CN109240335B - Aerospace vehicle approach landing guidance method - Google Patents

Aerospace vehicle approach landing guidance method Download PDF

Info

Publication number
CN109240335B
CN109240335B CN201811270811.8A CN201811270811A CN109240335B CN 109240335 B CN109240335 B CN 109240335B CN 201811270811 A CN201811270811 A CN 201811270811A CN 109240335 B CN109240335 B CN 109240335B
Authority
CN
China
Prior art keywords
aircraft
flight
height
energy
section
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
CN201811270811.8A
Other languages
Chinese (zh)
Other versions
CN109240335A (en
Inventor
刘智勇
何英姿
范松涛
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Beijing Institute of Control Engineering
Original Assignee
Beijing Institute of Control Engineering
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Beijing Institute of Control Engineering filed Critical Beijing Institute of Control Engineering
Priority to CN201811270811.8A priority Critical patent/CN109240335B/en
Publication of CN109240335A publication Critical patent/CN109240335A/en
Application granted granted Critical
Publication of CN109240335B publication Critical patent/CN109240335B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/10—Simultaneous control of position or course in three dimensions
    • G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
    • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft

Landscapes

  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • Traffic Control Systems (AREA)

Abstract

An aerospace vehicle approach landing guidance method mainly comprises the following steps: determining a flight energy-altitude corridor, sequentially dividing a flight track into a first transition section, an equal dynamic pressure flight section, a second transition section and a small sinking rate flight section according to the flight energy of an aircraft from big to small, determining an aircraft altitude characteristic profile of each track, adjusting the aircraft altitude characteristic profile of each track according to range deviation, determining an aircraft attack angle instruction, and simultaneously determining a resistance plate opening angle instruction and an aircraft inclination angle instruction, thereby completing the air-to-ground aircraft landing guidance work. The approach landing guidance method has the advantages of strong self-adaptive capacity, high precision, simple calculation and easy engineering realization.

Description

Aerospace vehicle approach landing guidance method
Technical Field
The invention relates to an approach landing guidance method for an aerospace vehicle, belongs to the technical field of landing guidance, and is particularly suitable for approach landing of aerospace vehicles, reusable vehicles and hypersonic vehicles.
Background
An Aerospace Vehicle (ASV) (also called an Aerospace plane) is a novel aircraft capable of working in both the Aerospace field and the Aerospace field, and combines the Aerospace technology and the Aerospace technology. The aerospace craft can take off horizontally like a common airplane, fly in the atmosphere at a hypersonic speed, can directly accelerate into the earth orbit to become an aerospace craft, and land at an airport like an airplane after returning to the atmosphere.
The approach landing section is the last stage of the return landing process of the aerospace craft, the aerospace craft has characteristics of unpowered flight and small lift-drag ratio in the stage, and is a key technology of the aerospace craft, which is greatly different from the landing of a general aircraft. The approach landing segment is located behind the terminal energy management segment, and the main purpose of the approach landing segment is to manage the energy of the aerospace vehicle and control the aircraft to arrive at the airport runway at the proper altitude, speed, subsidence rate and course. A typical approach landing segment starts at a ground speed of about 156m/s, a height of about 3000m from the ground, a distance of about 13880m from the airport runway, and ends at the airport runway at a speed of about 100m/s and a height of 0m from the ground.
The demonstration and verification of the unpowered launch test is an important stage of aerospace vehicle development, is used for verifying the landing performance during unpowered autonomous approach, and has a large initial condition range and increased guidance control difficulty compared with normal flight. An unpowered throwing test for the low-speed aircraft to fly off is carried out at initial flying speed of 50m/s, initial height to ground of 4025m, initial distance to airport runway of 11333m, and speed of 100m/s and height to ground of 0 m. It can be seen that it varies greatly from the initial conditions of a typical approach landing leg flight.
Therefore, the aerospace vehicle approach landing technology which is suitable for the large-range initial condition is researched and designed, the requirement of the aerospace vehicle for normal approach landing flight is met, the requirement of the unpowered launch test is met, and the aerospace vehicle approach landing technology has important practical significance as a general key technology.
In the aspect of normal flight autonomous landing, the automatic landing track design of the space shuttle adopts a steep glide section, an arc pulling section, an index transition section and a shallow glide section, four tracks can be selected for the landing section, which track is selected depends on the quality and energy condition of the space shuttle, and different guidance structures are adopted in different flight stages, so that the design requirement is met during landing. The Draper laboratory takes X-34 unpowered automatic landing as a research object, provides a trajectory design method based on a height profile, and compared with a trajectory design method based on time adopted by a space shuttle, the trajectory design method based on the height profile conforms to the physical essence of the trajectory profile and has obvious design advantages. It can be seen that the automatic landing technology of normal flight largely inherits the work of the American space shuttle, and is improved on the basis of the frame of unpowered automatic landing of the space shuttle. In the aspect of unpowered landing, a launch test of a scaling model of the unmanned space vehicle is carried out in Japan through an ALFLEX project, and a section of fixed track is added to the track shape of the ALFLEX unpowered landing for capture so as to be connected with an automatic landing section, so that the adaptability is poor.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: the method has the advantages of high guidance precision, simple calculation and easy engineering realization, and improves the initial condition adaptability of the aerospace vehicle.
The technical scheme of the invention is as follows:
an aerospace vehicle approach landing guidance method comprises the following steps:
1) determining a flight energy-altitude corridor;
2) judging whether the aircraft is located in a flight energy-height corridor or not according to the flight height and the energy of the aircraft, and entering a step 3 if the aircraft is not located in the flight energy-height corridor; if the aircraft is in the flight energy-altitude corridor, then step 4)
3) Carrying out longitudinal guidance and lateral guidance on the aircraft until the flight height and the energy of the aircraft are positioned in a flight energy-height corridor, and entering the step 4);
4) performing longitudinal guidance according to the range deviation of the aircraft, performing lateral guidance according to the lateral distance deviation of the aircraft to obtain longitudinal guidance information and lateral guidance information, and entering the step 5);
5) and (3) transmitting the longitudinal guidance information and the lateral guidance information obtained in the step (4) to an aircraft control system for controlling the aircraft to fly and finishing the landing guidance work of the aircraft during approach.
Said step 1) of determining the flight energy-altitude corridor comprises determining a lower altitude limit h of the flight energy-altitude corridorminRelationship to aircraft flight energy and upper altitude limit h of flight energy-altitude corridormaxThe relation with the flight energy of the aircraft is as follows:
11) a lower altitude limit h of the flight energy-altitude corridorminThe relationship to the flight energy of the aircraft is as follows:
Figure BDA0001845936000000031
wherein E is the flight energy of the aircraft, mu is the gravitational constant, R0Is the airport plane earth radius, qmaxFor aircraft movingUpper limit of pressure ρ0Is the atmospheric density, h, of the aircraft landing site levelsIs an atmospheric characteristic constant;
12) an upper altitude limit h of the flight energy-altitude corridormaxThe relationship to the flight energy of the aircraft is as follows:
Figure BDA0001845936000000032
where m is the aircraft mass, SrefFor aircraft reference area, CL maxThe coefficient of lift is the coefficient of lift at the maximum lift-drag ratio of the aircraft.
The method for laterally guiding the aircraft in the step 3) specifically comprises the following steps: and keeping the inclination angle of the aircraft to be zero as lateral guidance information, and transmitting the lateral guidance information to an aircraft control system to finish lateral guidance.
The aircraft longitudinal guidance method in the step 3) specifically comprises the following steps: determining an attack angle instruction of the aircraft, keeping an initial value of a starting angle instruction of the resistance plate unchanged and the determined attack angle instruction of the aircraft as longitudinal guidance information, and transmitting the longitudinal guidance information to an aircraft control system to complete longitudinal guidance;
the attack angle command alpha of the aircraft is specifically as follows:
Figure BDA0001845936000000041
Figure BDA0001845936000000042
where m is the aircraft mass, v is the aircraft velocity, ρ is the atmospheric density, SrefIs the aircraft reference area, kCD_α_0Is the ratio of flight angle of attack to drag coefficient of the aircraft, q is the real-time dynamic pressure of the aircraft, q is the ratio of flight angle of attack to drag coefficient of the aircraftcIs the dynamic pressure threshold value of the aircraft, k is the guidance instruction gain coefficient,
Figure BDA0001845936000000043
for a set fly-height-energy rate of change, θ is the aircraft track inclination, and D is the drag acceleration.
The method for performing longitudinal guidance according to the range deviation of the aircraft in the step 4) specifically comprises the following steps:
41) allocating the range deviation of the aircraft, and determining an attack angle instruction and a resistance plate opening angle instruction according to the allocation result of the range deviation;
42) and taking the attack angle instruction and the resistance plate opening angle instruction determined in the step 41) as longitudinal guidance information.
The step 41) is a method for determining an attack angle instruction according to the distribution result of the range deviation, and specifically comprises the following steps:
411) performing trajectory planning, and sequentially dividing a flight trajectory into a first transition section, an equal dynamic pressure flight section, a second transition section and a small sinking rate flight section from large to small according to the flight energy of the aircraft; determining the aircraft height characteristic profile of each section of track;
412) adjusting the aircraft height characteristic profile determined in the step 411) in real time according to the distribution result of the range deviation; and determining the aircraft attack angle instruction of each section of track in real time according to the adjusted aircraft height characteristic profile.
The aircraft height characteristic profile of each section of track determined in step 411) specifically includes:
A) the first transition section aircraft height characteristic profile hbThe method specifically comprises the following steps:
hb＝cb0+cb1E+cb2E2+cb3E3，
Ed＜E，
where E is aircraft energy, cb0、cb1、cb2And cb3The characteristic section coefficient of the aircraft height of the first transition section; edDesigning an energy design value at the intersection point of the first transition section and the equal dynamic pressure flight section;
B) the height characteristic profile h of the aircraft in the equal dynamic pressure flight sectiondThe method specifically comprises the following steps:
Figure BDA0001845936000000051
Figure BDA0001845936000000052
EL＜E≤Ed，
wherein h isd0To the starting height of the constant-pressure flight section, ElDesigning an energy design value at the intersection of the constant-pressure flight section and the second transition section;
C) the second transition section aircraft height characteristic profile hLThe method specifically comprises the following steps:
hL＝cL0+cL1E+cL2E2+cL3E3,
Eq＜E≤EL，
wherein, cL0、cL1、cL2And cL3For the second transition aircraft altitude characteristic profile coefficient, EqThe energy design value of the intersection point of the second transition section and the small-sinking-rate flight section is obtained;
D) the height characteristic profile h of the small-sinking-rate flight segment aircraftqThe method specifically comprises the following steps:
Figure BDA0001845936000000053
Figure BDA0001845936000000054
wherein h isfNominal altitude for aircraft terminal, EfNominal energy, g, for aircraft terminalsfFor nominal acceleration of gravity, v, of the terminalfFor the nominal speed of the terminal to be,
Figure BDA0001845936000000055
altitude-energy rate of change, θ, for flight segment with small sinking rateqFlight path inclination for flight section with small sinking rate, DqAcceleration of resistance for flight section of small sinking rate, CD_qIs the coefficient of resistance.
The step 412) of adjusting the aircraft altitude characteristic profile in real time includes:
determining the adjustment quantity delta h of a flight height profile, adjusting the height characteristic profile of the isokinetic pressure flight section aircraft in real time according to the delta h to obtain the adjusted height characteristic profile of the isokinetic pressure flight section aircraft, and adjusting the height characteristic profiles of the first transition section and the second transition section aircraft according to the real-time adjustment result of the height characteristic profile of the isokinetic pressure flight section aircraft to obtain the adjusted height characteristic profiles of the first transition section and the second transition section aircraft; and the altitude characteristic profile of the small-sinking-rate flight section aircraft is kept unchanged.
The height characteristic profile h of the aircraft of the adjusted equal dynamic pressure flight sectiond_ΔhFirst transition section aircraft altitude characteristic profile hb_ΔhAnd a characteristic profile h of the aircraft altitude of the second transition sectionL_ΔhThe method specifically comprises the following steps:
Figure BDA0001845936000000061
hL_Δh＝cL0_Δh+cL1_ΔhE+cL2_ΔhE2+cL3_ΔhE3，
hb_Δh＝cb0_Δh+cb1_ΔhE+cb2_ΔhE2+cb3_ΔhE3，
wherein, cL0_Δh、cL1_Δh、cL2_Δh、cL3_ΔhFor the adjusted second transition section aircraft altitude characteristic profile coefficient, cb0_Δh、cb1_Δh、cb2_ΔhAnd cb3_ΔhAnd obtaining the adjusted first transition section aircraft height characteristic section coefficient.
The method for determining the flight height profile adjustment quantity delta h specifically comprises the following steps:
Figure BDA0001845936000000062
Figure BDA0001845936000000063
Δsh＝(1-ks)Δs,
wherein,. DELTA.hkΔ h is the adjustment of the current guidance period flight altitude profile, Δ hk-1The adjustment of the flight height profile of the previous guidance period, dt is the magnitude of the guidance period, gammah>0，λ>0，EnIs the current energy of the aircraft, Δ s is the range deviation of the aircraft, ksThe distribution coefficient is adjusted for the flight distance, and k is the flight path of the aircraft in the second transition section or the flight section with small sinking rate s0; k when the flight path of the aircraft is located in the first transition section or the equal dynamic pressure flight sectionsIs not zero.
The step 412) is a method for determining the aircraft attack angle instruction of each section of track in real time according to the adjusted aircraft altitude characteristic profile, and specifically comprises the following steps:
αk＝αk-1+Δα，
Figure BDA0001845936000000071
Figure BDA0001845936000000072
wherein alpha iskFor angle of attack commands of the current guidance period, αk-1The angle of attack instruction of the previous guidance period, xi and omega are damping ratio of the guidance system and oscillation frequency of the guidance system, and nhFor real-time normal acceleration of the aircraft, hrefFor the height value corresponding to the current energy-height profile,
Figure BDA0001845936000000073
the height rate value corresponding to the current energy-height profile,
Figure BDA0001845936000000074
height acceleration value, n, corresponding to the current energy-height profilehIs the current normal acceleration, k, of the aircraftLIs the proportional coefficient of the lift coefficient and the attack angle, h is the aircraft height,
Figure BDA0001845936000000075
is the altitude rate of change of the aircraft, v is the speed of the aircraft, theta is the track inclination of the aircraft, sigmadIs the aircraft roll angle provided by the navigation system.
The step 41) is a method for determining the opening angle instruction of the resistance plate according to the distribution result of the range deviation, and specifically comprises the following steps:
Figure BDA0001845936000000076
Figure BDA0001845936000000077
Δsη＝ksΔs,
wherein eta isk-1For the drag-plate opening angle command of the previous guidance period, ηkFor the current guidance period flap opening angle command, gammaη>0, m is the aircraft mass, Δ s is the range deviation of the aircraft, ksAdjusting distribution coefficients for the voyage; k when the flight path of the aircraft is in the second transition section or the small-sinking-rate flight section s0; k when the flight path of the aircraft is located in the first transition section or the equal dynamic pressure flight sectionsIs not zero.
The step 4) is a method for lateral guidance according to the lateral distance deviation of the aircraft, and specifically comprises the following steps: determining an aircraft roll angle instruction, and taking the determined aircraft roll angle instruction as lateral guidance information; the aircraft roll angle command is determined as follows:
Figure BDA0001845936000000081
wherein y is the lateral distance deviation of the aircraft;
Figure BDA0001845936000000082
is the lateral velocity of the aircraft; k is a radical ofzp>0,kzd>0。
Compared with the prior art, the invention has the advantages that:
1) the invention provides a novel landing guidance process, which divides the whole process into a flight state adjusting section and a flight profile on-line adjusting section, and is favorable for adapting to initial conditions of large-scale changes such as normal landing flight, hanging flight and launching and the like.
2) The invention comprehensively utilizes the flight height and the opening angle of the resistance plate to adjust the flight range deviation, instead of utilizing a single means, improves the range adjustment capability of the aircraft, is insensitive to initial condition errors and various disturbances, is a simple mathematical operation formula, has low requirement on the performance of a flight control computer, and is easy to realize engineering.
3) The method is based on height online generation and adjustment instead of adopting a fixed height profile mode, improves the online range adjustment capability of the aerospace vehicle, avoids the problem of poor adaptability caused by adopting a fixed height profile in the prior art, and improves the guidance robustness.
4) According to the invention, the range is adjusted on line in real time based on the opening angle of the resistance plate, the height and the speed of the terminal are ensured by tracking the energy-height profile, instead of tracking the fixed speed profile by the resistance plate, the on-line range adjusting capability of the aerospace vehicle is improved, the problem of poor adaptability caused by tracking the fixed speed profile by the resistance plate in the prior art is avoided, and the guidance robustness is improved.
Drawings
FIG. 1 is a flow chart of the method of the present invention;
FIG. 2 is an energy level corridor and energy-level feature profile according to an embodiment of the present invention;
FIG. 3 is a plot of angle of attack commands in an embodiment of the present invention;
FIG. 4 is a roll angle command curve according to an embodiment of the present invention;
FIG. 5 is a dynamic pressure curve in an embodiment of the present invention;
FIG. 6 is a command curve of the opening angle of the resistance plate according to the embodiment of the present invention;
FIG. 7 shows μ in an embodiment of the present inventionhA curve;
FIG. 8 shows μ in an embodiment of the present inventionηCurve line.
Detailed Description
The aerospace vehicle approach landing technology suitable for the large-range initial conditions is high in guidance precision, can adapt to the initial conditions of large-range changes such as normal landing flight, hanging flight and launching and the like, is insensitive to initial condition errors and various disturbances, can better integrate longitudinal and transverse guidance, has low requirements on the performance of a flight control computer, and is easy to realize in engineering.
An aerospace vehicle approach landing guidance method, as shown in fig. 1, includes the following steps:
1) determining flight energy-altitude corridor
Determining the flight energy-altitude corridor comprises determining a lower altitude limit h of the flight energy-altitude corridorminRelationship to aircraft flight energy and upper altitude limit h of flight energy-altitude corridormaxThe relation with the flight energy of the aircraft is as follows:
11) a lower altitude limit h of the flight energy-altitude corridorminThe relationship to the flight energy of the aircraft is as follows:
Figure BDA0001845936000000091
wherein E is the flight energy of the aircraft, mu is the gravitational constant, R0Is the airport plane earth radius, qmaxUpper limit of dynamic pressure of aircraft, ρ0Large level of landing point of aircraftAir tightness, hsIs an atmospheric characteristic constant;
12) an upper altitude limit h of the flight energy-altitude corridormaxThe relationship to the flight energy of the aircraft is as follows:
Figure BDA0001845936000000101
where m is the aircraft mass, SrefFor aircraft reference area, CLmaxThe coefficient of lift is the coefficient of lift at the maximum lift-drag ratio of the aircraft.
2) Judging whether the aircraft is located in a flight energy-height corridor or not according to the flight height and the energy of the aircraft, and entering a step 3 if the aircraft is not located in the flight energy-height corridor; if the aircraft is located in the flight energy-height corridor, entering the step 4);
3) carrying out longitudinal guidance and lateral guidance on the aircraft until the flight height and the energy of the aircraft are positioned in a flight energy-height corridor, and entering the step 4);
keeping the inclination angle of the aircraft to be zero as lateral guidance information, and transmitting the lateral guidance information to an aircraft control system for lateral guidance of the aircraft; meanwhile, determining an attack angle instruction alpha of the aircraft, keeping an initial value of a starting angle instruction of the resistance plate unchanged and the determined attack angle instruction alpha of the aircraft as longitudinal guidance information, and transmitting the longitudinal guidance information to an aircraft control system to complete longitudinal guidance;
the attack angle command alpha of the aircraft is specifically as follows:
Figure BDA0001845936000000102
Figure BDA0001845936000000103
where m is the aircraft mass, v is the aircraft velocity, ρ is the atmospheric density, SrefTo flyReference area of the traveling gear, kCD_α_0Q is the ratio of the flight angle of attack to the drag coefficient of the aircraft, and q is the real-time dynamic pressure of the aircraft, as shown in FIG. 5, qcThe value range of the dynamic pressure threshold value of the aircraft and the dynamic pressure design value for switching the attack angle instruction mode is 1000 Pa-6000 Pa, k is a guidance instruction gain coefficient, k is more than 0 and less than 1,
Figure BDA0001845936000000104
for a set fly-height-energy rate of change, θ is the aircraft track inclination, and D is the drag acceleration.
4) Performing longitudinal guidance according to the range deviation of the aircraft, performing lateral guidance according to the lateral distance deviation of the aircraft to obtain longitudinal guidance information and lateral guidance information, and entering the step 5);
41) allocating the range deviation of the aircraft, and determining an attack angle instruction and a resistance plate opening angle instruction according to the allocation result of the range deviation;
42) and taking the attack angle instruction and the resistance plate opening angle instruction determined in the step 41) as longitudinal guidance information.
Carrying out range deviation distribution of the aircraft, determining range deviation adjusted by using an attack angle instruction and range deviation adjusted by using the resistance plates, and determining opening angle instructions of the resistance plates according to the range deviation adjusted by using the resistance plates;
44) taking the aircraft attack angle command of each track determined in the step 43) and the opening angle command of the resistance plate determined in the step 41) as longitudinal guidance information.
The step 41) is a method for determining an attack angle instruction according to the distribution result of the range deviation, and specifically comprises the following steps:
411) performing trajectory planning, and sequentially dividing a flight trajectory into a first transition section, an equal dynamic pressure flight section, a second transition section and a small sinking rate flight section from large to small according to the flight energy of the aircraft; determining the aircraft height characteristic profile of each section of track;
412) adjusting the aircraft height characteristic profile determined in the step 411) in real time according to the distribution result of the range deviation; and determining the aircraft attack angle instruction of each section of track in real time according to the adjusted aircraft height characteristic profile.
The aircraft height characteristic profile of each section of track determined in step 411) specifically includes:
A) the first transition section aircraft height characteristic profile hbThe method specifically comprises the following steps:
hb＝cb0+cb1E+cb2E2+cb3E3，
Ed＜E，
where E is aircraft energy, cb0、cb1、cb2And cb3The characteristic section coefficient of the aircraft height of the first transition section; edDesigning an energy design value at the intersection point of the first transition section and the equal dynamic pressure flight section; c. Cb0、cb1、cb2And cb3According to the initial height h of the equal dynamic pressure flight sectiond0Altitude-energy rate of change of equal dynamic pressure flight section
Figure BDA0001845936000000121
The coefficient c can be solved by four constraint conditions of the height at the tail end of the flight state adjusting section and the height-energy change rate at the tail end of the flight state adjusting sectionb0、cb1、cb2And cb3。
B) The height characteristic profile h of the aircraft in the equal dynamic pressure flight sectiondThe method specifically comprises the following steps:
Figure BDA0001845936000000122
Figure BDA0001845936000000123
EL＜E≤Ed，
wherein h isd0To the starting height of the constant-pressure flight section, ElDesigning an energy design value at the intersection of the constant-pressure flight section and the second transition section;
C) the second transition section aircraft height characteristic profile hLThe method specifically comprises the following steps:
hL＝cL0+cL1E+cL2E2+cL3E3,
Eq＜E≤EL，
wherein, cL0、cL1、cL2And cL3For the second transition aircraft altitude characteristic profile coefficient, EqThe energy design value of the intersection point of the second transition section and the small-sinking-rate flight section is obtained; c. CL0、cL1、cL2And cL3According to the height h of the tail end of the equal dynamic pressure flight sectiondfAltitude-energy rate of change of equal dynamic pressure flight section
Figure BDA0001845936000000124
Height h of starting end of flight segment with small sinking rateq0And altitude-energy rate of change of flight section with small sinking rate
Figure BDA0001845936000000125
And (4) determining.
D) The height characteristic profile h of the small-sinking-rate flight segment aircraftqThe method specifically comprises the following steps:
Figure BDA0001845936000000126
Figure BDA0001845936000000127
wherein h isfThe aircraft terminal nominal altitude; efNominal energy, g, for aircraft terminalsfFor nominal acceleration of gravity, v, of the terminalfFor the nominal speed of the terminal to be,
Figure BDA0001845936000000128
altitude-energy rate of change, θ, for flight segment with small sinking rateqFlight path inclination for flight section with small sinking rate, DqAcceleration of resistance for flight section of small sinking rate, CD_qIs the coefficient of resistance.
The step 412) of adjusting the aircraft altitude characteristic profile in real time includes:
determining the adjustment quantity delta h of a flight height profile, adjusting the height characteristic profile of the isokinetic pressure flight section aircraft in real time according to the delta h to obtain the adjusted height characteristic profile of the isokinetic pressure flight section aircraft, and adjusting the height characteristic profiles of the first transition section and the second transition section aircraft according to the real-time adjustment result of the height characteristic profile of the isokinetic pressure flight section aircraft to obtain the adjusted height characteristic profiles of the first transition section and the second transition section aircraft; the height characteristic profile h of the small-sinking-rate flight segment aircraftqRemain unchanged.
The height characteristic profile h of the aircraft of the adjusted equal dynamic pressure flight sectiond_ΔhFirst transition section aircraft altitude characteristic profile hb_ΔhAnd a characteristic profile h of the aircraft altitude of the second transition sectionL_ΔhThe method specifically comprises the following steps:
Figure BDA0001845936000000131
hL_Δh＝cL0_Δh+cL1_ΔhE+cL2_ΔhE2+cL3_ΔhE3，
wherein, cL0_Δh、cL1_Δh、cL2_ΔhAnd cL3_ΔhAccording to h, the adjusted height characteristic section coefficient of the second transition section aircraftd_ΔhDetermining; the method is specifically determined according to four constraint conditions of the height of the adjusted equal dynamic pressure flight tail end, the height-energy change rate of the equal dynamic pressure flight, the height of the shallow gliding start end and the height-energy change rate of the small sinking rate flight section.
hb_Δh＝cb0_Δh+cb1_ΔhE+cb2_ΔhE2+cb3_ΔhE3，
Wherein, cb0_Δh、cb1_Δh、cb2_ΔhAnd cb3_ΔhAccording to h, the adjusted first transition section aircraft altitude characteristic section coefficientd_ΔhDetermining; the method is specifically determined according to the height of the starting end of the adjusted equal dynamic pressure flight section, the equal dynamic pressure flight height-energy change rate, the current flight height and the current height-energy change rate.
The energy-height profile of the small-sinking-rate flight segment remains unchanged, namely:
Figure BDA0001845936000000132
the method for determining the flight height profile adjustment quantity delta h specifically comprises the following steps:
Figure BDA0001845936000000141
Figure BDA0001845936000000142
Δsh＝(1-ks)Δs,
wherein,. DELTA.hkΔ h is the adjustment of the current guidance period flight altitude profile, Δ hk-1The adjustment of the flight height profile of the previous guidance period, dt is the magnitude of the guidance period, gammah>0，λ>0，EnIs the current energy of the aircraft, Δ s is the range deviation of the aircraft, ksAdjusting distribution coefficient for voyage, k is more than or equal to 0sLess than or equal to 1; k when the flight path of the aircraft is in the second transition section or the small-sinking-rate flight sections＝0。
The step 412) is a method for determining the aircraft attack angle instruction of each section of track in real time according to the adjusted aircraft altitude characteristic profile, and specifically comprises the following steps:
αk＝αk-1+Δα，
Figure BDA0001845936000000143
Figure BDA0001845936000000144
wherein alpha iskFor angle of attack commands of the current guidance period, αk-1The angle of attack instruction of the previous guidance period, xi and omega are damping ratio of the guidance system and oscillation frequency of the guidance system, and nhFor real-time normal acceleration of the aircraft, hrefFor the height value corresponding to the current energy-height profile,
Figure BDA0001845936000000145
the height rate value corresponding to the current energy-height profile,
Figure BDA0001845936000000146
height acceleration value, n, corresponding to the current energy-height profilehIs the current normal acceleration, k, of the aircraftLIs the proportional coefficient of the lift coefficient and the attack angle, h is the aircraft height,
Figure BDA0001845936000000147
is the altitude rate of change of the aircraft, v is the speed of the aircraft, theta is the track inclination of the aircraft, sigmadIs the aircraft roll angle provided by the navigation system.
The step 41) is a method for determining the opening angle instruction of the resistance plate according to the distribution result of the range deviation, and specifically comprises the following steps:
Figure BDA0001845936000000151
Figure BDA0001845936000000152
Δsη＝ksΔs,
wherein eta isk-1For the drag-plate opening angle command of the previous guidance period, ηkFor the current guidance periodOpening angle command of force plate, gammaη>0, m is the aircraft mass, Δ s is the range deviation of the aircraft, ksAdjusting distribution coefficient for voyage, k is more than or equal to 0sLess than or equal to 1; k when the flight path of the aircraft is in the second transition section or the small-sinking-rate flight sections＝0。
The step 4) is a method for lateral guidance according to the lateral distance deviation of the aircraft, and specifically comprises the following steps: determining an aircraft roll angle instruction, and taking the determined aircraft roll angle instruction as lateral guidance information; the aircraft roll angle command is determined as follows:
Figure BDA0001845936000000153
wherein y is the lateral distance deviation of the aircraft;
Figure BDA0001845936000000154
is the lateral velocity of the aircraft; k is a radical ofzp>0,kzd>0。
5) And (3) transmitting the longitudinal guidance information and the lateral guidance information obtained in the step (4) to an aircraft control system for controlling the aircraft to fly and finishing the landing guidance work of the aircraft during approach.
Examples
The initial conditions of the aircraft are as follows: initial altitude 4025m, initial x position-11333 m, initial z position 100m, initial velocity 50m/s, initial heading angle 0 °.
The landing terminal conditions are: terminal altitude 0m, terminal x position 0m, terminal z position 0m, terminal speed 100m/s, terminal heading angle 0 °.
Step 1) determining a flight energy-height corridor;
11) solving the whole approach landing process according to the following formula, wherein the height lower limit h corresponding to each energy pointmin：
Figure BDA0001845936000000161
Wherein, mu is 3.9860×1014，R0＝6378375，qmax＝12000，ρ0＝1.1815，hs6370, specific energy-height lower limit hminAs shown in fig. 2.
12) Solving the whole approach landing process according to the following formula, wherein the height upper limit h corresponding to each energy pointmax
Figure BDA0001845936000000162
Wherein m is 3300, SrefSpecific energy-height upper limit h 5.453maxAs shown in fig. 2.
Step 2) judging whether the aircraft is positioned in a flight energy-height corridor or not according to the flight height of the aircraft, starting from approach landing guidance until the aircraft enters the energy-height corridor, wherein the flight section of the aircraft is a flight state adjusting section, and entering step 3); from the end of the flight state adjustment section to the arrival of the aircraft altitude at the airport altitude, the flight section of the aircraft is an altitude profile online adjustment section, and the step 4 is carried out
Step 3), performing longitudinal guidance and lateral guidance of a flight state adjusting section until the aircraft is positioned in a flight energy-height corridor, autonomously generating an energy-height characteristic profile, and entering step 4);
31) the aircraft lateral guidance method specifically comprises the following steps:
the roll angle command is held at zero, i.e., σ is 0.
32) The aircraft longitudinal guidance method specifically comprises the following steps:
η＝η0，
Figure BDA0001845936000000163
wherein eta is0＝42，qc＝3000，k＝0.01，
Figure BDA0001845936000000164
The opening angle command result of the drag plate is shown in FIG. 6, and the attack angle command result is shown inAs shown in fig. 3.
33) The energy-height characteristic profile autonomous generation method specifically comprises the following steps:
331) when E is less than or equal to Eq(Eq9500), the energy-height characteristic profile of the flight segment with small sinking rate is:
Figure BDA0001845936000000165
wherein E isf＝8674.1，hf＝0,
Figure BDA0001845936000000171
332) When E isq＜E≤ELWhen E is greaterL18000, the energy-height profile of the second transition segment is:
hL＝cL0+cL1E+cL2E2+cL3E3，
wherein, cL0＝338.0868、cL1＝-0.078948、cL2＝4.60412×10-6And cL3＝5.42317×10-12。
333) When E isL＜E≤EdWhen E is greaterdAt 35000, the energy-height characteristic profile of the iso-dynamic pressure section is:
Figure BDA0001845936000000172
wherein,
Figure BDA0001845936000000173
hd0＝2004.9。
334) when E isdWhen the energy-height characteristic section of the first transition section is less than E:
hb＝cb0+cb1E+cb2E2+cb3E3，
wherein, cb0＝85058.449、cb1＝6.52542、cb2＝1.66856×10-4And cb3＝1.3775519×10-9。
The energy-height profile is shown in fig. 2.
Step 4), carrying out height profile on-line adjustment section longitudinal guidance: according to the range deviation adjustment distribution law, the range deviation of the aircraft is distributed to a height profile for adjustment and a resistance plate opening angle for adjustment; adjusting the energy-height profile in real time by using a height profile adjusting law according to the flight distance deviation of the height profile responsible for adjustment; determining an attack angle instruction in real time by using a height tracking law according to the adjusted energy-height profile; and determining the opening angle of the resistance plate in real time by using a resistance plate opening angle guidance law according to the range deviation for adjusting the opening angle of the resistance plate. And simultaneously, carrying out lateral guidance of an online height profile adjusting section, and determining an aircraft roll angle instruction in real time by utilizing an aircraft lateral guidance law according to a lateral distance
41) According to the range deviation adjustment distribution law, the method for distributing the range deviation of the aircraft to the height profile responsible adjustment and the resistance plate opening angle responsible adjustment specifically comprises the following steps:
Figure BDA0001845936000000174
wherein, in the first transition section and the equal dynamic pressure section, k iss0.5; when the aircraft is in the second transition section and the flight section with small sinking rate, let ks＝0。
42) The method for adjusting the energy-height profile in real time by utilizing the height profile adjustment law according to the flight distance deviation of the height profile responsible for adjustment specifically comprises the following steps:
the height profile of the equal dynamic pressure flight section is adjusted in real time, the energy-height profile of the small-sinking rate flight section is kept unchanged, and the height profiles of the first transition section and the second transition section are correspondingly changed according to the real-time adjustment condition of the height profile of the equal dynamic pressure flight section.
After the height profile of the equal dynamic pressure flight section is adjusted by delta h, the form of the equal dynamic pressure flight height profile is as follows:
Figure BDA0001845936000000181
the energy-height profile of the small-sink-rate flight segment remains unchanged, namely:
Figure BDA0001845936000000182
after the equal dynamic pressure flight height profile is adjusted by delta h, the height profile of the second transition section is as follows:
hL_Δh＝cL0_Δh+cL1_ΔhE+cL2_ΔhE2+cL3_ΔhE3，
after the equal dynamic pressure flight height profile is adjusted by delta h, the height profile of the first transition section is as follows:
hb_Δh＝cb0_Δh+cb1_ΔhE+cb2_ΔhE2+cb3_ΔhE3，
at the current energy EnUnder the condition, the influence delta S of the altitude variation delta h on the voyagehSensitivity of voyage to altitude changes
Figure BDA0001845936000000183
Determined by the following formula.
Figure BDA0001845936000000184
μhThe results are shown in FIG. 7.
Determining the adjustment amount of the flight height profile in real time by using the following height profile adjustment law:
Figure BDA0001845936000000185
where dt is 0.02, gammah＝1，λ＝0.1。
43) According to the current energy-height profile, determining an attack angle instruction of the current guidance period in real time by using a height tracking law, wherein the specific method comprises the following steps:
the aircraft changes the lift coefficient by adjusting the attack angle, so that the normal acceleration adjustment is realized, the tracking guidance of the height is realized, and the attack angle instruction calculation formula is as follows:
Figure BDA0001845936000000191
where ω is 0.7 and ξ is 0.25. The angle of attack command results are shown in fig. 3.
44) The method for determining the opening angle of the resistance plate in real time by utilizing the guidance law of the opening angle of the resistance plate according to the course deviation of the adjustment of the opening angle of the resistance plate comprises the following steps:
at the current energy EnUnder the condition, the influence delta S of the altitude variation delta h on the voyagehThe sensitivity of the range to altitude changes is determined by the following equation.
Figure BDA0001845936000000192
As shown in particular in fig. 8.
The opening angle guidance law of the aircraft resistance plate is utilized to determine the opening angle of the aircraft resistance plate in real time:
Figure BDA0001845936000000193
wherein, γ η1. The flap opening angle command is shown in figure 6.
45) The method for determining the aircraft roll angle instruction in real time by utilizing the aircraft lateral guidance law according to the current lateral distance comprises the following specific calculation formula:
Figure BDA0001845936000000194
wherein k iszp＝0.0009，kzdThe roll angle command is as shown in fig. 4, 0.02Shown in the figure.
Those skilled in the art will appreciate that the details of the invention not described in detail in the specification are within the skill of those skilled in the art.

Claims (5)

1. An aerospace vehicle approach landing guidance method is characterized by comprising the following steps:
1) determining a flight energy-altitude corridor;
2) judging whether the aircraft is located in a flight energy-height corridor or not according to the flight height and the energy of the aircraft, and entering a step 3 if the aircraft is not located in the flight energy-height corridor; if the aircraft is located in the flight energy-height corridor, entering the step 4);
3) carrying out longitudinal guidance and lateral guidance on the aircraft until the flight height and the energy of the aircraft are positioned in a flight energy-height corridor, and entering the step 4);
4) performing longitudinal guidance according to the range deviation of the aircraft, performing lateral guidance according to the lateral distance deviation of the aircraft to obtain longitudinal guidance information and lateral guidance information, and entering the step 5);
5) transmitting the longitudinal guidance information and the lateral guidance information obtained in the step 4) to an aircraft control system for controlling the aircraft to fly and finishing the landing guidance work of the aircraft during approach;
the aircraft longitudinal guidance method in the step 3) specifically comprises the following steps: determining an attack angle instruction of the aircraft, keeping an initial value of a starting angle instruction of the resistance plate unchanged and the determined attack angle instruction of the aircraft as longitudinal guidance information, and transmitting the longitudinal guidance information to an aircraft control system to complete longitudinal guidance;
the attack angle command alpha of the aircraft is specifically as follows:
Figure FDA0003117915040000011
Figure FDA0003117915040000012
where m is the aircraft mass, v is the aircraft velocity, ρ is the atmospheric density, SrefIs the aircraft reference area, kCD_α_0Is the ratio of flight angle of attack to drag coefficient of the aircraft, q is the real-time dynamic pressure of the aircraft, q is the ratio of flight angle of attack to drag coefficient of the aircraftcIs the dynamic pressure threshold value of the aircraft, k is the guidance instruction gain coefficient,
Figure FDA0003117915040000013
in order to set the flight altitude-energy change rate, theta is the track inclination angle of the aircraft, and D is the resistance acceleration;
the method for performing longitudinal guidance according to the range deviation of the aircraft in the step 4) specifically comprises the following steps:
41) allocating the range deviation of the aircraft, and determining an attack angle instruction and a resistance plate opening angle instruction according to the allocation result of the range deviation;
42) taking the attack angle command and the opening angle command of the resistance plate determined in the step 41) as longitudinal guidance information;
the step 41) is a method for determining an attack angle instruction according to the distribution result of the range deviation, and specifically comprises the following steps:
411) performing trajectory planning, and sequentially dividing a flight trajectory into a first transition section, an equal dynamic pressure flight section, a second transition section and a small sinking rate flight section from large to small according to the flight energy of the aircraft; determining the aircraft height characteristic profile of each section of track;
412) adjusting the aircraft height characteristic profile determined in the step 411) in real time according to the distribution result of the range deviation; determining an aircraft attack angle instruction of each section of track in real time according to the adjusted aircraft height characteristic profile;
the aircraft height characteristic profile of each section of track determined in step 411) specifically includes:
A) the first transition section aircraft height characteristic profile hbThe method specifically comprises the following steps:
hb＝cb0+cb1E+cb2E2+cb3E3，
Ed<E，
where E is aircraft energy, cb0、cb1、cb2And cb3The characteristic section coefficient of the aircraft height of the first transition section; edDesigning an energy design value at the intersection point of the first transition section and the equal dynamic pressure flight section;
B) the height characteristic profile h of the aircraft in the equal dynamic pressure flight sectiondThe method specifically comprises the following steps:
Figure FDA0003117915040000021
Figure FDA0003117915040000022
EL<E≤Ed，
wherein h isd0To the starting height of the constant-pressure flight section, ELDesigning an energy design value at the intersection of the constant-pressure flight section and the second transition section;
C) the second transition section aircraft height characteristic profile hLThe method specifically comprises the following steps:
hL＝cL0+cL1E+cL2E2+cL3E3,
Eq<E≤EL，
wherein, cL0、cL1、cL2And cL3For the second transition aircraft altitude characteristic profile coefficient, EqThe energy design value of the intersection point of the second transition section and the small-sinking-rate flight section is obtained;
D) the height characteristic profile h of the small-sinking-rate flight segment aircraftqThe method specifically comprises the following steps:
Figure FDA0003117915040000031
Figure FDA0003117915040000032
wherein h isfNominal altitude for aircraft terminal, EfNominal energy, g, for aircraft terminalsfFor nominal acceleration of gravity, v, of the terminalfFor the nominal speed of the terminal to be,
Figure FDA0003117915040000033
altitude-energy rate of change, θ, for flight segment with small sinking rateqFlight path inclination for flight section with small sinking rate, DqAcceleration of resistance for flight section of small sinking rate, CD_qIs a coefficient of resistance;
the step 412) of adjusting the aircraft altitude characteristic profile in real time includes:
determining the adjustment quantity delta h of a flight height profile, adjusting the height characteristic profile of the isokinetic pressure flight section aircraft in real time according to the delta h to obtain the adjusted height characteristic profile of the isokinetic pressure flight section aircraft, and adjusting the height characteristic profiles of the first transition section and the second transition section aircraft according to the real-time adjustment result of the height characteristic profile of the isokinetic pressure flight section aircraft to obtain the adjusted height characteristic profiles of the first transition section and the second transition section aircraft; the altitude characteristic profile of the small-sinking-rate flight segment aircraft is kept unchanged;
the height characteristic profile h of the aircraft of the adjusted equal dynamic pressure flight sectiond_ΔhFirst transition section aircraft altitude characteristic profile hb_ΔhAnd a characteristic profile h of the aircraft altitude of the second transition sectionL_ΔhThe method specifically comprises the following steps:
Figure FDA0003117915040000034
hL_Δh＝cL0_Δh+cL1_ΔhE+cL2_ΔhE2+cL3_ΔhE3，
hb_Δh＝cb0_Δh+cb1_ΔhE+cb2_ΔhE2+cb3_ΔhE3，
wherein, cL0_Δh、cL1_Δh、cL2_Δh、cL3_ΔhFor the adjusted second transition section aircraft altitude characteristic profile coefficient, cb0_Δh、cb1_Δh、cb2_ΔhAnd cb3_ΔhThe adjusted first transition section aircraft height characteristic profile coefficient is obtained;
the method for determining the flight height profile adjustment quantity delta h specifically comprises the following steps:
Figure FDA0003117915040000041
Figure FDA0003117915040000042
Δsh＝(1-ks)Δs,
wherein,. DELTA.hkΔ h is the adjustment of the current guidance period flight altitude profile, Δ hk-1The adjustment of the flight height profile of the previous guidance period, dt is the magnitude of the guidance period, gammah>0，λ>0，EnIs the current energy of the aircraft, Δ s is the range deviation of the aircraft, ksThe distribution coefficient is adjusted for the flight distance, and k is the flight path of the aircraft in the second transition section or the flight section with small sinking rates0; k when the flight path of the aircraft is located in the first transition section or the equal dynamic pressure flight sectionsIs not zero;
the step 412) is a method for determining the aircraft attack angle instruction of each section of track in real time according to the adjusted aircraft altitude characteristic profile, and specifically comprises the following steps:
αk＝αk-1+Δα，
Figure FDA0003117915040000043
Figure FDA0003117915040000044
wherein alpha iskFor angle of attack commands of the current guidance period, αk-1The angle of attack instruction of the previous guidance period, xi and omega are damping ratio of the guidance system and oscillation frequency of the guidance system, and nhFor real-time normal acceleration of the aircraft, hrefFor the height value corresponding to the current energy-height profile,
Figure FDA0003117915040000045
the height rate value corresponding to the current energy-height profile,
Figure FDA0003117915040000051
height acceleration value, k, corresponding to the current energy-height profileLIs the proportional coefficient of the lift coefficient and the attack angle, h is the aircraft height,
Figure FDA0003117915040000052
is the altitude rate of change of the aircraft, v is the speed of the aircraft, theta is the track inclination of the aircraft, sigmadIs the aircraft roll angle provided by the navigation system.
2. The aerospace vehicle approach landing guidance method of claim 1, wherein the step 1) of determining a flight energy-altitude corridor comprises determining a lower altitude limit h of the flight energy-altitude corridorminRelationship to aircraft flight energy and upper altitude limit h of flight energy-altitude corridormaxThe relation with the flight energy of the aircraft is as follows:
11) a lower altitude limit h of the flight energy-altitude corridorminThe relationship to the flight energy of the aircraft is as follows:
Figure FDA0003117915040000053
wherein E is the flight energy of the aircraft, mu is the gravitational constant, R0Is the airport plane earth radius, qmaxUpper limit of dynamic pressure of aircraft, ρ0Is the atmospheric density, h, of the aircraft landing site levelsIs an atmospheric characteristic constant;
12) an upper altitude limit h of the flight energy-altitude corridormaxThe relationship to the flight energy of the aircraft is as follows:
Figure FDA0003117915040000054
where m is the aircraft mass, SrefFor aircraft reference area, CL maxThe coefficient of lift is the coefficient of lift at the maximum lift-drag ratio of the aircraft.
3. The aerospace vehicle approach landing guidance method according to claim 1, wherein the method for guidance of the aircraft in step 3) is specifically: and keeping the inclination angle of the aircraft to be zero as lateral guidance information, and transmitting the lateral guidance information to an aircraft control system to finish lateral guidance.
4. The aerospace vehicle approach landing guidance method according to claim 1, wherein the step 41) is a method for determining a flap opening angle command according to the distribution result of the range deviation, and specifically comprises the following steps:
Figure FDA0003117915040000061
Figure FDA0003117915040000062
Δsη＝ksΔs,
wherein eta isk-1Drag plate for last guidance periodOpening angle command, ηkFor the current guidance period flap opening angle command, gammaη>0, m is the aircraft mass, Δ s is the range deviation of the aircraft, ksAdjusting distribution coefficients for the voyage; k when the flight path of the aircraft is in the second transition section or the small-sinking-rate flight sections0; k when the flight path of the aircraft is located in the first transition section or the equal dynamic pressure flight sectionsIs not zero.
5. The aerospace vehicle approach landing guidance method according to claim 4, wherein the step 4) is a method for performing lateral guidance according to lateral distance deviation of the aircraft, and specifically comprises: determining an aircraft roll angle instruction, and taking the determined aircraft roll angle instruction as lateral guidance information; the aircraft roll angle command is determined as follows:
Figure FDA0003117915040000063
wherein y is the lateral distance deviation of the aircraft;
Figure FDA0003117915040000064
is the lateral velocity of the aircraft; k is a radical ofzp>0,kzd>0。
CN201811270811.8A 2018-10-29 2018-10-29 Aerospace vehicle approach landing guidance method Active CN109240335B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201811270811.8A CN109240335B (en) 2018-10-29 2018-10-29 Aerospace vehicle approach landing guidance method

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201811270811.8A CN109240335B (en) 2018-10-29 2018-10-29 Aerospace vehicle approach landing guidance method

Publications (2)

Publication Number Publication Date
CN109240335A CN109240335A (en) 2019-01-18
CN109240335B true CN109240335B (en) 2021-09-07

Family

ID=65079194

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201811270811.8A Active CN109240335B (en) 2018-10-29 2018-10-29 Aerospace vehicle approach landing guidance method

Country Status (1)

Country Link
CN (1) CN109240335B (en)

Families Citing this family (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN110347179B (en) * 2019-06-25 2022-06-03 湖北航天技术研究院总体设计所 Transverse flight energy management method of unpowered aircraft
CN110779512B (en) * 2019-10-22 2023-08-04 中国人民解放军63660部队 Flight test route planning method for accuracy identification of measurement and control equipment
CN112256045A (en) * 2020-05-11 2021-01-22 北京航天自动控制研究所 Rocket landing trajectory planning method and device
CN114200829B (en) * 2021-11-09 2023-06-23 西北工业大学 High-precision speed control method for supersonic large maneuvering target based on pseudo closed loop
CN114167888B (en) * 2021-11-19 2023-06-20 湖北航天技术研究院总体设计所 Method for controlling tail end position and speed of gliding hypersonic aircraft
CN114610056A (en) * 2022-02-24 2022-06-10 北京空天技术研究所 Aerospace vehicle landing segment online guidance method based on elliptical trajectory

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN102682625A (en) * 2011-02-07 2012-09-19 霍尼韦尔国际公司 Systems and methods for providing ITP clearance information
CN108398883A (en) * 2018-02-27 2018-08-14 北京控制工程研究所 A kind of RLV approaches track rapid deduction and determining method

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104731104B (en) * 2015-03-09 2015-11-04 北京航天自动控制研究所 A kind of longitudinal guidance method of hypersonic aircraft gliding flight section
KR101658019B1 (en) * 2015-08-07 2016-09-20 한국항공우주연구원 Flight guidance method of high-altitude pilotless plane
FR3064762B1 (en) * 2017-04-04 2020-07-31 Thales Sa MANAGEMENT OF THE DESCENT PHASE OF AN AIRCRAFT
CN107941087B (en) * 2017-10-18 2019-08-06 北京航空航天大学 A kind of superb steady gliding reentry guidance method of high lift-drag ratio based on resistance profiles

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN102682625A (en) * 2011-02-07 2012-09-19 霍尼韦尔国际公司 Systems and methods for providing ITP clearance information
CN108398883A (en) * 2018-02-27 2018-08-14 北京控制工程研究所 A kind of RLV approaches track rapid deduction and determining method

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
基于在线轨迹生成的末端能量管理段复合制导技术;刘智勇 等;《战术导弹技术》;20170831(第4期);第66-72页 *
重复使用运载器末端区域能量管理与自动着陆技术研究;孙春贞;《中国博士学位论文全文数据库 工程科技Ⅱ辑》;20090515(第05期);第1-150页 *

Also Published As

Publication number Publication date
CN109240335A (en) 2019-01-18

Similar Documents

Publication Publication Date Title
CN109240335B (en) Aerospace vehicle approach landing guidance method
CN110471450B (en) Method for directly planning reentry trajectory in altitude velocity profile
CN107643763B (en) Airplane unpowered return energy/track comprehensive control method
CN102759928B (en) Control method for flight path of airship on stratosphere
CN102393200B (en) General inertial navigation test method based on flight simulation
JP5473672B2 (en) Maximizing aircraft predictability with a continuous descent approach
Yu et al. Analytical entry guidance for coordinated flight with multiple no-fly-zone constraints
CN100541372C (en) Automatic homing control method under a kind of unmanned vehicle engine involuntary stoppage
CN111026157B (en) Intelligent aircraft guiding method based on reward remodeling reinforcement learning
CN109240323B (en) Aerospace vehicle reentry guidance method capable of analyzing structure in real time
CN108153323B (en) A kind of high-altitude unmanned vehicle high-precision reentry guidance method
CN105759609B (en) A kind of carrier-borne aircraft autonomous landing on the ship method based on explicit non-linear Model Predictive Control
CN113900448B (en) Aircraft prediction correction composite guidance method based on sliding mode interference observer
CN114065398B (en) Flight performance calculation method for high-aspect-ratio flexible aircraft
CN106444822A (en) Space vector field guidance based stratospheric airship&#39;s trajectory tracking control method
CN109703769B (en) Air refueling docking control method based on preview strategy
CN112744367A (en) Guidance control method and system for vertical launching and ignition phase in near space
CN103578299A (en) Method for simulating flight process of aircraft
CN107315419B (en) Three-section control strategy for realizing long-distance fixed-point hovering of airship
Hartmann et al. Control of departure and approach maneuvers of tiltwing VTOL aircraft
CN105759630A (en) Airplane 4D flight path simulation system and method based on fuzzy PID control
CN116301028A (en) Multi-constraint on-line flight trajectory planning middle section guiding method based on air suction hypersonic speed platform
CN113342039A (en) Landing control method for long-endurance scouting and batting type sliding running and take-off and landing unmanned aerial vehicle
Liu et al. Research on Trajectory Generation and Optimization in Continuous Descent Operations
Atmeh et al. A nonlinear automatic landing control system for a UAV

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant