciated that still some application may not need to comprise all factors.Similarly, in order to obtain the controlling factor that required particular result can add that some are known, even be not factor that this paper has provided or that discussed.

In addition, in the process of the system and method for the present invention of Miao Shuing, on some points of system, used the notion of gain and transport function in a preferred embodiment.Here use the notion of transport function be interpreted as widely comprising time lag or leading, amplify aspects such as (linear or non-linear), decay, integration.Like this, definition has comprised the notion of gain.Various forms of transport functions often with the Laplace transform symbolic representation, and can realize by analog form or digital form on mathematics.The example of analogue means comprises linear amplifier, capacitor, inductor, resistor and comprises a part or the network of all these devices.The example of digital device comprises the digital processing unit of binary adder-subtracter, comparer, reverser, control route marker and sequential processes digital encoded data.

Here used " composite set " this speech has broad interpretation, comprises the signal combination on the digital scale-of-two meaning, therefore comprises addition, subtraction, multiplication and division.These composite sets can comprise analogue means, amplify or transistor such as addition, also can comprise digital device, such as the shift register in the arithmetical logic device of binary adder, subtracter, comparer or central processing unit.

2. the discussion of pitch demand processor 26

2a. pitch demand processor 26

Fig. 2 a is the block diagram of pitch demand processor 26, and it is made brief description with respect to total system arrangement among Fig. 1.As previously mentioned, the basic function of pitch demand processor 26 is pilot driver bar input signal δ _CChange elevating rudder instruction δ into _C.FILTFeedforward instruction C ^*U _FFCPart, and generate the jociey stick position signalling δ that revises _{E, COR}For instructing C among Fig. 1 ^*U processor 30 is used.In the described arrangement of Fig. 2 a, pilot driver bar input signal δ _CBe coupled to sensory system 56.As known to, sensory system has been set up mechanical force on jociey stick, mobile jociey stick than midline position (card off-position put) in advance or backward the time this power must overcome.As known to, the power that sensory system is set up increases with the increase of jociey stick skew, and the characteristic of control force changes (need less power during low-speed operations, need bigger power during high-speed flight) with the change of aircraft flight speed.

Though the relation of very common power concerning prior art (being provided by sensory system 56 among Fig. 2 a) has been provided in the present invention, the relation that the present invention has replenished various power as follows changes the characteristic of control force: (1) reminds the pilot to owe speed state and supervelocity state; For providing the jociey stick of expection, the pilot evens up response when (2) landing; (3) the aircraft angle of attack reaches and/or control force increased when exceeding stall value (α ss) rapidly.This method that the present invention constitutes reaches and describes with reference to Fig. 5 a～Fig. 5 c in this method of owing to operate under speed state, supervelocity state and the landing flare state.When aircraft near or the present invention can increase the instruction driving stick force rapidly when being in stall conditions characteristic by Fig. 2 a in shown in the sensation order set 52, and describe with reference to Fig. 2 b.

Go on to say Fig. 2 a, sensory system 56 provides instruction driving bar displacement signal 58 and it is coupled on the jociey stick zero-deviation rejector 60.Jociey stick zero-deviation rejector 60 receives the linear variable differential transform device of the relative midline position of indicator jociey sticks (the card off-position is put) skew or other signal of sensor (by among Fig. 2 a shown in the arrow 58).Described with reference to Fig. 2 c, jociey stick zero-deviation rejector 60 has compensated the signal bias that exists in the jociey stick position sensor output signal or deviation (non-zero signal that sensor provided when for example the instruction driving bar was in the card off-position and puts).When the instruction driving bar was in the card off-position and puts, jociey stick zero-deviation rejection characteristic of the present invention can guarantee that aircraft is in trim condition and (for example equals C in flying speed ^*U reference velocity V _RFFThe time).

The jociey stick position signalling δ that revises _C.CORForm by jociey stick zero-deviation rejector 60 among Fig. 2 a, it is offered the transfer device 64 of Fig. 2 a, and as shown in Figure 1, it also is with and has offered instruction C ^*U processor 30, this point describes with reference to Fig. 3.

Transfer device 64 among Fig. 2 a is the jociey stick position signalling δ of the correction that is provided by jociey stick zero-deviation rejector 60 _C.CORChange signal C into ^*U _FFC, signal C ^*U _FFCBe elevating rudder command signal δ _{E.U * F}A part.As known to, the conversion of signals that transfer device 64 carries out is by other characteristic domination of used specific sensation system 56 and aircraft.As known to, the signal transformation of being undertaken by transfer device 64 can be listed in the function of flying speed, also can be arranged for showing the frequency response of any needs or hope.

Shown in the switch among Fig. 2 a 66, when aircraft did not use robot pilot, the signal that is provided by transfer device 64 was coupled on the adder 20 among Fig. 1 and as C ^*U feedforward command signal C ^*U _FFCWhen using robot pilot, signal is provided for switch 66, and robot pilot 68 provides its C ^*U feedforward instruction C ^*U _{A/P, FFC}Pattern.In addition, as previously mentioned, robot pilot can produce equivalent δ _CPerhaps mobile jociey stick.Under these states, switch 66 keeps state shown in Fig. 2 a.

2b. sensation order set 52

Two component of signals 103,97 that sensory system 56 uses among the sensation order set 52 constitutional diagram 2a shown in Fig. 2 b are used to produce suitable instruction driving stick force gradient.The sensation instruction type that the first sensation command signal component, 103 matching convention aircraft sensory systems use.Particularly, arrangement according to Fig. 2 b, represent the signal (with arrow 102 indications) of one or several flying quality parameter to offer look-up table 100, perhaps under another kind is arranged, the flying quality signal transition for showing as aircraft flight speed with such as the sensation command signal of the required power gradient of the function of other parameters such as aircraft pitch angular velocity.Arrangement according to Fig. 2 b, the sensation command signal 103 that is provided by look-up table 100 does not directly offer aircraft sensory system (for example sensory system among Fig. 1 56), on the contrary, it is provided for composite set 98, and it and the second sensation command signal component 97 are combined there.

According to the present invention and as described below, low flying speed state (or high angles of attack state) down the second sensation command signal component significantly change of flight person keep or move (new line instructs) required power when being in the jociey stick of card off-position after putting.According to the arrangement of Fig. 2 b, represent the signal of the aircraft filtering angle of attack (perhaps also can be described as the filtering flying speed) (shown in arrow 50) to offer composite set 90 as the addition input signal.The subtraction input signal that is added to composite set 90 is that a kind of angle of attack of representing slightly surpasses the shake signal (by among Fig. 2 b shown in the arrow 104) (or representing the signal of flying speed when shaking the bar flying speed) of bar angle of attack interval scale of aircraft.As known to, the shake signal of the bar angle of attack or the bar flying speed of shaking of representative can obtain from the system that stall warning can be provided such as alarm electronic system or other.

The output signal of composite set 90 offers the scaler 92 with a gain factor, and this gain factor may be constant or may calculate according to one or several flying quality signal of aircraft flap setting and/or indication aircraft flight speed.Scaler 92 has been determined the power gradient, and when aircraft was in the bar angle of attack part of shaking (or being lower than the bar flying speed of shaking) operation, the power gradient was added on the nominal sensation command force gradient.The gain of regulating scaler 92 according to wing flap and/or flying quality is also unnecessary as a rule, but can be comprised, so that obtain the accurate control as the variation of the control force of flying speed function.

The signal that is provided by scaler 92 is represented the filtered angle of attack and is slightly exceeded difference between the desired value of the bar angle of attack that shakes (or filtered flying speed and a little less than the difference between the suitable flying speed of the bar speed of shaking), and it is offered limiter 94.When the filtered angle of attack was not higher than the bar angle of attack that shakes (deducting little nargin), perhaps when filtered flying speed was not less than the bar flying speed of shaking (adding little nargin), limiter 94 had a limit value above freezing to eliminate the signal of generation.When aircraft exceeded the bar angle of attack that shakes (or being lower than the bar flying speed of shaking), the lower limit of limiter 94 was selected as determining that Fig. 2 b provides can determine that jociey stick is whether in the maximum, force at bayonet lock stop position rear portion.

The layout of Fig. 2 b continue to be discussed, and the signal that is provided by limiter 94 offers the second addition input signal of composite set 98 by speed restrictor 96.Speed restrictor 96 filtering or smoothing offer composite set 98 behind the signal and eliminate the signal height that may occur in the airborne signal of represent flying speed after the aircraft angle of attack and the filtering and suddenly change.As known to, other different layout also can be used to be similar to mode " smoothing " signal of Fig. 2 b medium velocity limiter 96 as the back wave filter that stagnates.As aforementioned, the signal that is provided by composite set 98 among Fig. 2 b offers aircraft sensory system (56 among Fig. 2 a) as system's sensation command signal 54.

2c. jociey stick zero-deviation rejector 60

The preferential at present control law that uses of jociey stick zero-deviation rejector 60 usefulness among Fig. 2 a is illustrated by Fig. 2 c.Description with reference to Fig. 2 a, the present invention adopts jociey stick position zero drift correction to come compensating offset formula offset component, when the dependent instruction jociey stick was in midline position (the card off-position is put), these skew formula deviation branches may appear in the signal that is provided by instruction driving bar position transducer.

In the described layout of Fig. 2 c, the signal that is provided by instruction driving bar linear variable differential converter (or other sensor) (with arrow 58 indications) is provided as the addition input signal of composite set 82, and the output signal of composite set 82 is for revising jociey stick position signalling δ _C.COROffer the subtraction input end of composite set 82 by compensation provided by the invention or corrected signal 81.

Shown in Fig. 2 c, provide by integrator 80 by corrected signal 81 provided by the invention, integrator 80 also links to each other with the subtraction input signal of composite set 72.The addition input signal of composite set 72 is provided by limiter 70 by instruction driving bar position sensor signal (arrow 58).The output signal of composite set 72 is by the gain factor K of scaler 74 ₁Regulate, and offer an end of switch 78.The switch slide plate links to each other with the input signal of integrator 80.Shown in frame 76, when connecting, switch 78 (for example is not in position shown in Fig. 2 c), and integrator 80 does not just have input signal.

The layout operation of Fig. 2 c is as follows.When not using aircraft autopilot and jociey stick to block off-position relatively to be equipped with very little displacement or not have displacement, switch 78 is connected and the input of integrator 80 and the output of scaler 74 is linked.Like this, when switch 78 is connected, the signal that is provided by instruction driving bar sensor links by the input with integrator 80 of limiter 70, composite set 72 and scaler 74, simultaneously the output signal that is provided by integrator 80 is fed back to known to the subtraction input end such as Technology professional of composite set 72, and this feedback arrangement meets the stagnate arrangement of wave filter of conventional back.Like this, except that the short time postponed, (as long as instruction driving bar sensor signal is within the scope of limiter 70) was provided with the signal that is provided by instruction driving bar sensor the zero-deviation corrected signal 81 that offers the subtraction input of composite set 82.Therefore, when jociey stick remains on that the card off-position is put or midline position and jociey stick sensor signal when being in limiter 70 scopes, the jociey stick control signal as many as zero of the correction that provides by composite set 82.In this respect, the scope that limiter 70 is set is within the limiter scope typical instruction driving bar sensor bias formula shifted signal, but quite different to bigger signal.

When exert oneself mobile jociey stick of pilot breaks away from the card off-position when putting, switch 78 disconnects, and causes integrator 80 not have signal input (shown in the frame 76).When not having signal to input to integrator 80, its output signal remains unchanged.Like this, the signal that offers composite set 82 subtractions inputs keeps the skew or the offset component of and instruction jociey stick sensor to equate.The deviation or the offset component that this means instruction driving bar sensor signal will can not appear in the jociey stick position signalling of the correction that is provided by composite set 82.

When steering force deficiency or force transducer indication power lost efficacy or made mistakes, switch 78 was started.Even this just allows the zero-deviation rejector also can work well under control force sensor signal failure state, and can make when putting mobile jociey stick switch 78 keep on-states from the card off-position as the pilot.Particularly, under a kind of like this failure state, the signal routes that comprises composite set 72, scaler 74 and integrator 80 will continue as stagnate wave filter and working of back.Yet, can not exceed the scope of limiter 70 owing to offer the input signal of composite set 72 additions input, thereby only have the loss of a few instructions jociey stick position signalling.Particularly, the skew of linear variable differential transform device of the present invention and other position transducer or deviation signal are not more than a few percent of maximum jociey stick degree of bias signal usually.Based on this, when the instruction driving bar leaves the card off-position and puts, if switch 78 just keeps connecting and aircraft pitch to be controlled required pilot's instruction driving bar almost do not influence or do not have a noticeable influence.And when in fact the instruction driving bar was in the card off-position and puts, it was zero output signal that this arrangement can continue to provide actual.

The instruction C ^*The discussion of U processor 30

As according to Fig. 1 discussed, the jociey stick position signalling of the correction that is provided by jociey stick zero-deviation rejector 60 in Fig. 2 offers instruction C again ^*U processor 30, this processor are transformed into the C that represents pilot's input with the jociey stick position signalling of revising again ^*U pitch demand signal (is the C among Fig. 1 ^*U _PilotcmdSignal).Same as in Fig. 1, being discussed, pilot C ^*U pitch demand signal is provided for composite set 22, and there it again with C ^*U compensation guard signal C ^*U _{C﹠amp; P}(in Fig. 5 a to Fig. 5 d, discussing) and C ^*The calculated value C of U pitch demand ^*U _Computed(will in Fig. 4, discuss) combination.

At instruction C shown in Figure 3 ^*In the U processor 30, the control jociey stick position signalling of correction _C.COROffer wave filter 110 or another equivalent device, this device can be with can be with the jociey stick position signalling δ that revises _C.CORThe mode that is transformed into the pitch control subsystem signal with desirable characteristics provides signal to form (for example, the required relation between control jociey stick displacement and the pitch attitude instruction).Represent the signal of surge pressure (or other represent the flying quality signal of flying speed) to offer scaler 112 shown in the arrow 120.A gain period set up by scaler 112 so that the output signal that Fig. 3 device is provided finally becomes the C that shows the stick force of consistent relatively every gram in whole flight envelope scope ^*U signal.The signal plus that multiplier 116 provides the signal and the wave filter 110 of scaler 112 output and export a desired C of pilot ^*U instructs C ^*U _PilotemdShown in switch 118, no matter whether robot pilot (among Fig. 3 68) switch, and the signal that multiplier 116 is exported all is input to the adder 22 among Fig. 1.If autopilot engagement is by the C of robot pilot generation ^*U command signal C ^*U _A/P.CMDBe input to adder 22 among Fig. 1 by switch 118.Under the situation that the robot pilot instruction occurs with the jociey stick instruction, switch 118 will remain on position shown in Figure 3.

4. calculate C ^*The discussion of U processor 32

4a. calculate C ^*U processor 32

Calculate C ^*U processor 32 is that pitch-axis stability augmentation and instruction control augmentation system are used for definite C based on current (referring to feedback here) value ^*The part of U.Produce following discriminant criterion in conjunction with equation (1) and (2):

C ^*U＝N _z+k _qq-K _V(V _CAS-V _REF) (4)

By calculating C ^*The signal processing that U processor 32 carries out is realized this process.

As will partly illustrating, at the C of presently preferred embodiment at 4c part and 4d ^*In the execution of U, vertically quicken N _zWith aircraft pitch angular velocity q be reference with the earth rather than with the airframe.

C in the equation (4) ^*Part is produced by the composite set among Fig. 4 a 132, and composite set 132 has a vertical acceleration signal N _zAs the addition input signal.Rate of pitch signal q is by the gain factor K of scaler 140 _qRegulate.The output signal of scaler 140 offers composite set 132 as the second addition input signal.Composite set 132 provides C ^*Criterion, input signal N _zPreferably described in Fig. 4 b, 4c and 4d, form with q, wherein vertical acceleration signal N _zPreferably the turning with respect to earth reference system compensates vertical acceleration signal (based on the symbols Z among Fig. 4 c _EST).Similarly, rate of pitch signal q preferably one with respect to the turning of earth reference system compensation rate of pitch signal (based on the symbol theta among Fig. 4 d _EST).Be used to form Z _ESTAnd θ _ESTProcessor will be below 4c part and 4d partly discuss.

The second portion speed stability signal K of equation (4) _V( _CAS-V _REF) in speed stability processor 151, form.C ^*Criterion is input to composite set 134 as input signal, speed stability signal K _V(V _CAS-V _REF) also import composite set 134.Composite set 134 provides whole calculating C ^*U signal C ^*U _ComputedGenerate the described error signal E of top part 1 for the composite set 22 among Fig. 1 _C*U

Calculating C ^*In the preferred embodiment of U signal processor 32, switch 154 ability permissible velocity stability processors 151 provide speed stability signal K when having only the robot pilot of use _V(V _CAS-V _REF).When using robot pilot, switch connection also receives zero input signal.Used robot pilot comprises the instrument of controlling flying speed in the currently preferred embodiments of the invention.If do not use robot pilot in the application-specific of the present invention, there is not flying speed control in perhaps using, this particular switch 154 just there is no need so.

According to classical C ^*Criterion, rate of pitch gain K _q=0.217.Can select different value according to pilot's hobby.But this value should be of value to C when practical flight speed is 400 feet per seconds ^*Generation, the C of generation ^*In, N _zWith q same contribution is arranged all.This is used to form N _z, q and K _V(V _CAS-V _REF) value be based upon on the basis of aircraft normal acceleration, rate of pitch and velocity error.These formation will be discussed below.

4b. turning compensation processor 200

The turning compensation processor 200 that Fig. 4 b shows is used for generation and is used in calculating C ^*Vertical acceleration signal N in U processor 32 and the speed stability processor 151 _zWith rate of pitch signal q.Turning compensation processor 200 is known in stability and Control System Design technical elements, thereby only does general the argumentation here.When craft inclination, effect lift aboard is still perpendicular to aerofoil.This causes the minimizing of the lift that vertically acts on earth surface, and then makes the aircraft altitude loss.For continuing the height flight when aircraft does not tilt, flight must increase lift and compensate the minimizing of lift perpendicular to the earth surface direction.

Therefore provide the compensation of turning to be used to increase the elevating rudder angle of pitch and compensate the lift of keeping the required increase of flying height, and do not need the required extra jociey stick input signal δ of pilot _CThe input signal of turning compensation processor 200 comprises roll attitude signal 170, pitch attitude signal 172, with reference to the vertical acceleration signal 174 of fuselage, rate of pitch 176, flight path vertical acceleration signal 150 and pitch attitude rate signal 152 with reference to fuselage.The output signal of turning compensation processor 200 comprises the compensation vertical acceleration signal N that turns _zWith the compensation rate of pitch signal q that turns.In current embodiment of the present invention, the compensation of turning only provides when the pitch angle is spent less than 30, then removes gradually during the high dip angle.

The present invention by use representative with the earth be with reference to rather than improve the indemnification arrangement of front with inertia as the flight path vertical acceleration signal 150 of reference and the improvement input signal of pitch attitude rate signal 152.The formation that with the earth is the flight path vertical acceleration signal 150 of reference will partly be discussed and shown in Fig. 4 c at 4c.The signal that with the earth is the pitch attitude rate signal 152 of reference will partly be discussed and shown in Fig. 4 d at 4d.

4c. normal acceleration complementary filter 264

It is the flight path vertical acceleration signal 150 of reference that normal acceleration complementary filter 264 produces with the earth, and it is the vertical acceleration signal N of reference that the latter transfers to generate with the earth _zUsually, normal acceleration N _zIt is aircraft acceleration about inertial reference system.Pitch-axis stability augmentation of the present invention and instruction control augmentation system 19 are attempted to control with the earth to be reference system rather than to be the aircraft acceleration of reference system with the aircraft.Use with the earth as reference system rather than can bring as the normal acceleration of reference system with inertia can be with the good handling characteristic (when for example aircraft have pitching input instruction also can not orbit the earth) of constant vertical speed (comprise zero vertical speed) around ground flight.Control is that the normal acceleration of reference system just can not resemble with the earth with inertia is aircraft to be flown the reference system with constant vertical speed.

Be that to obtain with the earth behind the vertical speed Z differential of reference be the normal acceleration Z of reference with the earth, still, can produce unwanted during the Z differential is not unconspicuous noise content.Shown in Fig. 4 c, the present invention uses complementary wave filter to come the vertical acceleration signal Z of combination results _MFTDForming with the earth with Z is the flight path normal acceleration estimated signal Z of reference _EST, the latter can be used for substituting the flight path vertical acceleration signal 150 of turning compensation processor 200.This has just caused turning processor 200 should provide one, and that turn compensation is the vertical acceleration signal N of reference with the earth _{Z, e}Signal N _{Z, e}Than common vertical acceleration signal N _zBetter, because it allows aircraft also can fly around ground under the situation that does not have pilot's input signal.With the earth is the normal acceleration N of reference _{Z, e}Can substitute vertical acceleration signal N _zOffer and calculate C ^*The other parts of U processor 30 and pitch-axis stability augmentation and instruction control augmentation system 19.

Vertical acceleration signal Z _MFTDGeneration be to be that the acceleration of reference is realized with the fuselage with the earth by mapping on the Z axle that is reference.This process is known in aircraft stability and Control System Design technical elements, and indicates in Fig. 4 c by acceleration processor 272.The vertical speed Z that with the earth is reference normally receives from the flight data system with inertial reference system.

Shown in Fig. 4 c, be the flight path normal acceleration estimated signal Z of reference with the earth _ESTProvide by composite set 282 with two addition input signals.First signal is a part of Z of the vertical acceleration signal that generates in the z-y plane _NzySecond signal is by multiply by gain factor K _EzThe vertical speed error E that provides of scaler 270 _zThe input signal of scaler 270 is that the composite set 254 of input of the vertical speed Z of reference and the signal integration that provided by composite set 278 provides by having with the earth, and integration is to finish on the integrator 268 in Fig. 4 c.The signal that is provided by composite set 278 is by the flight path normal acceleration estimated signal Z that is reference with the earth _ESTA part of Z with the vertical acceleration signal that generates along the aircraft longitudinal axis direction _NxAddition input form.

As known to, it be that the vertical speed z of reference drives with the earth is the flight path normal acceleration estimated signal Z of reference that normal acceleration complementary filter 264 utilizes with the earth _ESTSteady-state characteristic.With the earth is the flight path normal acceleration estimated signal Z of reference _ESTBe used for C ^*U computation processor and C ^*And/or C ^*In the U process.That turn compensation is the vertical acceleration signal N of reference with the earth _{Z, e}Preferentially be used to calculate C ^*The U processor, though can use also that turning compensated with the not compensation normal acceleration estimated signal Z of inertia as reference _ESTOr uncompensated be the vertical acceleration signal 150 of reference with inertia

4d. rate of pitch complementary filter

The rate of pitch complementary filter produces the pitch attitude rate signal 152 with respect to the partial water plane.Though normally with respect to the aircraft angular velocity of inertial reference system, pitch-axis stability augmentation of the present invention and instruction control augmentation system 19 are with reference to the pitch attitude speed of earth control aircraft for pitch attitude speed.Use with the earth to the rate of pitch of reference than favourable as the rate of pitch of reference, because it can provide the constant ability of maintenance flying height when fly around ground through inertia.Control is that the rate of pitch of reference can't make aircraft self fly around ground with the fuselage, needs the pilot to revise.

Be that to obtain with the earth behind the pitch attitude θ differential of reference be the rate of pitch θ of reference with the earth, can bring a part unwanted during differential θ is not unconspicuous noise content.Shown in Fig. 4 d, the pitch attitude rate signal θ that the present invention uses a complementary filter combination to generate _MFTDForming with the earth with θ is the pitch attitude velocity estimation signal θ of reference _EST, θ _ESTCan be used for substituting the pitch attitude rate signal 152 of turning compensation processor 200.This has caused turning processor 200 to provide one, and that turn compensation is the pitch attitude rate signal q of reference with the earth _eThis signal q _eBe better than common pitch attitude speed q, because it provides favourable long-term aircraft response.With the earth is the pitch attitude rate signal q of reference _eCan replace offering calculating C ^*The pitch attitude speed q of the other parts of U processor 30 and pitch-axis stability augmentation and instruction control augmentation system 19.

The pitch attitude rate signal θ that generates _MFTDLuffing speed q, lift-over speed γ and yawing velocity p by mapping inertia reference on earth reference system form.This task is known in aircraft stability and Control System Design technical elements, and passes through luffing speed processor 330 by shown in Fig. 4 d.The pitch attitude θ that with the earth is reference normally receives from the flight data system with inertial reference system.

Shown in Fig. 4 d, be the pitch attitude velocity estimation signal θ of reference with the earth _ESTBe to provide by composite set 332 with two addition input signals.First signal is the pitch attitude rate signal θ that is provided by luffing speed processor 330 _MFTDSecond signal is pitch attitude error signal E _θ, by having gain factor KE _θScaler 324 provide.The input signal of scaler 332 is provided by composite set 314.It is that the pitch attitude signal θ of reference is the pitch attitude velocity estimation signal θ of reference as addition input signal and one with the earth that composite set 314 has with the earth _ESTIntegration as the subtraction input signal.

Finish on the integrator 328 of integration in Fig. 4 d.

As known to, the rate of pitch complementary filter uses with the earth and adopts acquisition with the pitch attitude velocity estimation signal θ of the earth as reference as the pitch attitude θ of reference _ESTStability characteristic (quality).That turn compensation is the pitch attitude rate signal q of reference with the earth _ePreferentially be used to calculate C ^*The U processor, though also can use the compensation of turning with inertia as the rate of pitch q of reference, estimate θ as the uncompensated rate of pitch of reference with the earth _ESTOr uncompensated be the pitch attitude rate signal 152 of reference with inertia.

4e. speed stability processor 151

Speed stability processor 151 provides a speed stability feedback signal K who comprises direct trim signal 372 and fluctuating damping feedback signal 400 _VU _ErrorAnd speed stability processor 151 is determined the C based on the current location of pitching trim position 354 uniquely ^*U reference velocity V _REF

In described configuration, reference velocity V _REFSet up by pitching balancing device 354, for example hand is turned on and is closed or the passageway montant, and they have three positions, makes the normal position (access failure) of balancing device 354 produce null pitching trim signal 364.

When the pilot adjusts pitching balancing device 354, produce representative-1 or+1 pitching trim signal 364, this depends on that the pilot comes back or bows in trim.Pitching trim signal 364 multiply by the gain k at multiplier 370 places _t, be converted into rate of change (signal 378) again with reference to flying speed.As long as the pilot is adjusting pitching balancing device 354, the rate of change to reference flying speed 378 carries out integration at integrator 376 places with regard to continuing, and has so just formed a new reference speed signal 382.If needed, k _tValue can be the function of flying quality.

374,375 and 380 and V among Fig. 4 e _REFThe setting of initial value relevant.374 is initialization and the synchronization activity processor with an output signal trigger 375, when output signal trigger 375 is set at the input signal of its output signal of true time as integrator 376, at this moment, filtered flying speed V _CasShown in the dotted line among Fig. 4 e 380.When output signal trigger 375 is set at fictitious time, integrator 376 uses the rate of change of above-mentioned reference flying speed signal 378.

Initialization and synchronization activity processor 374 used logics will preferentially be set output signal trigger 375 under following any state be true, that is: (1) as long as the longitudinal acceleration of aircraft does not reduce to the preset value that is lower than in preset time, output signal trigger 375 is set at very in the schedule time T after taking off so in preset time T; (2) as flying speed difference U _ErrorBe in the predetermined scope and when this moment, balancing device was got back to null value, output signal trigger 375 was set at very.First condition is for providing an initial reference speed after taking off.Second condition is to finish the required working load of trim aircraft in order to reduce the pilot.

Initial reference rate signal 382 is limited by the limiter 384 with minimum and maximum speed limit.These restrictions are based upon on the basis of the flight envelope of being determined by the aircraft particular configuration.Minimum limit is described by the stall protection processor part of the 5th part.Using maximum speed limit is in order to make the reference velocity V of formation _REFAnd be equal to or less than maximum operating speed V with reference to Mach number accordingly _MOAnd maximum operation c Mach number M _MO

Though the pilot can adjust reference velocity V by pitching balancing device 354 _REF, but think that the speed that the trim aircraft exceeds outside the velocity range that limiter 384 provides is impossible.These characteristics suit the requirements, because it requires the pilot that jociey stick is remained on non-midline position when aircraft flies with speed of owing or supervelocity, thereby warning pilot aircraft is in speed state or the supervelocity state owed.

The reference velocity V that forms _REFThen at composite set 398 places and current filtered calibration flight speed V _CASCombine and form flying speed difference U _ERRORFiltered calibration flight speed is with V _CASOffer speed stability processor 151 for the basis and by Flight Data Unit.V _CASBefore delivering to processor 151 fully filtering to get rid of owing to disturb the signal content that brings.If set up suitable related gain and plan, also can use other flying speed signal, as Mach number or unfiltered demarcation flying speed if needed.

At composite set 396 places, U _ERRORBe combined to form reinforcement U with direct trim signal 3 72 _ERROR(signal 404).Directly trim signal 372 is calculated by direct trim signal processor 356.Directly the effect of trim signal 372 provides additional elevating rudder instruction so that the direct response of the emulation conventional airplane elevating rudder relevant with the balancing device connection.The flying quality of flying speed or flap configuration are used to generate direct trim signal 372, and directly whether trim signal 372 depends on the shown aircrafts of input signal 364 just with upwards pitching or downward pitching are imported appropriate signals by trim.

Strengthen U _ERRORThe flying speed gain K of signal 404 and multiplier 406 _VMultiply each other.This has caused initial velocity stability feedback signal 408.Flying speed gain K _VBe used to set the steering force on the required jociey stick that calculates with pound/joint when flying speed relative reference speed has deviation.If desired, K _VValue can be the function of flying quality.

At last, composite set 410 is stablized feedback signal 408 to initial velocity and is combined from the undulatory motion damping feedback signal 400 that undulatory motion damping feedback processor 344 receives.At speed stability feedback signal K _VU _ERRORIn comprise the undulatory motion damping purpose be to provide enough dampings for undulatory motion mode.Such processor 344 is known in aircraft stability and Control System Design technology.The signal that causes is speed stability feedback signal K _VU _ERRORAnd be used for aforementioned calculation C ^*U processor 32.At U _ERROROn can the dead band be set so that only work as U _ERRORSpeed stability just is provided during greater than predetermined value.Such dead band will provide neutral speed stability for the certain speed range that comprises trim speed.

5. the discussion of compensation protection processor 34

With reference to what Fig. 1 mentioned, the present invention includes a compensation protection processor (34 among Fig. 1), compensation guard signal C is provided ^*U _{C﹠amp; P}It with calculating C ^*U signal C ^*U _ComputdInstruct C with the pilot ^*U pitch demand C ^*U _PilotcmdBetween subtractive combination produce C together ^*U error signal EC ^*U.

With reference to Fig. 5 a, the compensation protection processor 34 in the currently preferred embodiments of the invention comprises that provides a C ^*U owes fast guard signal (C ^*U _US) stall protection processor 416; One provides C ^*U evens up compensating signal (C ^*U _FL) even up compensation processor 414; One provides C ^*U overspeed protection signal (C ^*U _OS) overspeed protection processor 412.By stall protection processor 416, the signal of evening up compensation processor 414 and overspeed protection 412 and providing offers the input end of composite set respectively by signal routes 422,420 and 418.According to described arrangement, the output signal of composite set 424 offers switch 428, and switch 428 can be connected by the aircraft autopilot connection signal, also can disconnect, and depends on the system design selection.

5a. stall protection processor 416

Stall protection processor 416 functions are as follows: (1) has set up the lower limit of the reference velocity that can be set by the aviator; (2) increased system speed stability feedback owing when operation speed, and the bigger revision directive and the pilot is recognized owe speed state of bowing is provided.

The stall protection processor 416 of present embodiment has two kinds of operation mode shown in Fig. 5 b.First kind of operation mode is only to provide in the starting stage of taking off operation to owe speed protection signal C ^*U _USAnd C ^*The lower limit of U reference velocity, V _{REF (MIN)}, the mode of taking off.Specifically, at initial take-off stage, C ^*U owes speed protection signal and C ^*U reference velocity minimum value all is used to provide the pitching Performance Characteristics that allows the pilot to take appropriate action when power failure.Second of the stall protection processor operation mode is the back operation mode of taking off shown in Fig. 5 b, and this moment, processor provided C ^*U owes fast guard signal and C ^*U reference velocity minimum value adapts to the operation of aircraft in the normal flight zone.In two kinds of operation mode, owe fast guard signal value and C ^*U reference velocity minimum value all is to set according to the low speed yellow region thresholding that is commonly called " yellow area summit ".As known in this technology, yellow area summit correspondence that aircraft can carry out that the angle of attack does not reach " bar shakes " or the speed the during banked turns of the degree of 40 when beginning to buffet, just, do not reach the aircraft crew member and be warned the state of flight that stall soon begins to buffet.

Mode and taking off can automatically switch between the mode of back before being configured in of Fig. 5 b taken off, and the switching of coherent signal is by shown in the ordinary tap S1-A, the S1-B that are expressed as reducing trim restriction mode, the S1-C.In embodiments of the present invention, the take off switching of back mode operation only just takes place when three kinds of conditions satisfy simultaneously.Specifically, switching occurs in: (1) is if pass by one period schedule time (for example 15 seconds) from flying up; And (2) aircraft flight speed is greater than the scheduled volume in yellow area summit (for example 4 joints); And (3) C that sets by the cockpit crew ^*U reference velocity C ^*U _REFExceed scheduled volume in yellow area summit (for example 4 joints).

(mode of taking off) C when switch S 1-A, S1-B, S1-C are in shown in Fig. 5 b the position ^*The lower limit V of U reference velocity _{REF (MIN)}Be based upon a value that is lower than the yellow area summit.Specifically, in Fig. 5 b, represent the signal V of the bar that shakes _SSOffer terminal 434 by aircraft electronic prealarming system or other similar source.At frame 436 places, V _SSBy constant K V _S1Regulate and provide a signal to give the end of switch S 1-A than the low aequum (for example half of yellow region velocity amplitude) in yellow area summit.Signal filtered after by switch S 1-A (by speed restrictor or the back wave filter 438 that stagnates) is with the elimination bar rate signal V that shakes _SSAny sudden change, and the used speed limit of currently preferred embodiment of the present invention was about for 4 joint/seconds.System is in when reducing trim restriction mode, is converted into C by speed restrictor (or the stagnant wave filter in back) 438 signals that provide 435 by switch S 1-B ^*The minimum reference velocity V of U _{REF (MIN)}Can notice like this, in the initial period of taking off, minimum C ^*The U reference velocity is lower than the yellow area summit, and this specific character allows the pilot to carry out trim when being lower than the yellow region summit, just, sets a C who is lower than the yellow region summit ^*The U reference velocity.It is essential that this characteristic only is only when the approximate velocity of the engine that is not in good state drops within the yellow region.

Continue with reference to figure 5b; be at described stall protection processor and reduce trim restriction mode operating period; the signal 435 that is provided by speed restrictor 438 forms stall protection simultaneously and owes speed reference signal 439, and stall protection is owed the subtraction input end of speed reference signal by switch S 1-C and composite set 440 and linked.Composite set 440 deducts stall protection and owes speed reference signal 439 from the signal of representing the filtered flying speed of aircraft, and represents the signal of the filtered flying speed of aircraft to be provided for the terminal 442 among Fig. 5 b and link to each other with the addition input of composite set 440.The output signal of composite set 440 has just been represented filtered flying speed signal and has been the difference between the stall protection set speed signal of owing speed setting like this.This rate signal that is provided by composite set 440 is by the gain factor KV at frame 444 places _S3Regulate and offer the addition input of composite set 446.

When aircraft flight speed be lower than owe speed reference signal just when aircraft flight speed be lower than selected when owing speed reference signal, gain factor KV _S3Selected to obtain the mobile of requirement or to keep the variation of jociey stick in the required power at midline position (the card off-position is put) rear portion.For example, with reference to the description of Fig. 4 e, speed stability feedback characteristics of the present invention has been set up predetermined control force gradient, and the control force gradient is below or above reference velocity (V in order to keep _REF) flying speed need the pilot to apply the instruction of extra jociey stick.In currently preferred embodiment of the present invention, to be higher or lower than C ^*Desired control force gradient was 3 a pounds/joint when speed of U reference velocity was flown.In these embodiments, the described gain factor KV that owes the speed protection configuration _S3Increase the power gradient of 12 a pounds additional/joint, selected it can work (according to the configuration of Fig. 5 b when owing velocity amplitude when air speed is lower than, the mode of taking off is approximately 0.5 times of yellow region summit when operating, being in when taking off back mode operation is the yellow region summit).The increase of control force is in order to remind the pilot to note owing speed state, not determine consciously with the low-speed control aircraft but can not hinder.

Come back to the configuration of Fig. 5 b, composite set 446 provides the C as the function of aircraft pitch angular velocity turning compensation ^*U owes rate signal C ^*U _USDamping.In the configuration of Fig. 5 b, represent the signal of rate of pitch turning compensation to offer terminal 448; Through a suitable gain factor KV _S2(shown in the frame 450) regulated; And offer the subtraction input end of composite set 446.The signal that is provided by composite set 446 is limited the trim restriction mode C that device 452 is handled and conduct reduces then ^*U owes speed command signal C ^*U _USOutput.The last restriction of limiter 452 preferably is taken as zero.That is, in a preferred embodiment of the invention, no matter when, when filtered flying speed is owed speed reference greater than stall protection, just do not provide C ^*U owes rate signal (promptly equalling zero).Like this, when air speed is higher than predetermined value (be in 0.5 times that is approximately the yellow region summit when taking off the mode operation concerning the embodiment of the invention, being in when taking off back mode operation is summit, amber district), 416 couples of C of stall protection processor ^*The just not effect of U protection compensating signal.

The lower limit of limiter 452 has been determined can be by the maximal value of owing the pitching signal of bowing that the speed protection processor provides.According to the present invention, the setting of higher limit should prevent to produce and can not owe speed protection downward pitch demand signal for what the pilot firmly operated the joystick override.

As aforementioned, when no longer needing to allow the pilot when being lower than the yellow region summit, to adhere to speed trim (setting C ^*U reference velocity V _REF) time, the stall protection processor among Fig. 5 b can change its operation mode automatically in take-off process.Switch and broken away from when reducing trim restriction mode (that is, turning to the back mode of taking off) when the configuration of Fig. 5 b, represent the signal on summit, amber district to offer the input end of speed restrictor 438 by terminal 437 and switch S 1-A.Reduce the used bar rate signal that shakes in the trim restriction mode as system, representing the signal on yellow region summit from electronic prealarming system or similar source, to obtain.And, represent the signal on yellow region summit from the bar speed of shaking, to obtain, vice versa does not have or does not almost have systemic loss of energy.About this point, the input signal of switch S 1-A or based on the yellow region summit, perhaps based on the bar device speed of shaking, perhaps based on them the two (compensating) through load factor.

Can notice, when the configuration of Fig. 5 b is in when taking off back mode, offer the owing speed reference signal 439 and the signal 435 that provides by speed restrictor 438 is provided of subtraction input end of composite set 440.In other words, when taking off back mode operation, the signal 435 that provides by speed restrictor 438 with owe velocity feedback reference factor (shown in the frame 456) and multiply each other by (shown in the frame 454), and offer the subtraction input end of composite set 440 by switch S 1-C by the output signal that multiplier 454 provides.Owe the function that velocity feedback reference gain 456 is flying speed (such as Mach numbers), it should make when owing speed reference signal 439 constant relatively under the low relatively flying speed (for example equaling one) and in higher flying speed and reduce.

Adjusting is by the jociey stick steering force signal of the determined increase of aircraft sensory system among Fig. 2 a when increasing for partial offset speed as the purpose of owing the velocity feedback reference gain of the function of flying speed.Specific purpose is in order to obtain the bar angle of attack (α that shakes _Ss) the jociey stick steering force that requires, all keeps unanimity relatively under all flight conditions (required scope is 15～25 pounds) of the bar angle of attack in currently preferred embodiments of the invention that shake.

Described with reference to the mode operation of taking off, composite set 440 has produced one and has represented the filtered flying speed of aircraft and owe the signal of difference between the speed reference signal 439 at 442 ends.The velocity contrast signal is regulated by the KVS3 at frame 444 places and is obtained to be lower than the instruction driving stick force gradient that need increase when owing speed reference in speed; Rate of pitch turning compensation damping realizes at composite set 446 places; The signal that causes is restricted (being limited device 452), and as C ^*U owes the speed protection signal and is provided, if there is not pilot override, and C ^*U owes the speed protection system will cause the pitch demand signal of bowing.

When the speed protection configuration of owing of Fig. 5 b is in when taking off back mode operation minimum C ^*U reference speed signal V _{REF (MIN)}Directly do not provide from the output signal of speed restrictor 438.On the contrary, shown in Fig. 5 b, the signal 435 that is provided by speed restrictor 438 is multiplied by (at multiplier 458 places) Mach number speed trim and suppresses gain (shown in 460), and the signal that obtains becomes by switch S 1-B and is minimum reference velocity V _{REF (MIN)}Just as the foregoing velocity feedback reference gain of owing, the trim of Mach number speed suppresses gain and regulates as the aircraft flight function of speed; Show a constant gain (such as) substantially in low-speed range, this constant gain reduces when speed is higher than predetermined Mach number (such as Mach 2 ship 0.6).The minimum C that determines when like this, taking off back mode operation by being configured in of Fig. 5 b ^*The U reference velocity is approximately equal to the yellow region summit when air speed is lower than predetermined Mach number, and can reduce when air speed is higher Mach number.The establishment of this relation is to owe the change in gain of velocity feedback gain to obtain α under all flying conditions substantially in order to replenish _SsThe steering force of consistent instruction driving bar relatively that requires.

5c. even up compensation processor 414

Fig. 5 c schematically compensation of evening up of illustration currently preferred embodiments of the invention disposes.In the configuration of Fig. 5 c, represent the signal of undercarriage terrain clearance to offer terminal 460 or other device that links to each other with look-up table 462, this device produces the flare command signal 461 (pitching downwards) that a change with the undercarriage height (increase) changes.Being used for the various signals of aircraft takeoffs and landings height indication are provided is known in technology, and they normally obtain from the aircraft radio height indicator.Represent the signal of undercarriage height preferably to pass through filtering and handle to revise the aircraft pitch attitude.

Look-up table 462 has been set up a kind of relation between the flare command signal of undercarriage height and output, the flare command signal of output in fact when aircraft landing emulation the ground influence that runs into of aircraft.Influence emulation by the ground of evening up compensation processor and providing and make C ^*The U system provides similar those not increase the flare control characteristic of steady aircraft.

Continuation is with reference to Fig. 5 c, and the compensating signal 461 of evening up that is provided by look-up table 462 offers limiter 467 by switch 466, as long as effective undercarriage altitude signal is just offering and evening up compensation processor, switch 466 just offers limiter 467 evening up compensating signal.Limiter is limited to zero (to determine that guarantor's new line compensating signal is not evened up compensation processor and provided) for 467 times, and the upper limit (being 0.54g in the current embodiment of the present invention) is determined flare command required when equaling to contact to earth.Pass through switch 468 as C by the output signal that limiter 467 provides ^*U evens up compensating signal C ^*U _FLOutput.

Only when aircraft was near mode and has reached the undercarriage height that can begin flare control, switch 468 just was switched on and provides the C that equates with the signal that is provided by limiter 467 ^*U evens up compensating signal C ^*U _FLIn currently preferred embodiment of the present invention, connect switch 468 used logics require be: (1) aircraft flown at least 60 seconds (in order when taking off, to prevent to even up compensation); (2) aircraft flap launches; (3) the undercarriage altitude signal has at least 1 second demonstration undercarriage height to be lower than 50 feet; (4) the undercarriage altitude signal is effective (just just being provided for the terminal 460 of evening up in the compensation processor).(perhaps, in above-mentioned condition, increasing the downward switch of undercarriage).If above-mentioned condition has any one not satisfy, switch 468 can not connected, and shown in frame 470, C ^*U evens up compensating signal C ^*U _FLBe set to zero (not compensation).Yet in case switch 468 is connected, even lose undercarriage altitude signal or the inefficacy of undercarriage altitude signal, it will still keep on-state.As described below, when evening up compensation when having generated, even the undercarriage altitude signal lost efficacy, this method still allows to even up compensation processor and operates in the mode that can not cause evening up the compensating signal sudden change.

The configuration of Fig. 5 c also comprises a speed restrictor 472, and when evening up compensation processor when a compensating signal just is being provided, even lost efficacy in the radio altimeter of undercarriage altitude signal or other source, speed limit 472 also can be controlled the operation of evening up compensation processor.In this, if one substitutes and to even up the compensation connection signal and offer and even up compensation processor (shown in Fig. 5 c center 474), speed restrictor 472 substitutes the source as one that evens up compensating signal, and it offers limiter 467 by the output signal that speed restrictor 472 provides by switch 466.Substitute and even up the signal that compensation flop signal (dotted line 475) can be an indication radio altimeter Signal Fail, or indication undercarriage altitude signal is considered to more insecure other available signal.

In normal operation (just switch 466 is in shown position), the C that provides by limiter 467 ^*U evens up compensating signal and links to each other with speed restrictor 472 and set up an initial value or deviation grade.Connected switch 466 if substitute the trigger pip 475 of evening up bucking-out system, the signal currency that is provided by limiter 472 just is established as the input signal (original state) of speed restrictor 467, if aircraft is not carried out the process of going around, the signal 476 of evening up compensating signal when representing ground connection so just offers limiter 472 by switch 474.Because the output signal of limiter 472 at that time will be as the function of time and linear increasing up to reaching to greatest extent (being 0.54gs in the present embodiment), thereby in fact the flare command signal that normally provided by look-up table 462 (be the function of time during operation, rather than the function of undercarriage height) has been provided speed restrictor 472.

Just begin if went around before the loss of undercarriage altitude signal, switch 474 is switched on and establishes following be limited to zero (shown in the frame 478) of speed restrictor 472.Be limited to zero and mean if the undercarriage altitude signal lost efficacy and just providing when evening up compensating signal at establish speed restrictor 472 when handling following that go around, it is zero C that speed restrictor 472 will provide a linear decrease ^*U evens up compensating signal.

The Technology professional will recognize that, also can produce satisfied C without other configuration of evening up compensation processor shown in Fig. 5 c ^*U evens up compensating signal.For example, the look-up table 462 among Fig. 5 c can be replaced by a loop or other device, and this circuit or other device are triggered and produce a flare command signal as time (rather than undercarriage height) function when aircraft altitude is lower than 50 feet.Secondly, if normally offer when the signal of limiter 467 is interrupted or validity is a problem, speed restrictor 472 can be able to be made that the signal smoothing ground that offers limiter 467 reduces to suitable limits value other install replace.

5d. overspeed protection processor 412

The control law (412 among Fig. 5 a) that Fig. 5 d illustration is realized by current preferred overspeed protection processor.The configuration of Fig. 5 d has produced a C at last ^*The overspeed protection signal C of U ^*U _OSUnless under the override state, otherwise no matter which kind of hypervelocity condition aircraft is under, this device all can produce a new line pitch demand signal.As C ^*The result that the U overspeed signal produces, pilot must keep on the instruction driving bar and be higher than usual power forward to keep or the increase hypervelocity.According to the present invention, set up C ^*The purpose of U hypervelocity pitch demand signal value is, can remind the driver that aircraft is in overspeed condition fully for keeping the required instruction driving stick force forward of overspeed condition, and if desired, the driver can select aircraft to be in the override state consciously.

Since the first overspeed protection signal 507 based on the calibration speed of aircraft the second overspeed protection signal 509 based on Mach speed, thereby basically, definite existence and the degree that exceeds the speed limit of Fig. 5 d device.Then two overspeed signals are compared, determine wherein greatlyyer one, and it is limited, so that prevent the generation of nose-down attitude command signal and it is limited in an appropriate value required for the compensation roll attitude.The signal that is restricted is again as C ^*U overspeed protection signal C ^*UOS is transfused to the adder among Fig. 5 a 424.

More particularly, according to Fig. 5 d, represent the signal of flying speed after the aircraft filtering to be added to the addition end of composite set 500 by terminal 502.What link to each other with the subtraction input end of composite set 500 is to operate flying speed (V with the representative of square frame 501 expressions a little more than the aircraft maximum _MO) the signal of flying speed.In currently preferred embodiment of the present invention, the signal that is input to the subtraction input end of composite set 500 compares V _MOHigh 6 joints.Flying speed and the V that has increased sign nargin after the aircraft filtering have been represented in the output of composite set 500 _MODifference between the speed, this output are used for controlling with extreme V with respect to the value of the overspeed signal of representing extreme maximum Mach speed after the constant-gain coefficient adjustment that scaler 504 is produced _MOValue for any synthetic overspeed signal on basis.

The flying speed difference signal is input to the addition end of composite set 506 after the conversion that provides in the frame 504.The signal of second addition end of input is represented the filtered flying speed speed of aircraft (time speed that flying speed changes after the filtering just), and this signal is added to terminal 508 and passes through the constant-gain coefficient conversion by scaler 510.The gain coefficient of scaler 510 is to set like this, make when aircraft near or surpass V _MOAnd flying speed suitably increases the overspeed protection that is produced by adder 506 and calculates the flying speed instruction when increasing.Also promptly, the gain coefficient of scaler 510 is to set up according to the gain coefficient of scaler 504, in order to the signal (being provided by composite set 500) of control representative flight hypervelocity with as the Relative Contribution of the signal of the representative flying speed speed of acceleration item.

In the configuration of Fig. 5 d, overspeed protection Mach command signal 509 is provided by composite set 512, and its producing method is similar to overspeed protection and calculates flying speed command signal 507.Specifically, represent the signal of aircraft Mach speed after the filtering to be input to the addition end (by terminal 514) of adder 520, the signal that the subtraction input end of composite set 520 receives is the maximum flight Mach of the aircraft speed (MMO) (shown in frame 516) after representative has added desirable nargin (as 0.01).The signal of composite set 520 outputs is transported to the addition input end of composite set 512 by the back (by frame 518) of converting.What be connected to composite set 512 another addition input ends is the signal (feeding terminal 524) of representing Mach speed after frame 522 is multiplied by suitable coefficient conversion.

Among Fig. 5 d, comparer 526 is compared the overspeed protection calibration flight speed command signal 507 that adder 506 produces with the overspeed protection Mach command signal 509 that composite set 512 is supplied with.The big limiter 528 that is imported among both.Shown in frame 530, the minimum value of limiter 528 is zero.But the maximal value of limiter 528 is not a constant, but decides according to the aircraft roll attitude.

More particularly, represent the signal of aircraft roll attitude (unit is degree) to be passed to terminal 532.Signal value (absolute value) is determined composite signal again through converting through frame 534, and the conversion of device shows the gain-roll attitude relation of the type at frame 536 places.Particularly, when the absolute value of aircraft roll attitude in 0～30 ℃ ° and when 30 ° to 60 ° scope internal linear of roll angle were decremented to 0, the gain coefficient of scaler 536 equaled one substantially.The output signal of scaler 536 is through an adjusting (at frame 538 places) of having set up the constant of maximum instruction driving stick force gradient, and this gradient can be generated by illustrated overspeed protection processor.Just, the used reduction coefficient of frame 538 to shown in the C that produced of configuration ^*U overspeed protection signal value has been set a upper limit, forms a upwards maximum of pitch attitude instruction, if select the aircraft overspeed condition for use, but pilot's override should amount.

Set up limiter 528 following to be limited to zero be not produce the nose-down attitude command signal in order to ensure the overspeed protection processor, otherwise should air speed further be increased instruction.The KB limit of regulating limiter 528 in a manner described can reduce the size of aircraft overspeed protection command signal when spiraling at a high speed, and the absolute value of roll attitude is between 30 °～60 ° in the case.If the roll attitude of aircraft meets or exceeds 60 °, the output signal of scaler 536 becomes zero so, thereby the KB limit of limiter 528 is based upon zero place.Therefore, when wide-angle tilt, there is not the overspeed protection signal to produce.Under the hypervelocity condition when aircraft flies in typical lift-over scope, regulate the back by 528 signal supplied of limiter by frame 540 and generate a suitably C of height ^*U overspeed protection signal C ^*U _OS(for example, the C of every pound of unit of representative that limiter 528 is produced ^*U signal is transformed into the C of the every g unit of representative ^*The signal of U).

Can notice, more than Shuo Ming C ^*The U overspeed protection system makes the crew increase alertness to hypervelocity to increase in order the to keep the aircraft hypervelocity flight essential strength of jociey stick forward mode.Also have, as protection previously discussed and compensation processor, the relation between instruction driving stick force and the signal has been set up in described configuration, protection and compensating signal that the pressure that makes the pilot can keep the instruction driving bar is fully generated to overcome the present invention.About this point, in currently preferred embodiments of the present invention, aircraft is wanted the override overspeed protection near design maximum speed with near the flying speed of design maximum Mach speed the time, and maximum stick force is approximately about 40 pounds.

6. the discussion of rate of pitch damping processor 36

The effect of rate of pitch damping processor 36 is to produce rate of pitch damping instruction Q _COOM, when this instruction is added to C ^*U integration instruction IC ^*U _COMWhen last, it will decay rudder face instruction short term response and regulate the short-term frequency.Rate of pitch damping processor 36 preferred for this invention has two basic characteristics: one, and it has a cut-out multiplier 631; Its two, it has a composite set 630, this device is used for combination back stagnate rate of pitch feedback signal 628 and ratio rate of pitch feedback signal 626.

When aircraft is parked in ground, cuts off multiplier 631 and cut off angle of pitch damping increment.Cut off the value that multiplier 631 provides according to calculation element 638, instruct Q with value in 0.0 to 1.0 scope and rate of pitch damping one _COMMultiply each other.Calculation element 638 determines that from sky/ground state the processor (not shown) receives a ground signal.This preferred processor is disclosed among the patent No 08/441,282 of the U.S., and the invention people is EVans etc., and the title of this invention is " system that sky/earth signal is provided for aircraft flight control system ", and authorizing day is on October 27th, 1998, and with reference to being included in this.Begin and carry out the transition to aerial state from the state of ground, ground signal will switch to " very " by " vacation ".This makes calculation element 638 is that zero to become output valve be 1 by output valve.Yet preferred calculation element 638 in a quite short time (nominally less than 10 seconds) makes output valve be skewed rather than step-like.Opposite condition then makes the output signal of calculation element 638 become zero with skewed by 1.Make rate of pitch damping instruction Q by the inclination output valve _COMReduce gradually.This makes and becomes level and smooth and not significantly sudden change from stability augmentation system to the transformation of non-stability augmentation system.

The stagnant rate of pitch feedback signal 628 in back links to each other with the addition input end of composite set 630 respectively with ratio rate of pitch feedback signal 626.The output signal of composite set 630 is instructed Q corresponding to the rate of pitch damping before by 631 multiplications of above-mentioned cut-out multiplier _COMRate of pitch damping instruction Q _COMForm by signal 626 and signal 628 merging.

The stagnant wave filter 618 in back generates the stagnant rate of pitch feedback signal 628 in back, and the stagnant wave filter 618 in back has K _Q1The changing function of/(τ S+1) form, wherein K _Q1With τ from chart shown in the frame 614.The input of wave filter 618 is the θ of compensation rate of pitch signal q or Fig. 4 d that turn _ESTK _Q1Preferably decide K when wing flap is packed up with the value of τ according to flap configuration or flying speed _Q1=1.0 and τ=1.5, K when wing flap puts down _Q1=1.5 and τ=1.0.This numerical value is packed up and the transformation of wing flap between putting down is (nominally 10～30 seconds) that gradually change in the given time at wing flap.The effect that is provided with of the stagnant wave filter 618 in back is stagnant rate of pitch feedback signals 628 after generating on the basis of compensation rate of pitch signal q of turning, and wherein the effect of q is to adjust the short-term frequency as feedback.

Ratio rate of pitch feedback signal 626 is generated by wave-shaping filter 624.Wave-shaping filter 624 has the changing function of the frequency content of removing the rate of pitch signal that can hinder the aircraft state of nature.The input signal of wave-shaping filter 624 is to take advantage of gain ratio of damping K by what gain 620 was revised _qTurning compensation rate of pitch signal q or θ _ESTK wherein _qProvide by illustrated box 619.As shown in Figure 6, K _qItem is preferably decided K when wing flap lifts according to flap configuration _q=1.0, K when wing flap puts down _q=1.35.Press the order of wing flap device when high speed and low-speed operations, gain ratio of damping K _qProduced sufficient short-term damping.Once more, the transformation between the yield value that wing flap is packed up and wing flap puts down is quite progressive.In the preferred embodiment, the wing flap device is used for determining K _Q1And K _qValue, wing flap down state are at first determined by the flying speed data.Use the wing flap device to determine K _Q1And K _qThe purpose of value is to change along with the flying quality of mistake in order to prevent to gain, and this mistake can occur sometimes when running into volcanic ash cloud, and uses flying speed can prevent that when the wing flap error in data gain from increasing.Trigger mechanism except that flap configuration also can be used for determining K _Q1And K _q, this depend on available data and deviser's hobby (as, use other logical combination of flap configuration and/or flying speed).

Basically, the compensation rate of pitch signal q that turns is input to rate of pitch damping processor 36, and with the gain ratio of damping K that depends on the wing flap device _qMultiply each other, through being added to after shaping filter 624 shapings on the stagnant rate of pitch feedback signal 628 in back, this feedback signal is provided by the back stagnant wave filter 618 that links to each other with reception turning compensation rate of pitch signal q the signal that draws again.

Can notice that rate of pitch damping processor 36 can be based on pitch attitude signal θ, replacing turning compensates rate of pitch signal q, or combines with the latter.This variation require gain 620 and 618 and wave filter 624 make suitable modification.Preferred input signal is that top the 4th part had been discussed and the turning of the ground reference shown in Fig. 4 b compensation rate of pitch signal q.

7.C ^*U integrator 28

C ^*28 couples of error signal E of U integrator _C*UIntegration and error signal E _C*UA part be added to elevating rudder pitch demand δ continuously _E.FILTOn.Finally, C ^*U integrator 28 will be eliminated error signal E _C*U, this is because elevating rudder pitch demand δ _{E, FILT}Elevating rudder moved, and this can make to be input to and calculate C ^*U signal C ^*U _ComputedSignal change thereby the C that the pilot is required ^*U instructs C ^*U _PilotemdWith calculating C ^*U signal C ^*U _ComputedWith compensation protection C ^*U signal C ^*Difference between the U δ sum is zero.

Currently used C ^*The preferred form of U integrator 28 has six devices: gain K _i, cut off table musical instruments used in a Buddhist or Taoist mass 642, anti-rolling-in device 648, decline gain apparatus 666, composite set 654 and integrator 662.Every kind of device all can carry out needed improvement to pitch-axis stability augmentation and instruction control augmentation system, though, have only integrator 662 device that is absolutely necessary.C ^*The essential structure of U integrator 28 is: error signal E _C*UWith the addition input signal of trim negative acknowledge character (NAK) TNS as composite set 654, the output signal of composite set 654 offers integrator 662 again.Resulting signal is the integration C that discusses in the part 1 ^*U instructs IC ^*U _COMAt error signal E _C*UBefore the input composite set 654, may exist some conditions to make the other value enter composite set 654 rather than error signal E _C*UBelow these conditions will be discussed.

According to chart shown in the frame 631, error signal is by integration gain factor K _i632 places convert in gain.By K _iRegulating error signal E _C*UFor at a high speed and low-speed conditions good short term response (flap configuration is indicated shown in frame 631) is provided.Frame 631 is set K during wing flap on receiving _i=5, when putting down wing flap, set K _i=8, putting down wing flap and receiving that value can asymptotic transition between the wing flap.If needed.Other device such as flying quality can be used for determining at a high speed or low-speed conditions.Gain K at 632 places _iError signal E after the variation _C*UBe shown among Fig. 7 as signal 634.

C ^*Second device of U integrator 28 is to cut off multiplier 642 (or gain rejector), when aircraft on the ground the time, cuts off multiplier 642 and cuts off C ^*U increases surely, cut off multiplier 642 the output signal of gain 632 with and multiply each other by the numerical value in 0.0～1.0 scope that calculation element 638 provides.Calculation element 638 is changed the used numerical value that changes of multiplier 642 in the mode that tilts basically between 0.0～1.0.Calculation element 638 illustrated in the 6th part in the above.C ^*U increases the steady reduction gradually along with the inclination of output valve.This makes that be a level and smooth inapparent change procedure by stability augmentation system to the filtration of non-stability augmentation system.

C ^*The 3rd device of U integrator 28 is anti-rolling-in logical units 646, and when elevating rudder (or aircraft) can not or not need to respond, anti-rolling-in logical unit 646 can prevent that integrator from increasing elevating rudder pitch demand δ _{E, FILT}Increase signal.At C ^*Be provided with change-over switch 648 in the U integrator 28, when it is in normal closed position, can not disturb the output signal of multiplier 642.Switch generally cuts out, unless run into a series of certain criteria.Determine that the logic when switch 648 is opened provides as shown in Figure 7 and by device 646.When satisfying one of following condition, the preferred logic of device 646 is opened switch, that is: (1) aircraft on the ground; (2) aircraft tail is near ground and error signal E _C*URequire bigger new line; (3) aircraft is in big angle of attack state and error signal E _C*URequire bigger new line; (4) elevating rudder reach range of application maximal value and error signal E _C*URequire elevating rudder on the restriction direction, bigger displacement to be arranged.In these conditions, under any one condition, had better not allow integrator to carry out work, because the requirement of bigger pitching is not produced or do not allow to produce the bigger change of pitch attitude.Open the addition input end that switch 648 causes 0.0 value is transported to composite set 662, replace the output of multiplier 642.

The 4th device is composite set 654, is used to make up error signal E front part 1 discussion and as shown in Figure 1 _C*UThe addition input signal of (or switch 648 is zero when opening) and trim negative acknowledge character (NAK) TNS.Used trim negative acknowledge character (NAK) can obtain from stabilizator position rate (being converted into per second elevating rudder angle), and this method is in U.S. Patent No. 08/441,682In obtain the explanation, the invention people be E.E.Coleman etc., exercise question is " method and apparatus of automatic trim aircraft stabilizator ", is issued on Dec 2nd, 1997, and the reference be included in this.The trim negative acknowledge character (NAK) can produce in the operation by present available autopilot system.During unloading operation, trim negative acknowledge character (NAK) TNS is added on the output signal of multiplier 642, elevating rudder moved towards the appropriate location about stabilizator, and can not produce significant aircraft response.

C ^*The 5th device of U integrator 28 is decline gain apparatus 666, and it comprises a decay gain coefficient k _dAnd the output signal that makes integrator 662 in predetermined period decays to zero gradually.When aircraft was on the ground, decline gain apparatus 666 was eliminated gradually and is increased surely.This is a kind of good characteristic because next time flight before integration C ^*U command signal IC ^*U _COMShould be set to zero.Be preferably in the short period time and after aircraft lands the decay of back short time and just finish this decay.Among Fig. 7, decline gain apparatus 666 is provided for integration C ^*U instructs IC ^*U _COMSignal, and this signal and decay gain k _dMultiply each other.When aircraft was on the ground and the output signal of decline gain apparatus 666 offered composite set 654, the switch 652 that is positioned on the output signal route of decline gain apparatus 666 can be opened.Preferred decay gain k _dLess than q.

The 6th device is integrator 662.Integrator 662 has calculated error signal E _C*UA part and the integration C that has used of part 1 discussion is provided ^*U instructs IC ^*U _COMPut it briefly, offer C ^*The error signal E of U integrator 28 _C*UChange back and trim negative acknowledge character (NAK) TNS combination by storage gain K, then integration and produce C ^*U integration instruction IC ^*U _COM

Though illustrated and illustrated the preferred embodiments of the present invention, be appreciated that under the situation that does not depart from spirit of the present invention and scope, can carry out various changes.C ^*The U control law is novel, and develops for relaxing the static-stability aircraft.Yet be appreciated that other non-static-stability aircraft that relaxes also can benefit from the present invention.

Claims

1. an aircraft pitch increases steady and command control system, has the response of representative instruction elevating rudder, jociey stick position, normal acceleration, and the input signal of aircraft pitch angular velocity; This aircraft pitch that is used to change the aircraft pitch angle increases steady and command control system presents stable speed responsive characteristic; This aircraft pitch increases steady and command control system comprises:

(f) integrator is used to receive described difference signal and is used for difference signal is guided into null value, and integrator generates output signal, receives first signal combiner by second input end.

2. increase steady and command control system according to the described aircraft pitch of claim 1, wherein calculate C ^*The U processor also comprise first composite set and second composite set and:

(a) one first gain factor Kq is used to apply hybrid gain in the aircraft pitch angular velocity signal, then at described first composite set place combination rate of pitch signal and speed error signal; And

(b) one second gain factor Kv is used to apply speed stability and gains in speed error signal, then at described second composite set place group speed error signal and the rate of pitch signal.

3. increase steady and command control system according to the described aircraft pitch of claim 1, wherein vertical acceleration signal is the earth reference vertical acceleration signal, and the rate of pitch signal is an earth reference rate of pitch signal.

4. increase steady and command control system according to the described aircraft pitch of claim 1, wherein angle of pitch instruction processing unit comprises the bias voltage correction device, and being used to receive the jociey stick position signalling and be used to provide of connection proofreaied and correct the jociey stick position signalling to instruction C ^*The U processor is as the signal of representing the jociey stick position, and the bias voltage correction device is used for reducing the bias voltage of jociey stick position when jociey stick is in neutral position.

5. increase steady and command control system according to the described aircraft pitch of claim 4, wherein the bias voltage correction device comprises a device for limiting, is used for correction signal is defined in a certain numerical value within the preset range, and this scope is represented jociey stick position signalling numerical value.

6. increase steady and command control system according to the described aircraft pitch of claim 4, wherein the bias voltage correction device comprises:

(a) one the 3rd composite set is used to make up jociey stick position signalling and the signal of representing required correcting value;

(b) switch has to activate and conciliates moving state, and this switch is set in the state separated when in fact movable jociey stick is removed neutral position, and actuating state is set in the numerical value identical with the jociey stick position signalling to correction signal.

7. increase steady and command control system according to the described aircraft pitch of claim 6, wherein the bias voltage correction device also comprises a smoothing filter, is used to receive the also correction signal of output smoothing when switch is in its actuating state of jociey stick position signalling.

8. increase steady and command control system according to the described aircraft pitch of claim 1, wherein angle of pitch instruction processing unit also comprises a sensation order set, be used to provide the power gradient to output signal to the device that is used to regulate the required power of mobile jociey stick, described sensation order set is used to increase the required power of mobile jociey stick during near stalling speed or stall angle when aircraft.

9. increase steady and command control system according to the described aircraft pitch of claim 8, feel that wherein order set comprises:

(a) one first power gradient component signal has a predetermined value that is relevant at least a flying quality input signal;

(b) one the 4th composite set has first and second input ends, and first input is rate signal or angle of attack signal, and second input is with reference to the stalling speed signal or with reference to shaking bar angle of attack signal, this 4th composite set generates output error signal;

(c) scaler is used to revise output error signal, generating the second power gradient component signal, and

(d) one the 5th composite set is used to receive the every output and the power output gradient output signal of the first and second power gradient component signals.

10. increase steady and command control system according to the described aircraft pitch of claim 9, feel that wherein the order set scaler comprises:

(a) scaler is used to increase the numerical value of output error signal; And

(b) delimiter is used to limit the scope of acceptable output error signal numerical value.

11. increase steady and command control system according to the described aircraft pitch of claim 10, feel that wherein order set also comprises an angular velocity limiter, is used to eliminate the acute variation of the second power gradient component signal.

12. increase steady and command control system according to the described aircraft pitch of claim 1, wherein instruct C ^*The U processor comprises a wave-shaping filter, this wave-shaping filter has an input end and an output terminal, represent the signal of jociey stick position to be provided for input end, it is the instruction elevating rudder response signal with desirable characteristics that wave-shaping filter is used to change the signal of representing the jociey stick position.

13. increase steady and command control system according to the described aircraft pitch of claim 12, wherein instruct C ^*The U processor also comprises a scaler, is used for dispatch command elevating rudder response signal to cause the stick force of the harmonious every gram that spreads all over flight envelope.

14. increase steady and command control system according to the described aircraft pitch of claim 1, wherein calculate C ^*The U processor also comprises a speed stability processor, and it comprises:

(a) one first composite set is used to make up the signal of representing reference velocity and the signal of representing air speed, to generate the speed error signal of representing difference between them; And

(b) multiplying assembly is used to apply power and gains in speed error signal.

15. increase steady and command control system according to the described aircraft pitch of claim 14, wherein the speed stability processor also comprises one second composite set, is used to make up fluctuating damping feedback signal and the speed error signal that gains and revise.

16. increase steady and command control system according to the described aircraft pitch of claim 14, wherein the speed stability processor also comprises one second composite set, be used to make up a direct trim signal and the speed error signal that gains and revise, directly trim signal provides additional elevating rudder instruction, so that simulate the summary responses that each elevating rudder activates for balancing device.

17. increase steady and command control system according to the described aircraft pitch of claim 1; also comprise a compensation and protection processor; be used for when air speed is lower than predetermined value, generating and owe fast signal, owe the input end that fast signal is provided for the secondary signal combiner, as an error signal component.

18. increase steady and command control system according to the described aircraft pitch of claim 1; also comprise a compensation and protection processor; be used for when air speed just is operated, generating overspeed signal beyond the scheduled operation scope; overspeed signal is provided for an input end of secondary signal combiner, as an error signal component.

19. increase steady and command control system according to the described aircraft pitch of claim 1, also comprise:

(a) a rate of pitch damping processor is used for damping elevating rudder instruction short period response and regulates the short period frequency, and this rate of pitch damping processor generates a rate of pitch damping command signal;

(b) one the 3rd signal combiner, being used to of connection receives rate of pitch damping command signal and integrator output signal, is used to provide this combination to second input end for first signal combiner.

20. increase steady and command control system according to the described aircraft pitch of claim 1; also comprise an elevating rudder loop gain device; be used for revising the integrator output signal, so that protection is with respect to the stability of the aircraft evelvator response of each shake mode by an elevating rudder loop gain.

21. increase steady and command control system according to the described aircraft pitch of claim 1, also comprise a structural modal wave filter, be used to revise the output of first signal combiner, so that protection is with respect to the stability of the aircraft evelvator response of each structural modal of aircraft.

22. a method that is used to generate the elevating rudder command signal that presents the stabilized speed response characteristic comprises the steps:

(d) utilize output signal to revise the elevating rudder command signal.

23. one for utilizing C ^*Improving one's methods of the aircraft pitch angle control system of criterion comprises:

(a) detection speed stability signal is used to make aircraft to be returned to its trim speed, and the speed stability signal comprises:

(i) detect the signal of representing air speed and the signal of representing reference velocity; And

(ii) import with the formation speed error signal by input of one first composite set Combined toy plane speed and reference velocity, speed error signal is used to formation speed stability signal; And

(b) by one second composite set group speed stability signal and C ^*Criterion.

24. it is described for utilizing C according to claim 23 ^*Improving one's methods of the aircraft pitch angle control system of criterion also comprises by an integrator and guides speed error signal into null value.

25. a method that is used to generate the elevating rudder command signal with speed stability comprises following each step:

26. in accordance with the method for claim 25, also comprise the step that generates the stall protection signal, wherein combined command elevating rudder response signal comprises additional stall protection signal with the step of calculating the elevating rudder response signal.

27. in accordance with the method for claim 26, wherein the stall protection signal is the aircraft bar function of speed of shaking.

28. in accordance with the method for claim 26, wherein the stall protection signal is the shake function of the bar angle of attack of aircraft.

29. in accordance with the method for claim 25, also comprise the step that generates the overspeed protection signal; Wherein combined command elevating rudder response signal comprises additional overspeed protection signal with the step of calculating the elevating rudder response signal.

30. in accordance with the method for claim 25, comprise also generating the step of owing fast guard signal that wherein combined command elevating rudder response signal comprises the additional speed protection signal of owing with the step of calculating the elevating rudder response signal.

31. sensation order set, be used to have among the aircraft of jociey stick, the sensation order set is used to provide the power gradient to output signal to the device that is used to regulate the required power of mobile jociey stick, the sensation order set is used for increasing the required power of mobile jociey stick during near stalling speed or stall angle when aircraft, feels that order set comprises:

(c) scaler is used for the round-off error output signal to form the second power gradient component signal; And

32. according to the described sensation order set of claim 31, wherein scaler comprises:

(a) scaler is used to increase the numerical value of error output signal; And

(b) device for limiting is used to limit the scope of acceptable error output signal numerical value.

33. according to the described sensation order set of claim 31, wherein correcting device also comprises the angular velocity limiter, is used to eliminate the acute variation among the second power gradient component signal.

34. bias voltage correction system, be used among the aircraft with jociey stick and angle of pitch control system, the position of jociey stick is by the representative of jociey stick position signalling, and the bias voltage correction system is used for suppressing any jociey stick bias voltage when jociey stick is in neutral position, and the bias voltage correction system comprises:

(b) switch has to activate and conciliates moving state, and switch is set to the state separated when in fact jociey stick removes its neutral position, and the state of separating allows correction signal to be retained in its last numerical value place, and actuating state allows correction signal to equal the jociey stick position signalling.

35. according to the described bias voltage correction of claim 34 system, also comprise smoothing filter, it is traditional delay filter, is used to receive the also correction signal of output smoothing when switch is in actuating state of jociey stick position signalling.

36. according to the described bias voltage correction of claim 35 system, also comprise device for limiting, be used for correction signal is defined in some jociey stick position signalling numerical value within the preset range that limits.