Aircraft altitude speed straight climb control method

Abstract

The present invention discloses a method for controlling the altitude and speed of an aircraft in a straight line, including: associating the velocity inclination angle and the velocity inclination change rate with the flight altitude and the flight speed, constructing a first association model, so that the aircraft accelerates the climb and turns flat at the end of the climb section; associating the flight speed and the velocity change rate with the flight altitude and the velocity inclination angle, constructing a second association model, so that the altitude and speed of the aircraft at the end of the climb section gradually tend to cruise; obtaining the velocity inclination change rate and the velocity change rate based on the first association model and the second association model; constructing a flight trajectory model and a thrust specific impulse model, and obtaining the real-time angle of attack instruction and fuel flow rate instruction of the aircraft based on the velocity inclination change rate and the velocity change rate. The present invention is applied to the field of trajectory control, making the state variables and trajectory forms of the aircraft more reasonable, satisfying the dynamic pressure and overload distribution constraints of the aircraft, and making the control quantity easy to solve, which is of great significance for the rapid planning of flight missions.

Description

Technical Field

[0001] The invention relates to the technical field of aircraft trajectory planning, and in particular to an aircraft altitude speed linear climbing control method. Background Art

[0002] The advantages of ramjet engine in high specific impulse and thrust-to-drag ratio make it a power device for aircraft. In order to give full play to the performance advantages of ramjet engine in high-altitude and high-speed flight, it is usually accelerated to a suitable working speed by booster, and then ignited to quickly climb to the cruise working point. The climbing process must meet the terminal constraints, and the engine performance is also coupled with trajectory parameters such as altitude, dynamic pressure, and angle of attack, making the climbing stage a difficult point in trajectory planning for ramjet engine aircraft.

[0003] The mission envelope of an air-breathing aircraft is large. The ramjet engine is coupled to the flight state. The propulsion characteristics of the aircraft at different altitudes and speeds vary greatly. The selection of the flight state has a significant impact on the model characteristics. The flight performance at the beginning and end of the climb phase will have an impact on the overall flight mission. The design of the climb trajectory needs to solve how to make the missile climb from the ramjet engine start-up point to the cruise state, which is of great value to improving the performance of the aircraft and expanding the application field. Summary of the invention

[0004] In view of the deficiencies in the above-mentioned prior art, the present invention provides a method for controlling the altitude and speed of an aircraft in a straight line, so that the state variables and trajectory forms of the aircraft are more reasonable, while satisfying the dynamic pressure and overload distribution constraints of the aircraft. The control quantity is easy to solve, which is of great significance for the rapid planning of flight missions.

[0005] To achieve the above object, the present invention provides a method for controlling the altitude and speed of an aircraft in a straight line, comprising the following steps:

[0006] Step 1, associating the velocity inclination angle and velocity inclination change rate of the aircraft with the flight altitude and the flight speed, and constructing a first association model, so that the aircraft accelerates the climb and turns flat at the end of the climb segment;

[0007] Step 2, associating the flight speed and speed change rate of the aircraft with the flight altitude and speed inclination, and constructing a second association model, so that the altitude and speed of the aircraft at the end of the climb phase gradually approach the cruising state;

[0008] Step 3, obtaining the real-time velocity inclination change rate and velocity change rate of the aircraft based on the first association model and the second association model;

[0009] Step 4, constructing a flight trajectory model and a thrust-specific impulse model of the aircraft, and obtaining the real-time angle of attack instruction and fuel flow rate instruction of the aircraft based on the real-time velocity inclination angle change rate, velocity change rate, the flight trajectory model, and the thrust-specific impulse model of the aircraft.

[0010] Compared with the prior art, the present invention has the following beneficial technical effects:

[0011] The present invention establishes a first association model to accelerate the climb of an aircraft and level off at the end of a climb segment, and simultaneously establishes a second association model to gradually approach the altitude and speed of the aircraft at the end of the climb segment to cruise. The real-time angle of attack command and fuel flow rate command of the aircraft are calculated based on the first association model and the second association model, and the angle of attack and fuel flow rate of the climb segment of the ramjet aircraft are adjusted simultaneously, so that the altitude of the aircraft climbing terminal point is adjustable, and there is a corresponding speed for a certain altitude, so that the state variables and trajectory forms of the aircraft are more reasonable, and the dynamic pressure and overload distribution constraints of the aircraft are satisfied at the same time, and the control quantity is simple to solve, which is of great significance for the rapid planning of flight missions. BRIEF DESCRIPTION OF THE DRAWINGS

[0012] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the structures shown in these drawings without paying creative work.

[0013] Figure 1 Flow chart of a method for controlling the height and speed of an aircraft in a straight line climb according to an embodiment of the present invention;

[0014] Figure 2 Schematic diagram of a height speed straight climbing mode in an embodiment of the present invention;

[0015] Figure 3 Schematic diagram of ramjet engine performance in an embodiment of the present invention, wherein: (a) is a graph showing the variation of specific impulse with height, (b) is a graph showing the variation of specific impulse with speed, (c) is a graph showing the variation of specific impulse with fuel flow rate, and (d) is a graph showing the variation of specific impulse with angle of attack;

[0016] Figure 4 Schematic diagram of the change of control quantity and state during the flight of the aircraft in the embodiment of the present invention, wherein: (a) is the altitude range curve, (b) is the Mach number time curve, (c) is the speed inclination time curve, (d) is the angle of attack instruction, (e) is the fuel flow rate instruction, and (f) is the mass change curve.

[0017] The realization of the purpose, functional features and advantages of the present invention will be further explained in conjunction with embodiments and with reference to the accompanying drawings. DETAILED DESCRIPTION

[0018] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0019] In addition, in the present invention, descriptions such as "first", "second", etc. are only used for descriptive purposes and cannot be understood as indicating or implying their relative importance or implicitly indicating the number of the indicated technical features. Therefore, the features defined as "first" or "second" may explicitly or implicitly include at least one of the features. In the description of the present invention, the meaning of "plurality" is at least two, such as two, three, etc., unless otherwise clearly and specifically defined.

[0020] In addition, the technical solutions between the various embodiments of the present invention can be combined with each other, but it must be based on the fact that ordinary technicians in the field can implement it. When the combination of technical solutions is contradictory or cannot be implemented, it should be deemed that such combination of technical solutions does not exist and is not within the scope of protection required by the present invention.

[0021] The climbing phase is when the booster engine finishes working and the ramjet engine takes over, climbing from the current altitude speed to the cruising altitude speed. For the climbing phase, the purpose is to climb to the cruising altitude smoothly, taking into account both the dynamic pressure constraint and the overload constraint. Specifically, the control command should balance climbing and acceleration. If the angle of attack is too small, a large part of the aircraft's ability is used for acceleration, which may cause excessive axial overload. On the other hand, if the speed is increased very quickly at low altitude, the dynamic pressure may also exceed the upper limit of the dynamic pressure, which will cause damage to the structure of the aircraft. If the angle of attack is too small, a large part of the cruise aircraft's ability is used to lift its head and climb, and the speed inclination angle change rate is large, which may cause the normal overload to exceed the constraint. On the other hand, the speed increases slowly and the altitude increases quickly, which may cause a sharp decrease in dynamic pressure, affecting the air intake. The working performance of the ramjet engine may be greatly reduced, and there may even be a risk of flameout. Based on this, the present embodiment discloses a method for controlling the altitude and speed of an aircraft in a straight line climb. By simultaneously adjusting the angle of attack and the fuel flow rate of a ramjet aircraft in a climb segment, the altitude of the aircraft's climb terminal point can be adjusted, and there is a corresponding speed for a certain altitude, so that the state variables and trajectory forms of the aircraft are reasonable and the dynamic pressure and overload distribution constraints are satisfied at the same time.

[0022] refer to Figure 1 The aircraft altitude speed straight climb control method in this embodiment specifically includes the following steps:

[0023] Step 1, associating the velocity inclination angle and velocity inclination change rate of the aircraft with the flight altitude and the flight speed, and constructing a first association model, so that the aircraft accelerates the climb and turns flat at the end of the climb segment;

[0024] Step 2, associating the flight speed and speed change rate of the aircraft with the flight altitude and speed inclination, and constructing a second association model, so that the altitude and speed of the aircraft at the end of the climb phase gradually approach the cruising state;

[0025] Step 3, obtaining the real-time velocity inclination change rate and velocity change rate of the aircraft based on the first correlation model and the second correlation model;

[0026] Step 4, construct the flight trajectory model and thrust-specific impulse model of the aircraft, and obtain the real-time angle of attack command and fuel flow rate command of the aircraft based on the real-time velocity inclination angle change rate, velocity change rate, flight trajectory model and thrust-specific impulse model of the aircraft.

[0027] In order to accelerate the aircraft to climb and level off at the end of the climbing segment, the first association model in this embodiment is specifically constructed as follows:

[0028] ;

[0029] in, is the velocity inclination angle, is the velocity angle change rate, is the initial climbing speed inclination angle, is the flight altitude, To climb to the initial flight altitude, is the climbing end altitude, For the flight speed.

[0030] In order to make the height and speed of the aircraft slowly approach cruising at the end of the climbing phase, the second association model in this embodiment is specifically constructed as follows:

[0031] ;

[0032] in, is the speed change rate, is the speed control coefficient, is the flight altitude, It is a constant determined by the flight altitude and flight speed after the boost is completed, that is, the flight speed is expressed as a linear function of the flight altitude, and the speed change rate is set as a function of the speed inclination angle.

[0033] At a known real-time flight altitude , initial climb speed inclination , climbing end height , speed control coefficient And constant On the premise of , by combining the equations of the first association model and the second association model, the current flight speed, speed inclination, flight speed change rate and speed inclination change rate of the aircraft can be calculated.

[0034] The fuel flow rate of the ramjet gas generator is adjustable, and the engine specific impulse can be expressed as a four-dimensional function of Mach number, flight altitude, gas generator fuel flow rate, and flight angle of attack, namely:

[0035] ;

[0036] in, is the specific impulse, is the incoming flow Mach number, is the gas generator fuel flow rate, is the angle of attack;

[0037] Therefore, when the real-time flight speed, speed inclination, flight speed change rate and speed inclination change rate of the aircraft are known, by constructing the aircraft's flight trajectory model and thrust specific impulse model, the aircraft's real-time angle of attack command and fuel flow rate command can be calculated to complete the aircraft's altitude and speed linear climb control.

[0038] The flight trajectory model in this embodiment is specifically:

[0039] ;

[0040] in, For thrust, is the mass of the aircraft, is the aerodynamic drag, is the acceleration due to gravity, is the aerodynamic lift, is the flight speed, is the distance from the center of the earth, is the rate of change of track direction angle, is the latitude, is the track direction angle, is the rate of change of longitude, is the latitude change rate, is the rate of change of the distance from the center of the earth.

[0041] The geometrical dimensions and other structural parameters of the ramjet engine will affect the specific impulse. However, in the design and calculation stage of the external flight trajectory, the geometrical dimensions and other structural parameters of the ramjet engine should have been determined. Therefore, the thrust and specific impulse of the ramjet engine can be simplified and modeled. The construction process of the thrust specific impulse model in this embodiment includes:

[0042] First, the inlet entrance velocity coefficient is calculated based on the incoming flow Mach number, which is:

[0043] ;

[0044] in, is the inlet velocity coefficient, is the incoming flow specific heat ratio, is the incoming flow Mach number;

[0045] Secondly, the static temperature of the incoming flow is obtained based on the current flight altitude, and the total temperature of the incoming flow is calculated based on the static temperature of the incoming flow and the inlet velocity coefficient, which is:

[0046] ;

[0047] ;

[0048] in, For the coming flow of calm and warmth, is the total temperature of the incoming flow, is the gas function;

[0049] Then, calculate the intake air flow rate as:

[0050] ;

[0051] in, is the intake air flow rate, is the incoming flow density, is the speed of sound, is the inlet area of ​​the air intake duct, is the flow coefficient;

[0052] Then, the air-fuel ratio is calculated based on the intake air flow rate as:

[0053] ;

[0054] in, is the air-fuel ratio, is the fuel flow rate;

[0055] Then, the constant-pressure specific heat capacity and gas constant of the mixed gas in the afterburner are calculated based on the air-fuel ratio:

[0056] ;

[0057] in, is the specific heat capacity of the mixed gas at constant pressure, is the constant pressure specific heat capacity of the rich fuel gas, is the constant pressure specific heat capacity of the air at the afterburner inlet, is the gas constant of the mixed gas, is the rich fuel gas constant, is the air gas constant of the afterburner chamber;

[0058] The total temperature of the afterburner outlet gas, the heating ratio of the afterburner, the mass increase coefficient and the comprehensive parameters of the afterburner are further calculated as follows:

[0059] ;

[0060] ;

[0061] ;

[0062] ;

[0063] ;

[0064] in, is the mixed gas fuel flow rate, is the total temperature of the mixed gas, is the total temperature of the rich fuel gas, is the calorific value released by the secondary combustion of the rich fuel gas in the afterburning chamber, is the heating ratio of the afterburning chamber, is the total air temperature at the afterburner inlet, is the mass increase coefficient, is the comprehensive parameter of the afterburner chamber, It is a combination parameter with no real meaning. is the ratio of the specific heats of the mixed gases, is the incoming gas constant;

[0065] The flow in the nozzle is regarded as an isentropic steady flow, and the nozzle outlet flow function is:

[0066] ;

[0067] in, is the nozzle outlet flow function, is the nozzle exit velocity coefficient, is the nozzle throat area, is the nozzle outlet area, and the nozzle outlet velocity coefficient can be obtained ;

[0068] Ignoring additional drag, the thrust coefficient of the ramjet is:

[0069] ;

[0070] ;

[0071] in, is the ramjet thrust coefficient, is the inlet area of ​​the air intake, is the aerodynamic function;

[0072] Finally, the thrust specific impulse model is obtained as follows:

[0073] ;

[0074] ;

[0075] in, For thrust, is the dynamic pressure, For specific impulse.

[0076] At this point, the thrust-specific impulse model has been established. The relationship between specific impulse and altitude, Mach number, fuel flow rate, and angle of attack is as follows: Figure 3 By substituting the real-time flight speed, speed inclination, flight speed change rate and speed inclination change rate of the aircraft obtained according to the first association model and the second association model into the flight trajectory model and the thrust specific impulse model, the real-time angle of attack command and fuel flow rate command of the aircraft can be obtained to complete the aircraft height speed linear climb control.

[0077] In this embodiment, in order to ensure that the solid ramjet engine works in the best state, the constant altitude and constant speed strategy is adopted in the cruise phase, which consists of two conditions: constant altitude and constant speed, namely:

[0078] ;

[0079] ;

[0080] in, For thrust, is the aerodynamic drag, For aerodynamic lift.

[0081] The following is a further description of the aircraft altitude speed linear climb control method in this embodiment in conjunction with a specific simulation example. The simulation parameters are set as shown in Table 1.

[0082] Table 1

[0083]

[0084] The linear relationship between aircraft altitude and speed is as follows: Figure 2 As shown in , the dynamic pressure constraint of the aircraft can be satisfied by setting a reasonable speed control coefficient. Figure 4 (a) It can be seen that the aircraft successfully completed the climb and turned to enter the cruise phase. Under the same initial conditions, the aircraft climbed to different altitudes and maintained a fixed altitude during the cruise phase. Obviously, the higher the climb altitude, the longer the climb phase time and the longer the total flight time. Figure 4(b) It can be seen that due to the continuous increase in the altitude of the climb stage, the speed continues to increase under the altitude speed linear climb control method, reducing the possibility of engine flameout due to the speed reduction. Due to different cruising altitudes, the cruising speed is also different. The higher the climbing altitude, the greater the terminal speed. The aircraft is always accelerating from launch to the end of the climb, and in the later stage of the climb, the altitude change is small, and the corresponding speed change is gentle, which is conducive to a smooth transition to the cruise stage.

[0085] Depend on Figure 4 (c) It can be seen that after the boost ends, the speed angle begins to decrease, and the speed of change is related to the climb time. In the late climb, the change is relatively gentle, and it smoothly turns into the cruise phase. Figure 4 (d) It can be seen that the change trend of the angle of attack command is the same, because the initial speed inclination is large, and a negative angle of attack is adopted to reduce the speed inclination. Then it gradually increases, which is related to the change in altitude. After the altitude increases rapidly, the dynamic pressure decreases rapidly. In order to ensure sufficient lift to continue climbing, the angle of attack continues to increase to compensate for the trend of reduced lift caused by the reduction in dynamic pressure.

[0086] Depend on Figure 4 (e) It can be seen that when climbing to different altitudes, the fuel flow rate is different, and the overall trend is decreasing. This corresponds to the change in Mach number. In the early stage of climbing, a larger fuel flow rate is required to continue climbing. In the later stage of climbing, the altitude speed changes slowly, and there is little demand for acceleration, so the fuel flow rate continues to decrease. At the same time, the aircraft has different altitudes and speeds. The different altitudes are caused by the tilt control strategy, and the different speeds are caused by different fuel flow rates. Figure 4 (f) It can be seen that the dynamic pressure in the boost phase increases rapidly due to acceleration. The dynamic pressure decreases after the stage change because the altitude increases and the atmospheric density decreases. Although the speed is increasing, the effect of the density reduction is greater than the effect of the speed increase, causing the dynamic pressure to eventually decrease. Due to the different speeds at the end of the climb, the final dynamic pressure is different, and the dynamic pressure decreases with increasing altitude.

[0087] The above description is only a preferred embodiment of the present invention, and does not limit the patent scope of the present invention. All equivalent structural changes made by using the contents of the present invention specification and drawings under the inventive concept of the present invention, or directly / indirectly applied in other related technical fields are included in the patent protection scope of the present invention.

Claims

1. A method for controlling the altitude and speed of an aircraft in a straight line, characterized in that: The steps include: Step 1: Associating the velocity inclination angle and velocity inclination change rate of the aircraft with the flight altitude and the flight speed, and constructing a first association model to accelerate the aircraft to climb and level off at the end of the climb segment. Specifically, the first association model is: in, is the velocity inclination angle, is the velocity angle change rate, is the initial climbing speed inclination angle, is the flight altitude, To climb to the initial flight altitude, is the climbing end altitude, is the flight speed; Step 2, the flight speed and speed change rate of the aircraft are associated with the flight altitude and speed inclination, and a second association model is constructed so that the altitude and speed of the aircraft at the end of the climb phase gradually approach the cruising altitude. Specifically, the second association model is: in, is the flight speed, is the speed change rate, is the speed control coefficient, is the flight altitude, is a constant value, is the velocity inclination angle; Step 3, obtaining the real-time velocity inclination change rate and velocity change rate of the aircraft based on the first association model and the second association model; Step 4, constructing a flight trajectory model and a thrust-specific impulse model of the aircraft, and obtaining the real-time angle of attack instruction and fuel flow rate instruction of the aircraft based on the real-time velocity inclination angle change rate, velocity change rate, the flight trajectory model, and the thrust-specific impulse model of the aircraft.

2. The method for controlling the aircraft's altitude and speed in a straight line according to claim 1, characterized in that: In step 4, the flight trajectory model is specifically: in, is the speed change rate, For thrust, is the mass of the aircraft, is the angle of attack, is the aerodynamic drag, is the acceleration due to gravity, is the velocity inclination angle, is the velocity angle change rate, L is the aerodynamic lift, v is the flight speed, r is the distance from the center of the earth, is the rate of change of track direction angle, is the latitude, is the track direction angle, is the rate of change of longitude, is the latitude change rate, is the rate of change of the distance from the center of the earth.

3. The method for controlling the aircraft's altitude and speed in a straight line according to claim 1, characterized in that: In step 4, the construction process of the thrust specific impulse model includes: The inlet entrance velocity coefficient is calculated based on the incoming flow Mach number: in, is the inlet velocity coefficient, is the incoming flow specific heat ratio, is the incoming flow Mach number; The static temperature of the incoming flow is obtained based on the current flight altitude, and the total temperature of the incoming flow is calculated based on the static temperature of the incoming flow and the inlet velocity coefficient, which is: in, For the coming flow of calm and warmth, is the total temperature of the incoming flow, is the gas function; Calculate the intake air flow rate as: in, is the intake air flow rate, is the incoming flow density, a is the speed of sound, is the inlet area of ​​the air intake duct, is the flow coefficient; The air-fuel ratio is calculated based on the intake air flow rate: in, is the air-fuel ratio, is the fuel flow rate; Based on the air-fuel ratio, the constant-pressure specific heat capacity and gas constant of the mixed gas in the afterburner are calculated as follows: in, is the specific heat capacity of the mixed gas at constant pressure, is the constant pressure specific heat capacity of the rich fuel gas, is the constant pressure specific heat capacity of the air at the afterburner inlet, is the gas constant of the mixed gas, is the rich fuel gas constant, is the gas constant of the air at the afterburner inlet; The total temperature of the afterburner outlet gas, the heating ratio of the afterburner, the mass increase coefficient and the comprehensive parameters of the afterburner are further calculated as follows: in, is the mixed gas fuel flow rate, is the total temperature of the mixed gas, is the total temperature of the rich fuel gas, Q is the calorific value released by the secondary combustion of the rich fuel gas in the afterburning chamber, is the heating ratio of the afterburning chamber, is the total air temperature at the afterburner inlet, is the mass increase coefficient, is the comprehensive parameter of the afterburner chamber, is the combined parameter, is the ratio of specific heats of the mixed gas, is the incoming gas constant; The flow in the nozzle is regarded as an isentropic steady flow, and the nozzle outlet flow function is: in, is the nozzle outlet flow function, is the nozzle exit velocity coefficient, is the nozzle throat area, is the nozzle outlet area; Ignoring additional drag, the thrust coefficient of the ramjet is: in, is the ramjet thrust coefficient, is the inlet area of ​​the air intake duct, is the aerodynamic function; Finally, the thrust specific impulse model is obtained as follows: in, F For thrust, q is the dynamic pressure, For specific impulse.

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