Two-stage horizontal take-off and landing sub-orbital aircraft appearance aerodynamic trajectory integrated design method

Abstract

The invention discloses a two-stage horizontal take-off and landing sub-orbital aircraft appearance aerodynamic trajectory integrated design method. Full trajectory design of the horizontal take-off and landing sub-orbital aircraft in a two-stage rocket power mode is completed in a trajectory splicing mode, height-attack angle and speed-attack angle node control schemes are adopted in the large-attack-angle climbing sections I and II correspondingly, and horizontal accelerated flight of the aircraft in the active section is achieved; considering the characteristic that the aerodynamic design and the trajectory scheme of the sub-orbital aircraft are tightly coupled, the aerodynamic configuration and trajectory integrated optimization design is adopted, and the aerodynamic configuration and trajectory integrated optimization scheme of the aircraft is simplified into the inner layer optimization of the sub-orbital aircraft profile parameters and the outer layer optimization of the full trajectory scheme optimization. The hour-level 10000-kilometer remote top-speed transportation task under the two-stage rocket power mode is achieved, the time of the remote top-speed transportation task is greatly shortened, and a guarantee is provided for the remote top-speed transportation task.

Description

Technical Field

[0001] The present invention relates to the technical field of aerospace vehicle design, and in particular to a method for integrating the shape, aerodynamics and ballistics of a two-stage horizontal take-off and landing suborbital vehicle. Background Art

[0002] The suborbital long-distance transportation system utilizes aerospace technology to achieve ultra-fast long-distance transportation, typically reaching tens of thousands of kilometers per hour. With the ability to reach long distances within an hour, it could spur emerging industries such as one-hour transportation logistics and mass space tourism. It holds practical social, scientific, economic, and military applications in supporting my country's development into a powerhouse in science and technology, aerospace, and transportation. Suborbital vehicle design is crucial for this system, and relatively mature research cases exist internationally. Three types of suborbital long-distance transportation systems are available based on their takeoff and landing methods: the vertical takeoff and landing (VTOL) SpaceX Starship, the vertical takeoff and horizontal landing (VTOL) US Space Shuttle, and the horizontal takeoff and landing (HTL) German Sanger. Horizontal takeoff and landing (HTL) requires more flexible landing and takeoff sites, utilizing existing airport runways and other facilities, reducing reliance on dedicated launch and recovery sites. This improves accessibility and affordability, facilitating launch and recovery operations from more locations worldwide, and facilitating the development and expansion of commercial space applications, including potential applications in space tourism and satellite launch services.

[0003] The design of a suborbital vehicle is a complex system encompassing subsystems / modules such as geometry, aerodynamics, weight, propulsion, performance, and control and stability. During the conceptual / preliminary design phase, the vehicle's aerodynamic shape, aerodynamic forces, and missile design are the core of the suborbital vehicle system design. Existing two-stage horizontal takeoff and landing suborbital vehicles, the "Sanger" and "Hotol," utilize an air-breathing engine / combined propulsion system as their first-stage propulsion system. These systems offer advantages during the initial takeoff acceleration phase, such as low thrust, high specific impulse, high efficiency, and a mature technical foundation. However, they also face challenges such as limited flight speed and range, sensitivity to weather conditions, and low reliability and high risk. In comparison, rocket propulsion as the first-stage propulsion system offers advantages such as a wide flight envelope, high autonomy, and high thrust. However, the high thrust required for rapid initial takeoff acceleration necessitates precise control of thrust direction and magnitude. Furthermore, precise flight control and system integration must comprehensively consider multiple factors, such as aerodynamic forces and thrust vectoring, posing significant technical challenges. Therefore, it is necessary to develop an integrated aerodynamic and ballistic design method for the rocket-powered two-stage horizontal take-off and landing suborbital vehicle. By designing a reasonable horizontal take-off control strategy and a vehicle shape that meets the requirements of horizontal take-off and extremely high-speed long-distance transportation, the ballistic parameters of the suborbital vehicle can meet the mission requirements. Summary of the Invention

[0004] In view of this, the present invention provides a method for designing the shape, aerodynamic and ballistic integration of a two-stage horizontal take-off and landing suborbital aircraft, which adopts a two-stage rocket propulsion for horizontal take-off and landing; the full ballistics includes a runway segment, a climb segment I, a climb segment II and a reentry segment; wherein, the runway segment adopts a sub-stage rocket propulsion, and when the aircraft is at an altitude of y d When the takeoff altitude H0 is exceeded, the climb phase I is entered; the climb phase I adopts a sub-stage rocket propulsion and sets multiple sequences of increasing altitudes H i and the corresponding angle of attack control sequence α i Coordinate the relationship between the aircraft's flight speed and altitude increase; when the aircraft's altitude y d Exceeding the climb altitude H climb , the flight speed exceeds V climb After the first stage rocket propulsion fuel is exhausted, it enters the climb phase II; the climb phase II uses the second stage rocket propulsion and sets multiple sequences of increasing speed V j and the corresponding angle of attack control sequence α j Coordinate the relationship between the flight speed and the increase in angle of attack of the aircraft; when the propulsion fuel of the second-stage rocket is exhausted, it enters the unpowered reentry phase;

[0005] Taking the maximum lift-to-drag ratio in the reentry phase as the optimization goal and considering the aerodynamic lift coefficient constraint of the aircraft during horizontal takeoff, the aircraft's shape design parameters are optimized to obtain the optimal shape design parameters.

[0006] Based on the optimal shape design parameters, the maximum terminal speed of climb segment II is optimized, and the control node (H i ,α i ,V j ,α j ) is optimized to obtain the optimal active segment trajectory.

[0007] Preferably, in the climbing section I, 3 to 5 sequences of increasing altitude and corresponding angle of attack control sequences are set; in the climbing section II, 2 to 4 sequences of increasing speed and corresponding angle of attack control sequences are set.

[0008] Preferably, the take-off altitude H0 is 0.1-0.8 km; the climbing altitude H climb 60~70km, flight speed V climb More than 2.0 km.

[0009] The optimal angle of attack control scheme for the taxiing phase adopts the angle of attack-altitude node linear interpolation control scheme.

[0010] The optimal angle of attack control scheme for the reentry phase adopts the angle of attack-velocity node linear interpolation control scheme.

[0011] Preferably, the two-stage rocket propulsion uses liquid hydrogen and liquid oxygen propellants.

[0012] Preferably, the dynamic equation of the center of mass of the aircraft is

[0013]

[0014] Where P is thrust, Q is drag, m is aircraft mass, Y is lift on the aircraft, θ is trajectory inclination, σ is track yaw angle, γ s is the trajectory roll angle (roll angle around the velocity vector), V is the flight speed, and g is the earth's gravitational acceleration.

[0015] Preferably, the kinematic equation of the center of mass of the aircraft during the taxiing phase is:

[0016]

[0017] Where x d ,y d , z d are the coordinates of the center of mass of the aircraft in a certain inertial coordinate system;

[0018] The kinematic equations of the center of mass of the aircraft in climb segments I and II are:

[0019]

[0020] Among them, H V represents the transformation matrix from velocity system to ballistic system; V B V represents the transformation matrix from the velocity coordinate system to the projectile coordinate system; G represents the conversion matrix from the launch system to the velocity system; α and β represent the angle of attack and sideslip angle respectively, θ represents the velocity inclination angle, σ represents the track yaw angle; ω e =[ω ex ω ey ω ez ] T Represents the three-axis components of the earth's rotation angular velocity in the ground coordinate system; a ij ,i,j=1,2,3 represents the acceleration component of the implicate acceleration in the launch coordinate system; b ij ,i,j=1,2,3 represents the acceleration component of the Coriolis acceleration in the launch coordinate system; R0=[R 0x R 0y R 0z ] T It represents the three components of the launch point's geocentric radius in the launch coordinate system; g r and g ωe Represents the acceleration components of the earth's gravitational acceleration g in the radial and tangential directions; P e is the thrust of the aircraft engine, m represents the mass of the aircraft; X = Cx qS M and is the drag and lift under flight conditions, C x and They represent the derivatives of the drag coefficient and lift coefficient with respect to the angle of attack α, q is the dynamic pressure of the aircraft; S M is the characteristic area of ​​the aircraft;

[0021] The kinematic equation of the center of mass of the spacecraft during the reentry phase is:

[0022]

[0023] in, is the derivative of the aircraft velocity V, velocity inclination θ, track yaw angle σ, geocentric longitude λ, and geocentric latitude φ; ω e is the Earth's rotational angular velocity ω e Modulus value; V = [V x V y V z ] T Indicates the components of the three axes of the aircraft in the return coordinate system.

[0024] Preferably, the optimization of the aircraft's shape design parameters is as follows:

[0025]

[0026] Among them, X1 is the appearance design parameter; C L is the lift coefficient, C L0 is the lift coefficient required for the suborbital vehicle to take off horizontally, C D is the drag coefficient; x c represents the pressure center coefficient; case 1 and case 2 represent the typical operating conditions of the horizontal takeoff of a two-stage suborbital spacecraft and the reentry phase of a two-stage suborbital spacecraft, respectively; and Design lower and upper bounds for suborbital vehicle shape variables.

[0027] Preferably, the control node of the aircraft (H i ,α i ,V j ,α j ) is optimized as follows:

[0028]

[0029] Among them, H f represents the flight altitude at the end of the active segment; θ ff is the velocity inclination angle of the aircraft at the end of the entire flight segment; η max and q maxis the maximum overload and maximum dynamic pressure of the aircraft during flight; α i and H i (i=1,2,3,4) represents the angle of attack-altitude control node of the two-stage aircraft in the high angle of attack climb phase, α j and V j (j=5,6,7) represents the angle of attack-speed control node of the second-stage aircraft in the high angle of attack climb phase, and the lower and upper bounds of its design variables are denoted as and

[0030] Beneficial effects:

[0031] The present invention adopts a ballistic splicing method to complete the full ballistic design of the horizontal take-off and landing suborbital aircraft in a two-stage rocket power mode, and adopts an altitude-angle of attack and speed-angle of attack node control schemes in the high-angle-of-attack climb segment I and the high-angle-of-attack climb segment II of the aircraft to realize horizontal accelerated flight of the aircraft active segment, thereby providing speed guarantee for the long-range and extremely high-speed flight of the suborbital aircraft; considering the characteristics of the close coupling between the aerodynamic design and the ballistic scheme of the suborbital aircraft, an integrated optimization design of the aerodynamic shape and the ballistic scheme is adopted, and the integrated optimization scheme of the aerodynamic shape and the ballistic scheme of the aircraft is simplified into an inner-layer optimization of the suborbital aircraft shape parameters and an outer-layer optimization of the full ballistic scheme optimization, thereby realizing a long-range and extremely high-speed transportation mission of 10,000 kilometers per hour under a two-stage rocket power mode, greatly reducing the time of the long-range transportation mission, and providing guarantee for the long-range and extremely high-speed transportation mission.

[0032] The present invention provides a demonstration scheme for the rapid iteration of the overall design of a two-stage horizontal take-off and landing suborbital aircraft, which can shorten the iteration cycle of the overall design stage of the suborbital aircraft and accelerate the development speed of the overall design of domestic suborbital aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] Figure 1 This is a flow chart for optimizing the aerodynamic and ballistic integrated design of a two-stage horizontal take-off and landing suborbital vehicle based on the present invention.

[0034] Figure 2 This is a schematic diagram of the parameterized appearance of the suborbital first-stage and second-stage vehicles of the present invention, wherein: Figure 2 (a) is a schematic diagram of the parameterized shape of the booster sub-level. Figure 2 (b) is a schematic diagram of the parameterized appearance of the second sub-level above.

[0035] Figure 3 This is a state parameter change diagram after the aerodynamic and ballistic integrated design method of the two-stage horizontal take-off and landing suborbital aircraft of the present invention, wherein: Figure 3 (a) is the speed change curve, Figure 3 (b) is the dynamic pressure overload curve, Figure 3 (c) is the velocity angle variation curve, Figure 3(d) is the curve of altitude changing with distance, Figure 3 (e) is the total overload variation curve. DETAILED DESCRIPTION

[0036] The present invention is described in detail below with reference to the accompanying drawings and embodiments.

[0037] The present invention provides a method for integrated design of the aerodynamic and ballistic profile of a two-stage horizontal take-off and landing suborbital aircraft. In view of the close coupling characteristics of the aerodynamic profile, aerodynamic force and trajectory of a suborbital aircraft, a trajectory splicing method is used to complete full-process trajectory modeling of the horizontal take-off and landing suborbital aircraft, and a node control method is introduced to complete the control of the two-stage horizontal take-off active segment, thereby realizing horizontal accelerated flight of the active segment of the aircraft. Then, by optimizing the passive segment lift-to-drag ratio to maximize the flight, the shape design parameters of the aircraft are optimized. Furthermore, by optimizing the active segment terminal velocity to maximize the flight, the control node parameters of the active segment are optimized, thereby completing the integrated optimized design of the aerodynamic profile and trajectory of the aircraft under the suborbital aircraft's long-range, extremely high-speed transport mission of 10,000 kilometers per hour.

[0038] The specific steps include:

[0039] Step 1: Complete parametric modeling based on the aerodynamic shape design method of a two-stage horizontal take-off and landing suborbital aircraft, and select key aerodynamic shape parameters that determine the performance of the aircraft as the aerodynamic shape design parameters to be optimized through demonstration and analysis.

[0040] Step 2: The primary mission profile for a two-stage, horizontal takeoff and landing suborbital vehicle consists of four phases: parallel acceleration during a two-stage roll, a high-angle climb, interstage separation to boost the first stage back to the second stage for continued powered high-angle climb, and unpowered reentry of the second stage. Considering these mission profiles, typical operating conditions for both two-stage and single-stage flight were selected.

[0041] Step 3: Both stages of the suborbital spacecraft are propelled by rocket power. Due to the requirements of technological maturity, cleanliness and environmental protection, and high thrust-to-impulse ratio, liquid hydrogen and liquid oxygen propellants are selected to provide power for the suborbital spacecraft.

[0042] Step 4: Full trajectory assembly of the two-stage horizontal take-off and landing suborbital vehicle:

[0043] Step 4.1: State equation analysis of the runway phase of the suborbital vehicle's two-stage parallel runway acceleration.

[0044] According to the basic theorem of dynamics, the dynamic equation of the center of mass of the aircraft established in the track coordinate system is:

[0045]

[0046] Where P is thrust, Q is drag, m is aircraft mass, Y is lift on the aircraft, θ is trajectory inclination, σ is track yaw angle, γs is the trajectory roll angle (roll angle around the velocity vector), V is the flight speed, and g is the earth's gravitational acceleration. Correspondingly, the kinematic equation describing the change in the center of mass of the aircraft in space is

[0047]

[0048] Where x d ,y d , z d are the coordinates of the center of mass of the aircraft in a certain inertial coordinate system.

[0049] When the height y d When the take-off altitude H0 is exceeded, the aircraft dynamics switches to step 4.2. Preferably, the take-off altitude H0 is 0.1 to 0.8 km.

[0050] Step 4.2: State equation analysis of climb segment I of a two-stage high angle of attack climb of a suborbital vehicle.

[0051] Since the active phase of horizontal launch is long and the range is long, the center of mass dynamic equation established in the velocity coordinate system is as follows, considering the influence of the earth's rotation and the earth's oblateness:

[0052]

[0053] Among them, H V represents the transformation matrix from velocity system to ballistic system; V B V represents the transformation matrix from the velocity coordinate system to the projectile coordinate system; G represents the conversion matrix from the launch system to the velocity system; α and β represent the angle of attack and sideslip angle respectively, θ represents the velocity inclination angle, σ represents the track yaw angle; ω e =[ω ex ω ey ω ez ] T Represents the three-axis components of the earth's rotation angular velocity in the ground coordinate system; a ij ,i,j=1,2,3 represents the acceleration component of the implicate acceleration in the launch coordinate system; b ij ,i,j=1,2,3 represents the acceleration component of the Coriolis acceleration in the launch coordinate system; R0=[R 0x R 0y R 0z ] T It represents the three components of the launch point's geocentric radius in the launch coordinate system; g r and g ωe Represents the acceleration components of the earth's gravitational acceleration g in the radial and tangential directions; P e is the thrust of the aircraft engine, m represents the mass of the aircraft; X = C x qSM and is the drag and lift under flight conditions, C x and They represent the derivatives of the drag coefficient and lift coefficient with respect to the angle of attack α, q is the dynamic pressure of the aircraft; S M is the characteristic area of ​​the aircraft.

[0054] In the climb segment I, set multiple sequences of increasing altitudes H i and the corresponding angle of attack control sequence α i , coordinate the growth of the suborbital vehicle's flight speed and altitude in climb phase I to achieve the expected flight speed and altitude indicators.

[0055] Correspondingly, the kinematic equation describing the change of the center of mass of the aircraft is shown in Equation (2). When the flight altitude y d Exceeding the climb altitude H climb , the speed exceeds V climb When the fuel of the first booster stage of the aircraft is exhausted, the aircraft dynamics switches to step 4.3. Preferably, the climbing height H climb Take 60~70km; flight speed V climb Take 2.0km / s or above.

[0056] Step 4.3: State equation analysis of climb phase II of the suborbital vehicle's interstage separation to boost the first stage back to the second stage and continue the powered high-angle climb.

[0057] After the fuel of the first stage of the two-stage suborbital spacecraft is exhausted, the first stage uses near-field return to achieve reuse and reduce costs. The state equation of the second stage spacecraft is consistent with step 4.2.

[0058] In climb segment II, set multiple sequences of increasing speed V j and the corresponding angle of attack control sequence α j , coordinating the increase in the suborbital vehicle's flight speed and angle of attack during climb phase II to achieve the desired flight speed and angle of attack. When the second-stage fuel is exhausted, the suborbital vehicle enters the unpowered reentry phase, and the dynamics switch to step 4.4.

[0059] Step 4.4: Analyze the equation of state of the unpowered reentry phase of the suborbital vehicle second stage.

[0060] During the reentry phase, the spacecraft is in a constant mass flight phase under the influence of the earth's gravity, aerodynamic forces and aerodynamic torque, and without control forces. Its state equation in the return coordinate system is summarized as follows

[0061]

[0062] in, is the derivative of the aircraft velocity V, velocity inclination θ, track yaw angle σ, geocentric longitude λ, and geocentric latitude φ; ω e is the Earth's rotational angular velocity ω e Modulus value; V = [V x V y V z ] T Indicates the components of the three axes of the aircraft in the return coordinate system.

[0063] The angle of attack control scheme for the reentry phase adopts a linear angle of attack-velocity control scheme. f When the aircraft enters the terminal parking state, the flight mission of the suborbital aircraft ends. f Take 2 to 3 km.

[0064] Step 5: Optimization of the aerodynamic and trajectory integrated design of the two-stage horizontal take-off and landing suborbital vehicle

[0065] Considering the mission requirement of long-distance high-speed transportation of 10,000 kilometers per hour for suborbital vehicles, the present invention takes the maximum terminal velocity of the active segment (i.e., the terminal velocity of the second-stage high-angle-of-attack climb segment II) as the optimization target. Considering the characteristics of the close coupling between the aerodynamic design and the ballistic scheme of the suborbital vehicle, the aerodynamic shape and ballistic integrated optimization design is adopted, and the integrated optimization design scheme is simplified into the inner optimization of the suborbital vehicle shape parameters and the outer optimization of the full ballistic scheme optimization. The optimization process is as follows: Figure 1 shown.

[0066] Step 5.1: Aerodynamic Optimization of the Two-Stage Horizontal Takeoff and Landing Suborbital Vehicle

[0067] Considering the mission requirements of suborbital spacecraft horizontal takeoff and long-distance high-speed transportation at a rate of 10,000 kilometers per hour, the aerodynamic lift coefficient constraints for the two-stage spacecraft during horizontal takeoff and the pressure center coefficient range constraints to ensure spacecraft stability are set. With the goal of maximizing the lift-to-drag ratio in the reentry phase, an optimization mathematical model is established as follows:

[0068]

[0069] Among them, X1 is the aerodynamic shape design parameter to be optimized determined in step 1; C L is the lift coefficient, C L0 is the lift coefficient required for the suborbital vehicle to take off horizontally, C D is the drag coefficient; x c represents the pressure center coefficient; case 1 and case 2 represent the typical operating conditions of the two-stage suborbital vehicle's taxiing phase and reentry phase, respectively; and Design lower and upper bounds for suborbital vehicle shape variables.

[0070] Step 5.2: Aerodynamic-trajectory optimization of a two-stage horizontal takeoff and landing suborbital vehicle

[0071] Considering the mission requirements of suborbital spacecraft horizontal takeoff and long-range high-speed transportation at a rate of 10,000 kilometers per hour, the angle of attack-altitude node control method is adopted in the high-angle-of-attack climb segment I of the two-stage spacecraft to regulate the speed and altitude of the horizontal takeoff suborbital spacecraft to achieve horizontal takeoff. The angle of attack-speed node control method is adopted in the high-angle-of-attack climb segment II of the second-stage spacecraft to regulate the speed of the second-stage spacecraft to provide support for subsequent long-range high-speed transportation. The maximum terminal velocity of the active segment (i.e., the terminal velocity of the second-stage spacecraft's powered flight high-angle-of-attack climb segment II) is used as the objective function, and the trajectory optimization mathematical model is established as follows

[0072]

[0073] Among them, H f represents the flight altitude at the end of the active segment; θ ff is the velocity inclination angle at the end of the aircraft climb phase II; η max and q max is the maximum overload and maximum dynamic pressure of the aircraft during flight; α i and H i represents the angle of attack-altitude control node of the high angle of attack climb segment I of the two-stage aircraft, i=1,2,3,…,α j and V j represents the angle of attack-speed control node of the second-stage vehicle in the high-angle-of-attack climb phase II, j = 5, 6, 7, ...; and are the lower and upper bounds of the design variable X2, respectively.

[0074] The following is a more detailed explanation with a specific example:

[0075] Example 1:

[0076] Step 1: Complete parametric modeling based on the aerodynamic shape design method of the two-stage horizontal take-off and landing suborbital aircraft. Through demonstration and analysis, the key aerodynamic shape parameters that determine the performance of the aircraft are selected as the aerodynamic shape design parameters to be optimized. The design parameter statistics of the booster stage and the upper two stages are shown in Table 1. The corresponding parametric shape diagram is shown in Figure 2 shown.

[0077] Table 1 Statistics of the external parameters of the two-stage horizontal take-off and landing suborbital aircraft

[0078]

[0079] Step 2: The primary mission profile for a two-stage horizontal takeoff and landing suborbital vehicle consists of four phases: parallel acceleration during a two-stage roll, a high-angle climb, interstage separation to boost the first stage back to the second stage for continued powered high-angle climb, and unpowered reentry of the second stage. Considering this mission profile, representative operating conditions for both two-stage and single-stage flight were selected for subsequent simulation and optimization.

[0080] Step 3: Both stages of the suborbital vehicle are propelled by liquid hydrogen and liquid oxygen rockets. The propulsion parameters of the liquid hydrogen and liquid oxygen rocket engines are shown in Table 2.

[0081] Table 2 Liquid hydrogen and liquid oxygen rocket engine power parameters

[0082]

[0083]

[0084] Step 4: trajectory splicing based on the three-degree-of-freedom center-of-mass dynamic equation and kinematic equation;

[0085] Step 4.1: Construct the center-of-mass dynamic equation and kinematic equation of the suborbital vehicle during the runway phase, as shown in Equations (1) and (2);

[0086] When the height y d When the take-off altitude H0=500m is exceeded, the aircraft dynamics switches to step 4.2.

[0087] Step 4.2: Construct the center of mass dynamic equation of climbing segment I as shown in Equation (3);

[0088] In the climb segment I, four sequences of increasing altitudes H1, H2, H3, and H4 and corresponding angle of attack control sequences α1, α2, α3, and α4 are set to coordinate the relationship between the flight speed and altitude increase of the suborbital vehicle to achieve the expected indicators of flight speed and altitude.

[0089] When the flight altitude y d Exceeding the climb altitude H climb = 60km, speed exceeds V climb =2.3km / s, after the fuel of the first booster stage of the spacecraft is exhausted, the spacecraft dynamics switches to step 4.3.

[0090] Step 4.3: In the climb phase II, after the fuel of the suborbital vehicle's booster stage is exhausted, the stage uses near-field return to achieve reuse and reduce costs. The state equation of the second-stage spacecraft is consistent with that in step 4.2.

[0091] During climb phase II, three sequential speed increases (V1, V2, and V3) and corresponding angle-of-attack control sequences (α5, α6, and α7) are set to coordinate the relationship between the suborbital vehicle's flight speed and the increase in angle of attack, achieving the desired flight speed and angle of attack. When the upper second stage fuel is depleted, the suborbital vehicle enters an unpowered, passive phase, and the dynamics switch to step 4.4.

[0092] Step 4.4: Construct the state equation of the reentry phase as shown in Equation (4);

[0093] The reentry phase is a passive phase, and its angle of attack control scheme adopts a linear angle of attack-velocity control scheme. f =2km, the spacecraft can be considered to have entered the terminal shutdown state and the flight mission of the suborbital spacecraft is completed.

[0094] Step 5: Integrated design of the aerodynamic and ballistic profile of a two-stage horizontal take-off and landing suborbital vehicle

[0095] Step 5.1: Aerodynamic Optimization of the Two-Stage Horizontal Takeoff and Landing Suborbital Vehicle

[0096] Considering the mission requirements of the suborbital vehicle's horizontal takeoff and the altitude and speed requirements of the climb phase, the optimization mathematical model is established as follows:

[0097] find X1=[C′ T ,S′1,S′2,S′3,W′1,W′2,W′3,C″ T ,S″,W″]

[0098] max f(X1)=C L / C D ,case 2

[0099]

[0100] Among them, the design variables of the shape parameterization are shown in Table 1; C L represents the lift coefficient, C D is the lift coefficient; x c represents the pressure center coefficient; and The lower and upper bounds of the suborbital vehicle shape design variables are shown in Table 1. The shape parameters after the shape-aerodynamic optimization of the two-stage horizontal take-off and landing suborbital vehicle are shown in Table 3.

[0101] Table 3. Parameter statistics of the two-stage horizontal take-off and landing suborbital vehicle

[0102]

[0103] Step 5.2: Carry out aerodynamic-trajectory optimization of the aircraft based on the shape parameters optimized in step 5.1

[0104] Taking the maximum velocity at the end of climb phase II as the objective function, the trajectory optimization mathematical model is established as follows:

[0105] find X2=[α1,α2,α3,α4,H1,H2,H3,H4,α5,α6,α7,V5,V6,V7]

[0106] max f(X2)=V f

[0107]

[0108] The control node parameters obtained after the above optimization are shown in Table 4.

[0109] Table 4 Design statistics of attack angle-altitude / speed nodes in the active phase of suborbital horizontal takeoff and landing

[0110]

[0111] According to the control node parameters in Table 4, the above dynamic and kinematic state equations are integrated to obtain the trajectory curve profile. The resulting state is as follows: Figure 3 As shown. Figure 3 (a) The speed variation curve shows that the aerodynamic shape design method of a two-stage horizontal take-off and landing suborbital aircraft disclosed in the present invention can achieve a cross-speed range flight of 0-6.4 km / s for about 60 minutes; Figure 3 (d) The curve of altitude versus range shows that the aerodynamic shape design method of a two-stage horizontal take-off and landing suborbital aircraft disclosed in the present invention can achieve long-distance transportation in a large airspace with an altitude of 0-115 km and a range of 11,500 km; Figure 3 (b) Dynamic pressure change curve and Figure 3 (e) The total overload variation curve shows that the maximum dynamic pressure is less than 100 kPa and the overload is less than 3.5 g, which preliminarily meets the structural strength requirements for hypersonic vehicle design; Figure 3 (c) The velocity inclination change curve shows that the terminal velocity inclination angle is within 10°, which meets the horizontal landing requirement.

[0112] The aerodynamic shape design method of a two-stage horizontal take-off and landing suborbital aircraft disclosed in the present invention can preliminarily complete the long-range and high-speed mission of the orbital aircraft in a large airspace and wide speed range, and provide a model for the subsequent domestic horizontal take-off and landing suborbital aircraft shape design.

[0113] In summary, the above are only preferred embodiments of the present invention and are not intended to limit the scope of protection of the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.

Claims

1. A method for integrated design of the aerodynamic and ballistic profile of a two-stage horizontal take-off and landing suborbital vehicle, characterized in that: It adopts two-stage rocket propulsion for horizontal take-off and landing; the full trajectory includes the run-out phase, climb phase I, climb phase II and re-entry phase; among them, the run-out phase adopts a sub-stage rocket propulsion, when the aircraft is at an altitude of y d When the takeoff altitude H0 is exceeded, the climb phase I is entered; the climb phase I adopts a sub-stage rocket propulsion and sets multiple sequences of increasing altitudes H i and the corresponding angle of attack control sequence α i Coordinate the relationship between the aircraft's flight speed and altitude increase; when the aircraft's altitude y d Exceeding the climb altitude H climb , the flight speed exceeds V climb After the first stage rocket propulsion fuel is exhausted, it enters the climb phase II; the climb phase II uses the second stage rocket propulsion and sets multiple sequences of increasing speed V j and the corresponding angle of attack control sequence α j Coordinate the relationship between the flight speed and the increase in angle of attack of the aircraft; when the propulsion fuel of the second-stage rocket is exhausted, it enters the unpowered reentry phase; Taking the maximum lift-to-drag ratio in the reentry phase as the optimization goal and considering the aerodynamic lift coefficient constraint of the aircraft during horizontal takeoff, the aircraft's shape design parameters are optimized to obtain the optimal shape design parameters. Based on the optimal shape design parameters, the maximum terminal speed of climb segment II is optimized, and the control node (H i ,α i ,V j ,α j ) is optimized to obtain the optimal active segment trajectory.

2. The method according to claim 1, wherein In the climbing section I, 3 to 5 sequences of increasing altitude and corresponding angle of attack control sequences are set; in the climbing section II, 2 to 4 sequences of increasing speed and corresponding angle of attack control sequences are set.

3. The method according to claim 1, wherein The take-off height H0 is 0.1-0.8 km; the climbing height H climb 60~70km, flight speed V climb More than 2.0 km.

4. The method according to claim 1, wherein The angle of attack control scheme in the taxiing phase adopts the angle of attack-altitude node linear interpolation control scheme.

5. The method according to claim 1, wherein The angle of attack control scheme of the reentry phase adopts the angle of attack-velocity node linear interpolation control scheme.

6. The method according to claim 1, wherein The two-stage rocket is powered by liquid hydrogen and liquid oxygen propellants.

7. The method according to claim 1, wherein The dynamic equation of the center of mass of the aircraft is: Where P is thrust, Q is drag, m is aircraft mass, Y is lift on the aircraft, θ is trajectory inclination, σ is track yaw angle, γ s is the trajectory roll angle, V is the flight speed, and g is the earth's gravitational acceleration.

8. The method according to claim 7, wherein The kinematic equation of the center of mass of the aircraft during the taxiing phase is: Where x d ,y d , z d are the coordinates of the center of mass of the aircraft in a certain inertial coordinate system; The kinematic equations of the center of mass of the aircraft in climb segments I and II are: Among them, H V represents the transformation matrix from velocity system to ballistic system; V B V represents the transformation matrix from the velocity coordinate system to the projectile coordinate system; G represents the conversion matrix from the launch system to the velocity system; α and β represent the angle of attack and sideslip angle respectively, θ represents the velocity inclination angle, σ represents the track yaw angle; ω e =[ω ex ω ey ω ez ] T Represents the three-axis components of the earth's rotation angular velocity in the ground coordinate system; a ij ,i,j=1,2,3 represents the acceleration component of the implicate acceleration in the launch coordinate system; b ij ,i,j=1,2,3 represents the acceleration component of the Coriolis acceleration in the launch coordinate system; R0=[R 0x R 0y R 0z ] T It represents the three components of the launch point's geocentric radius in the launch coordinate system; g r and g ωe Represents the acceleration components of the earth's gravitational acceleration g in the radial and tangential directions; P e is the thrust of the aircraft engine, m represents the mass of the aircraft; X = C x qS M and is the drag and lift under flight conditions, C x and They represent the derivatives of the drag coefficient and lift coefficient with respect to the angle of attack α, q is the dynamic pressure of the aircraft; S M is the characteristic area of ​​the aircraft; The kinematic equation of the center of mass of the spacecraft during the reentry phase is: in, is the derivative of the aircraft velocity V, velocity inclination θ, track yaw angle σ, geocentric longitude λ, and geocentric latitude φ; ω e is the Earth's rotational angular velocity ω e Modulus value; V = [V x V y V z ] T Indicates the components of the three axes of the aircraft in the return coordinate system.

9. The method according to claim 1, wherein The optimization of the aircraft's shape design parameters is specifically as follows: Among them, X1 is the appearance design parameter; C L is the lift coefficient, C L0 is the lift coefficient required for the suborbital vehicle to take off horizontally, C D is the drag coefficient; x c represents the pressure center coefficient; case 1 and case 2 represent the typical operating conditions of the horizontal takeoff of a two-stage suborbital spacecraft and the reentry phase of a two-stage suborbital spacecraft, respectively; and Design lower and upper bounds for suborbital vehicle shape variables.

10. The method according to claim 1, wherein Aircraft control node (H i ,α i ,V i ,α j ) is optimized as follows: Among them, H f represents the flight altitude at the end of the active segment; θ ff is the velocity inclination angle at the end of the entire flight segment of the aircraft; η max and q max is the maximum overload and maximum dynamic pressure of the aircraft during flight; α i and H i (i=1,2,3,4) represents the angle of attack-altitude control node of the two-stage aircraft in the high angle of attack climb phase, α j and V j (j=4,5,6) represents the angle of attack-speed control node of the second-stage vehicle in the high angle of attack climb phase, and the lower and upper bounds of its design variables are denoted as and

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