CN104260889A - Hanging frame for helicopter to deliver aircraft at low speed and aircraft attitude control method - Google Patents
Hanging frame for helicopter to deliver aircraft at low speed and aircraft attitude control method Download PDFInfo
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- CN104260889A CN104260889A CN201410438725.9A CN201410438725A CN104260889A CN 104260889 A CN104260889 A CN 104260889A CN 201410438725 A CN201410438725 A CN 201410438725A CN 104260889 A CN104260889 A CN 104260889A
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Abstract
The invention discloses a hanging frame for a helicopter to deliver an aircraft at the low speed and an aircraft attitude control method. The hanging frame comprises a main steel cable, hanging frame steel cables, damping plates, a hanging frame main structure, a stable umbrella, compressing components and separation release components. One end of the main steel cable is connected with the helicopter, and the other end of the main steel cable is connected with the hanging frame main structure through the hanging frame steel cables; the damping plates are used for reducing left-and-right swing of the hanging frame; the stable umbrella is used for removing front-and-back swing of the hanging frame and rotating, around the main steel cable, of the hanging frame; each compressing component comprises a compressing switch and an adjustable compressing foot, and used for separating signal detection; and the aircraft is hung below the hanging frame main structure through the separation release components. The aircraft attitude control method includes the following steps that the aircraft performs separation signal detection based on a reliable separation signal detection circuit formed by the compressing switches; before separation, open-loop control over a longitudinal channel and a heading channel of the aircraft is carried out, and close-loop control over a roll angle of a transverse channel is carried out; after separation, close-loop control over the pitch angle and normal acceleration of the longitudinal channel of the aircraft is carried out, and close-loop control over the roll angle of the transverse channel is carried out; and close-loop control over lateral acceleration of the heading channel is carried out.
Description
Technical field
The present invention relates to and be a kind ofly applicable to hanger and the Spacecraft Attitude Control that helicopter low speed throws in unpowered lift formula aircraft, belong to aircraft manufacturing technology design field.
Background technology
Track reenters or sub-track reenters lift formula aircraft dependence aerodynamic force and gravity can realize horizontal landing, not necessarily need the main power system used or configuration is complicated, such as space shuttle, X-37B, spaceship No. two (SpaceShipTwo), The Commitments (Dream Chaser) etc.Owing to not using or not configuring main power system, this kind of aircraft need carry out extension by other carrying platform and flies captive test and throw in flight test in development process, such as transport plane or helicopter, to verify aircraft low speed aerodynamic characteristic and approach Guidance and control technology.The space shuttle of earlier US adopts the back of the body of the Boeing-747 transport plane after the reequiping mode of carrying on the back to carry out approach flight test, X-37B successively adopts UH-60 Black Hawk helicopter in early stage Project R&D process, CH-47 props up and exerts dry helicopter and white knight (White Knight) transport plane carries out inputs flight test as carrier aircraft, and the helicopter after recent The Commitments then adopts a frame repacking completes extension and flies captive test and input flight test.
Adopt utility helicopter to carry out unpowered lift formula aircraft extension as carrier aircraft fly captive test and throw in flight test, there is the advantages such as improvement cost is low, the preparatory period is short, convenient, flexible, safe.But, because helicopter flight speed is low, aircraft faces the problem of control ability deficiency after being separated with the hanger of helicopter, throwing in initial condition (IC) deviation, be separated interference, likely attitude instability under the combined action of the factors such as winds aloft interference and aircraft static stability, thus cannot safe landing on predetermined runway.Therefore, adopt helicopter low speed to throw in unpowered lift formula aircraft, attitude stabilization and the control problem of aircraft in low dynamic pressure situation after throwing in must be solved.Be exactly specifically before throwing in, make the attitude motion As soon as possible Promising Policy of carry state aircraft throw in requirement, or make input initial condition (IC) deviation as far as possible within allowed band, increase the probability allowing to throw in; After aircraft is separated with hanger, keeps attitude stabilization, reduces the impact of throwing in initial condition (IC) deviation and interference, and accelerate to improve flying speed rapidly by diving, and then promote the control ability of pneumatic rudder face.
At present, adopt helicopter to complete as carrier aircraft the unpowered lift formula checking aircraft throwing in flight test at home and abroad in the documents and materials published to have: the ALFLEX (1996) of Japan, the X-40A (1998 and calendar year 2001) of the U.S., the Phoenix (2004) of Germany, the The Commitments (2013) of the U.S..These checking aircraft all achieve flight test success, and the technical scheme taked respectively has feature.ALFLEX hanger is comparatively complicated, and hanger adopts fin stabilization, and aircraft adopts angle of attack closed loop control; Simply, helicopter is by single cable wire hanging hanger and aircraft, and hanger is with drogue for X-40A and The Commitments hanging scheme; The hanging frame plan of Phoenix is simple, and helicopter is by three cable wire hanging aircraft, and Vee formation hanging point layout ensures that hanger and attitude of flight vehicle are stablized.Above technical scheme all has distinct disadvantage, restricted application.Such as, ALFLEX hanger composition is complicated, practical operation difficulty; X-40A and The Commitments are comparatively large by the impact of winds aloft under carry state, and this hanging protocols call aircraft static stability is good, can tolerate the impact of interference; Phoenix is comparatively remarkable by the impact of carrier aircraft, and three cable wires are easily wound around, and therefore requires strict to carrier aircraft flight.In addition, the aircraft manufacturing technology scheme that above technical scheme comprises also has shortcoming, such as, needs the angle of attack signal used to be difficult to accurate acquisition, adopts pitch angle and roll angle controlled reset to be thus difficult to ensure to be separated rear attitude stabilization simultaneously.
Summary of the invention
Technology of the present invention is dealt with problems and is: overcome the deficiencies in the prior art, the hanger providing a kind of helicopter low speed to throw in aircraft and Spacecraft Attitude Control, attitude stabilization and the control problem of aircraft in low dynamic pressure situation after throwing in when adopting helicopter low speed to throw in unpowered lift formula aircraft can be solved, there is the advantages such as system composition is simple, antijamming capability is strong, applied widely.
Technical solution of the present invention is:
Helicopter low speed throws in the hanger of aircraft, comprising: main cable, hanger cable wire, damping panel, hanger main structure, drogue, hold down assembly and be separated releasing unit; Hold down assembly and comprise clamping switch and adjustable presser foot;
One end of main cable is connected on helicopter, and the other end is by hanger cable wire connection suspension frame main structure; Damping panel is vertically welded in hanger main structure, swings for reducing hanger; Drogue is connected to the tail end of hanger main structure, for eliminating hanger swing and rotating around main cable; Four adjustable presser foots held down assembly are arranged in hanger main structure, four clamping switch are arranged on the back side of aircraft, four adjustable presser foots and four clamping switch one_to_one corresponding, predetermincd tension is applied by adjustable presser foot after hanger main structure is connected with aircraft, clamping switch is pushed down by adjustable presser foot, makes clamping switch be in off-state; Aircraft hangs over below hanger main structure by being separated releasing unit.
Described hanger cable wire is at least 2, before and after hanger at least each 1, and adjustable length, by the length of hanger cable wire before and after adjustment and then the pitch attitude of adjustment hanger.
Then four clamping switch parallel connections are between two cascaded, each clamping switch all comprises two autonomous channels, thus form two separate remaining separation signal testing circuits, realize separation signal by the on off mode detecting each remaining circuit two ends to detect, namely any one remaining line conduction produces separation signal, being open circuit when clamping switch is compacted, is path during clamping switch free state.
Before and after described separation releasing unit each one, often cover separation releasing unit includes hanger attaching parts and is separated priming system, hanger attaching parts is connected with hanger main structure, hanger attaching parts is connected with the hanger of aircraft by being separated priming system, and separation priming system detonates to realize hanger under instruction control and is separated with aircraft.
Based on the Spacecraft Attitude Control that hanger realizes, step is as follows:
(1) helicopter is by described hanger carry aircraft flight, and after the steering wheel of aircraft powers on, the flight-control computer of aircraft periodically carries out separation signal detection by two remaining separation signal testing circuits, and sense cycle is not more than 40ms;
(2) if the separation signal of not detecting, then aircraft carries out gesture stability in the following manner, returns step (1) afterwards; If the separation signal of detecting, then enter step (3);
Vertical passage and the course passage of aircraft all adopt opened loop control, and vertical passage elevating rudder drift angle instruction DEC is preset value, and course channel direction angle of rudder reflection instruction DRC is 0;
The interconnection of aircraft adopts roll angle closed loop control, and roll angle instruction GAMAC is set to 0, and to reduce to disturb the impact on the motion of aircraft roll attitude, the control law of interconnection is:
DAC=LIM (DACUL, KGAMA* (GAMAC-GAMA)-KWX*WX, DACLL), wherein GAMAC=0
GAMA is roll angle, WX is roll angle speed, KGAMA and KWX is ride gain, DAC is the instruction of aileron angle of rudder reflection, DACUL and DACLL is respectively higher limit and the lower limit of aileron angle of rudder reflection instruction DAC, allows range of use to determine according to aileron angle of rudder reflection, LIM (,) be amplitude limit link;
(3) detect after separation signal, the flight-control computer of aircraft periodically judges flying speed, and the judgement cycle is not more than 40ms;
(4) if the flying speed of aircraft is less than preset value, then aircraft carries out gesture stability in the following manner, returns step (3) afterwards; If the flying speed of aircraft is greater than preset value, then enter step (5);
The vertical passage of aircraft adopts pitch angle closed loop control, and pitch angle instruction THETAC is preset value, and span is-30 ° to-60 °, and accelerate to realize diving by commanded pitch attitude, the control law of vertical passage is:
AYC=LIM (AYCUL, KTHETA* (THETAC-THETA), AYCLL), wherein THETAC is preset value; DEC=LIM (DECUL, KAY* (AYC-AY)-KWZ*WZ, DECLL)
THETA is pitch angle, WZ is pitch rate, AY is normal acceleration, KTHETA, KWX and KAY are ride gain, AYC is normal acceleration instruction, AYCUL and AYCLL is respectively higher limit and the lower limit of normal acceleration instruction AYC, determine according to normal acceleration allowed band, DEC is the instruction of elevating rudder drift angle, DECUL and DECLL is respectively higher limit and the lower limit of elevating rudder drift angle instruction DEC, allows range of use to determine according to elevating rudder drift angle, LIM (,) be amplitude limit link;
The interconnection of aircraft adopts roll angle closed loop control, and roll angle instruction GAMAC is set to 0, and to reduce to disturb the impact on the motion of aircraft roll attitude, the control law of interconnection is:
DAC=LIM (DACUL, KGAMA* (GAMAC-GAMA)-KWX*WX, DACLL), wherein GAMAC=0
GAMA is roll angle, WX is roll angle speed, KGAMA and KWX is ride gain, DAC is the instruction of aileron angle of rudder reflection, DACUL and DACLL is respectively higher limit and the lower limit of aileron angle of rudder reflection instruction DAC, allows range of use to determine according to aileron angle of rudder reflection, LIM (,) be amplitude limit link;
The course passage of aircraft adopts lateral acceleration closed loop control, and considers the interchannel coupling of interconnection and course, and to eliminate the impact of unfavorable sideslip, the control law of course passage is:
DRC=LIM (DRCUL, KXY* (KGAMA* (GAMAC-GAMA)-KWX*WX)-KAZ*AZ-KWY* (s/ (s+wwy)) * WY, DRCLL), wherein GAMAC=0,
AZ is lateral acceleration, WY is yawrate, s/ (s+wwy) is for washing out the transfer function of network, KXY, KAZ and KWY are ride gain, and DRC is rudder instruction, DRCUL and DRCLL is respectively higher limit and the lower limit of rudder instruction DRC, range of use is allowed to determine according to rudder, LIM (,) and be amplitude limit link;
(5) flying speed of aircraft is greater than preset value, represents that the underriding accelerator of aircraft terminates.
Amplitude limit link LIM (UL, X, LL) be input as variable X, amplitude limit link is used for output to be limited between higher limit UL and lower limit LL, higher limit UL is not less than lower limit LL, when X is not less than UL, the output of amplitude limit link is UL, and when X is not more than LL, the output of amplitude limit link is LL, and when X is between UL and LL, the output of amplitude limit link is X.
The present invention's beneficial effect is compared with prior art:
(1) helicopter is by single main cable hanging hanger and aircraft, system is formed simple, avoid the wrapping phenomena that many cable wires may occur, reduce the requirement to helicopter flight, reduce the adverse effect to helicopter flight, therefore practical operation is easier.
(2) single main cable and hanger and aircraft form a single pendulum system, even if also certainly exist the simple harmonic motion of fore-and-aft direction and left and right directions when stabilized flight, aircraft also can rotate around main cable in addition, adopt drogue can eliminate aircraft swing campaign and around main cable rotational motion, adopt damping panel can reduce aircraft roll motion, therefore hanger and external disturbance can greatly be alleviated on the impact of attitude of flight vehicle by drogue and damping panel combined action, thus the requirement can relaxed environmental conditions and helicopter flight, or increase the probability allowing to throw in.
(3) parallel connection of multiple binary channel clamping switch and two separate remaining separation signal testing circuits in series are adopted, can ensure can normally work when any one clamping switch fault or any one sense channel fault, thus possess failure-free separation signal measuring ability.
(4) detect based on failure-free separation signal, by taking different Attitude Control Strategies before aircraft is separated from hanger with after being separated, attitude of flight vehicle can be increased to move jamproof ability, the requirement to aircraft static stability can be relaxed, even quiet unstable flight device also can ensure to be separated rear attitude stabilization and carry out underriding on request and accelerate, therefore Applicable scope is wider.In addition, after being separated, gesture stability guarantees that aircraft accelerates without rolling with without breakking away and carrying out underriding by commanded pitch attitude, can obtain best acceleration effect.
(5) aircraft manufacturing technology does not adopt the angle of attack to feed back and angle of side slip feedback, reduces the requirement of measuring the angle of attack and angle of side slip.Aircraft manufacturing technology adopts roll angle feedback, pitch angle feedback and normal acceleration feedback, can prevent roll angle, pitch angle attitude instability and normal g-load from transfiniting.
Accompanying drawing explanation
Fig. 1 is hanger of the present invention composition schematic diagram;
Fig. 2 is separation signal testing circuit schematic diagram of the present invention;
Fig. 3 is that the present invention is separated releasing unit connection structure schematic diagram;
Fig. 4 is Spacecraft Attitude Control diagram of circuit of the present invention;
Fig. 5 is that the present invention is separated front interconnection gesture stability schematic diagram;
Fig. 6 is that the present invention is separated rear vertical passage gesture stability schematic diagram;
Fig. 7 is that the present invention is separated rear interconnection and course passage gesture stability schematic diagram.
Detailed description of the invention
Below in conjunction with accompanying drawing, the hanger of proposed a kind of helicopter low speed input aircraft and Spacecraft Attitude Control are described in further detail:
As shown in Figure 1, the invention provides a kind of helicopter low speed and throw in the hanger of aircraft, comprising: main cable 1, hanger cable wire 2, damping panel 3, hanger main structure 4, drogue 5, hold down assembly 6 be separated releasing unit 7; Hold down assembly and 6 comprise clamping switch and adjustable presser foot;
Main cable 1 one end connects helicopter, and the other end connects hanger main structure 4 by hanger cable wire 2; Damping panel 3 is vertically welded in hanger main structure 4, swings for reducing hanger; Drogue 5 is connected to the tail end of hanger main structure 4, for eliminating hanger swing and rotating around main cable 1; Four hold down assembly 6 adjustable presser foot be arranged in hanger main structure 4, four clamping switch are arranged on the back side of aircraft 8, four adjustable presser foots and four clamping switch one_to_one corresponding, predetermincd tension is applied by adjustable presser foot after hanger main structure 4 is connected with aircraft 8, clamping switch is pushed down by adjustable presser foot, makes clamping switch be in off-state; Aircraft 8 hangs over below hanger main structure 4 by being separated releasing unit 7.
Described hanger cable wire 2 is at least 2, before and after hanger at least each 1, and adjustable length, by the length of hanger cable wire before and after adjustment and then the pitch attitude that can adjust hanger and aircraft.
As shown in Figure 1, the present invention adopts four hanger cable wires, hanger and aircraft can be made to bow by the length of length or shortening hanger rear portion two cable wires increasing anterior two cable wires of hanger, namely pitch angle reduces, otherwise, hanger and aircraft can be made to come back by the length of length or increase hanger rear portion two cable wires shortening anterior two cable wires of hanger, namely pitch angle increases.
As shown in Figure 2, then four clamping switch parallel connections are between two cascaded, each clamping switch all comprises two autonomous channels, thus form two separate remaining separation signal testing circuits, realize separation signal by the on off mode detecting each remaining circuit two ends to detect, namely any one remaining line conduction produces separation signal, is open circuit when clamping switch is compacted, is path during clamping switch free state;
To be separated before and after releasing unit 7 each one, as shown in Figure 3, often cover separation releasing unit 7 includes hanger attaching parts 701 and is separated priming system 702, hanger attaching parts 701 is connected with hanger main structure 4, hanger attaching parts 701 is connected with the hanger 801 of aircraft 8 by being separated priming system 702, and separation priming system 702 detonates to realize hanger under instruction control and is separated with aircraft 8.
Hanger of the present invention, except will meeting helicopter carry and discharging the requirement of lift formula aircraft, also should be and throws in that moment attitude of flight vehicle is stable to create conditions.Be exactly specifically, when helicopter stabilized flight, hanger and aircraft pitch attitude, swing, swinging and rotating around main cable all should meet input requirement, and throw in moment attitude of flight vehicle like this and will be in expected range, input afterwards attitude of flight vehicle is not easy unstability.
Main cable 1 length, hanger cable wire 2 length, damping panel 3 area, drogue 5 lanyard length, drogue 5 eigenwert are the important design parameter of hanger, determine that the step of these parameters is as follows:
(a1) length of main cable 1 is rule of thumb tentatively determined, generally between 10m to 40m.Main cable can not be too short, otherwise hanger and aircraft are subject to the impact of lifting airscrew purling serious, can not be oversize, otherwise attitude of flight vehicle is too large by the impact of main cable simple harmonic motion.
(b1) drogue 5 lanyard length is rule of thumb tentatively determined, generally between 3m to 15m.Drogue lanyard can not be too short, otherwise drogue affects comparatively large by hanger and aircraft, can not be oversize, otherwise affect larger by disturbing the drogue caused to swing on hanger and attitude of flight vehicle.
(c1) in conjunction with relevant design experience, hanger cable wire 2 length, damping panel 3 area and drogue 5 eigenwert is tentatively determined according to the mass property of hanger and aircraft.Hanger rope length should be convenient to the pitch attitude adjusting hanger and aircraft.Damping panel is used for producing aerodynamic damping power when hanger swings, and flaps area is decisive parameter, and when the gross area meets the demands, damping panel weight should be as far as possible little, to installation site and the form not requirement of damping panel.The pulling force that drogue produces equals the product of dynamic pressure and eigenwert.
(d1) carry out extension to fly captive test and carry out simulation to throw in, hanger design parameters is made a flight test checking.
(e1) according to the pitch angle of aircraft before throwing in, and the swinging of aircraft, swing and around main cable rotational motion, hanger design parameters is suitably adjusted.
(f1) step (c1) and (d1) is repeated, until hang hanger and flight state when flying captive test can meet input requirement.
As shown in Figure 4, based on hanger of the present invention, achieve a kind of Spacecraft Attitude Control, step is as follows:
(1) helicopter is flown by described hanger carry aircraft 8, and after the steering wheel of aircraft 8 powers on, the flight-control computer of aircraft 8 periodically carries out separation signal detection by two remaining separation signal testing circuits, and sense cycle is not more than 40ms;
(2) if the separation signal of not detecting, then aircraft 8 carries out gesture stability in the following manner, returns step (1) afterwards; If the separation signal of detecting, then enter step (3);
The vertical passage of aircraft 8 and course passage adopt opened loop control, and vertical passage elevating rudder drift angle instruction DEC is preset value, and course channel direction angle of rudder reflection instruction DRC is 0;
As shown in Figure 5, the interconnection of aircraft 8 adopts roll angle closed loop control, and roll angle instruction GAMAC is set to 0, and to reduce to disturb the impact on the motion of aircraft roll attitude, the control law of interconnection is:
DAC=LIM (DACUL, KGAMA* (GAMAC-GAMA)-KWX*WX, DACLL), wherein GAMAC=0
GAMA is roll angle, WX is roll angle speed, KGAMA and KWX is ride gain, DAC is the instruction of aileron angle of rudder reflection, DACUL and DACLL is respectively higher limit and the lower limit of aileron angle of rudder reflection instruction DAC, allows range of use to determine according to aileron angle of rudder reflection, LIM (,) be amplitude limit link.
Ride gain KGAMA and KWX, adopts flight control system conventional design method to determine for separation moment flight state.
Carry state attitude of flight vehicle determines primarily of helicopter and hanger, and the impact of pneumatic control surface deflection on attitude motion is less.Spacecraft Attitude Control of the present invention, carried out active attitude control before aircraft is separated with hanger.The object of pitch channel and jaw channel opened loop control provides initial condition (IC) for throwing in moment aircraft manufacturing technology, the object of roll channel closed loop control reduces aircraft to swing on the impact of roll attitude, and the roll angle of input moment aircraft is remained in expected range.
(3) detect after separation signal, the flight-control computer of aircraft 8 periodically judges flying speed, and the judgement cycle is not more than 40ms;
(4) if the flying speed of aircraft 8 is less than preset value, then aircraft 8 carries out gesture stability in the following manner, returns step (3) afterwards; If the flying speed of aircraft 8 is greater than preset value, then enter step (5);
As shown in Figure 6, the vertical passage of aircraft 8 adopts pitch angle closed loop control, and pitch angle instruction THETAC is preset value, and span is-30 ° to-60 °, such as THETAC=-45 °, and accelerate to realize diving by commanded pitch attitude, the control law of vertical passage is:
AYC=LIM (AYCUL, KTHETA* (THETAC-THETA), AYCLL), wherein THETAC is preset value; DEC=LIM (DECUL, KAY* (AYC-AY)-KWZ*WZ, DECLL)
THETA is pitch angle, WZ is pitch rate, AY is normal acceleration, KTHETA, KWX and KAY are ride gain, AYC is normal acceleration instruction, AYCUL and AYCLL is respectively higher limit and the lower limit of normal acceleration instruction AYC, determine according to normal acceleration allowed band, DEC is the instruction of elevating rudder drift angle, DECUL and DECLL is respectively higher limit and the lower limit of elevating rudder drift angle instruction DEC, allows range of use to determine according to elevating rudder drift angle, LIM (,) be amplitude limit link.
As shown in Figure 7, the interconnection of aircraft 8 adopts roll angle closed loop control, and roll angle instruction GAMAC is set to 0, and to reduce to disturb the impact on the motion of aircraft roll attitude, the control law of interconnection is:
DAC=LIM (DACUL, KGAMA* (GAMAC-GAMA)-KWX*WX, DACLL), wherein GAMAC=0
GAMA is roll angle, WX is roll angle speed, KGAMA and KWX is ride gain, DAC is the instruction of aileron angle of rudder reflection, DACUL and DACLL is respectively higher limit and the lower limit of aileron angle of rudder reflection instruction DAC, allows range of use to determine according to aileron angle of rudder reflection, LIM (,) be amplitude limit link.
As shown in Figure 7, the course passage of aircraft 8 adopts lateral acceleration closed loop control, and considers the interchannel coupling of interconnection and course, and to eliminate the impact of unfavorable sideslip, the control law of course passage is:
DRC=LIM (DRCUL, KXY* (KGAMA* (GAMAC-GAMA)-KWX*WX)-KAZ*AZ-KWY* (s/ (s+wwy)) * WY, DRCLL), wherein GAMAC=0
AZ is lateral acceleration, WY is yawrate, s/ (s+wwy) is for washing out the transfer function of network, KXY, KAZ and KWY are ride gain, and DRC is rudder instruction, DRCUL and DRCLL is respectively higher limit and the lower limit of rudder instruction DRC, range of use is allowed to determine according to rudder, LIM (,) and be amplitude limit link.
Ride gain KTHETA, KWX, KAY, KGAMA, KWX, KXY, KAZ and KWY, adopt flight control system conventional design method to determine for flight state after separation.
Spacecraft Attitude Control of the present invention, making aircraft accelerating without rolling with without carrying out underriding by commanded pitch attitude when sideslip, can obtain best acceleration effect like this after aircraft is separated with hanger.Pitch angle instruction THETAC preset value can be optimized by the analysis of aircraft Six-degree-of-freedom Simulation, to obtain best underriding acceleration effect.
(5) flying speed of aircraft 8 is greater than preset value, and represent that underriding accelerator terminates, the requirement according to follow-up phase Guidance and control is carried out gesture stability by aircraft 8, THETAC and GAMAC is no longer preset value.
Described amplitude limit link LIM (UL, X, LL) be input as variable X, amplitude limit link is used for output to be limited between higher limit UL and lower limit LL, higher limit UL is not less than lower limit LL, when X is not less than UL, the output of amplitude limit link is UL, and when X is not more than LL, the output of amplitude limit link is LL, and when X is between UL and LL, the output of amplitude limit link is X.
The content be not described in detail in specification sheets of the present invention belongs to the known technology of professional and technical personnel in the field.
Claims (6)
1. helicopter low speed throws in the hanger of aircraft, it is characterized in that comprising: main cable (1), hanger cable wire (2), damping panel (3), hanger main structure (4), drogue (5), hold down assembly (6) be separated releasing unit (7); Hold down assembly (6) comprise clamping switch and adjustable presser foot;
One end of main cable (1) is connected on helicopter, and the other end connects hanger main structure (4) by hanger cable wire (2); Damping panel (3) is vertically welded in hanger main structure (4), swings for reducing hanger; Drogue (5) is connected to the tail end of hanger main structure (4), for eliminating hanger swing and rotating around main cable (1); The adjustable presser foot of four hold down assembly (6) is arranged in hanger main structure (4), four clamping switch are arranged on the back side of aircraft (8), four adjustable presser foots and four clamping switch one_to_one corresponding, predetermincd tension is applied by adjustable presser foot after hanger main structure (4) is connected with aircraft (8), clamping switch is pushed down by adjustable presser foot, makes clamping switch be in off-state; Aircraft (8) hangs over hanger main structure (4) below by being separated releasing unit (7).
2. a kind of helicopter low speed according to claim 1 throws in the hanger of aircraft, it is characterized in that: described hanger cable wire (2) is at least 2, before and after hanger at least each 1, and adjustable length, by the length of hanger cable wire before and after adjustment and then the pitch attitude of adjustment hanger.
3. a kind of helicopter low speed according to claim 1 throws in the hanger of aircraft, it is characterized in that: then four clamping switch parallel connections are between two cascaded, each clamping switch all comprises two autonomous channels, thus form two separate remaining separation signal testing circuits, realize separation signal by the on off mode detecting each remaining circuit two ends to detect, namely any one remaining line conduction produces separation signal, being open circuit when clamping switch is compacted, is path during clamping switch free state.
4. a kind of helicopter low speed according to claim 1 throws in the hanger of aircraft, it is characterized in that: described each one of separation releasing unit (7) front and back, often cover separation releasing unit (7) includes hanger attaching parts (701) and is separated priming system (702), hanger attaching parts (701) is connected with hanger main structure (4), hanger attaching parts (701) is connected with the hanger (801) of aircraft (8) by being separated priming system (702), separation priming system (702) is detonated to realize hanger under instruction control and is separated with aircraft (8).
5., based on the Spacecraft Attitude Control that claim 1 hanger realizes, it is characterized in that step is as follows:
(1) helicopter is by described hanger carry aircraft (8) flight, after the steering wheel of aircraft (8) powers on, the flight-control computer of aircraft (8) periodically carries out separation signal detection by two remaining separation signal testing circuits, and sense cycle is not more than 40ms;
(2) if the separation signal of not detecting, then aircraft (8) carries out gesture stability in the following manner, returns step (1) afterwards; If the separation signal of detecting, then enter step (3);
Vertical passage and the course passage of aircraft (6) all adopt opened loop control, and vertical passage elevating rudder drift angle instruction DEC is preset value, and course channel direction angle of rudder reflection instruction DRC is 0;
The interconnection of aircraft (6) adopts roll angle closed loop control, and roll angle instruction GAMAC is set to 0, and to reduce to disturb the impact on the motion of aircraft roll attitude, the control law of interconnection is:
DAC=LIM (DACUL, KGAMA* (GAMAC-GAMA)-KWX*WX, DACLL), wherein GAMAC=0
GAMA is roll angle, WX is roll angle speed, KGAMA and KWX is ride gain, DAC is the instruction of aileron angle of rudder reflection, DACUL and DACLL is respectively higher limit and the lower limit of aileron angle of rudder reflection instruction DAC, allows range of use to determine according to aileron angle of rudder reflection, LIM (,) be amplitude limit link;
(3) detect after separation signal, the flight-control computer of aircraft (8) periodically judges flying speed, and the judgement cycle is not more than 40ms;
(4) if the flying speed of aircraft (8) is less than preset value, then aircraft (8) carries out gesture stability in the following manner, returns step (3) afterwards; If the flying speed of aircraft (8) is greater than preset value, then enter step (5);
The vertical passage of aircraft (8) adopts pitch angle closed loop control, and pitch angle instruction THETAC is preset value, and span is-30 ° to-60 °, and accelerate to realize diving by commanded pitch attitude, the control law of vertical passage is:
AYC=LIM (AYCUL, KTHETA* (THETAC-THETA), AYCLL), wherein THETAC is preset value; DEC=LIM (DECUL, KAY* (AYC-AY)-KWZ*WZ, DECLL)
THETA is pitch angle, WZ is pitch rate, AY is normal acceleration, KTHETA, KWX and KAY are ride gain, AYC is normal acceleration instruction, AYCUL and AYCLL is respectively higher limit and the lower limit of normal acceleration instruction AYC, determine according to normal acceleration allowed band, DEC is the instruction of elevating rudder drift angle, DECUL and DECLL is respectively higher limit and the lower limit of elevating rudder drift angle instruction DEC, allows range of use to determine according to elevating rudder drift angle, LIM (,) be amplitude limit link;
The interconnection of aircraft (6) adopts roll angle closed loop control, and roll angle instruction GAMAC is set to 0, and to reduce to disturb the impact on the motion of aircraft roll attitude, the control law of interconnection is:
DAC=LIM (DACUL, KGAMA* (GAMAC-GAMA)-KWX*WX, DACLL), wherein GAMAC=0
GAMA is roll angle, WX is roll angle speed, KGAMA and KWX is ride gain, DAC is the instruction of aileron angle of rudder reflection, DACUL and DACLL is respectively higher limit and the lower limit of aileron angle of rudder reflection instruction DAC, allows range of use to determine according to aileron angle of rudder reflection, LIM (,) be amplitude limit link;
The course passage of aircraft (6) adopts lateral acceleration closed loop control, and considers the interchannel coupling of interconnection and course, and to eliminate the impact of unfavorable sideslip, the control law of course passage is:
DRC=LIM (DRCUL, KXY* (KGAMA* (GAMAC-GAMA)-KWX*WX)-KAZ*AZ-KWY* (s/ (s+wwy)) * WY, DRCLL), wherein GAMAC=0,
AZ is lateral acceleration, WY is yawrate, s/ (s+wwy) is for washing out the transfer function of network, KXY, KAZ and KWY are ride gain, and DRC is rudder instruction, DRCUL and DRCLL is respectively higher limit and the lower limit of rudder instruction DRC, range of use is allowed to determine according to rudder, LIM (,) and be amplitude limit link;
(5) flying speed of aircraft (8) is greater than preset value, represents that the underriding accelerator of aircraft terminates.
6. a kind of Spacecraft Attitude Control according to claim 5, it is characterized in that: amplitude limit link LIM (UL, X, LL) be input as variable X, amplitude limit link is used for output to be limited between higher limit UL and lower limit LL, and higher limit UL is not less than lower limit LL, and when X is not less than UL, the output of amplitude limit link is UL, when X is not more than LL, the output of amplitude limit link is LL, and when X is between UL and LL, the output of amplitude limit link is X.
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