CN114291292A - Aerospace vehicle parallel separation design method - Google Patents
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Abstract
The invention provides a parallel separation design method of aerospace vehicles, which comprises the steps of firstly designing the determined separation Mach number and height according to the overall scheme, carrying out CFD numerical calculation of a first-stage aircraft, and selecting an attack angle corresponding to the lift force of the first-stage aircraft as a separation attack angle; secondly, designing the relative positions of the primary aircraft and the secondary aircraft to minimize the influence of the gravity of the secondary aircraft on the moment characteristic of the primary aircraft and the pneumatic interference in the separation process; then according to the shock wave systems of the primary aircraft and the secondary aircraft within the range of the separation attack angle and the Mach number, the design of the head shape of the secondary aircraft matched with the shock wave systems is developed; and finally, determining the preset rudder deflection angles of the elevators of the primary aircraft and the secondary aircraft by adopting a numerical calculation method, so that the primary aircraft and the secondary aircraft are safely separated, and the attitude change range is smaller. The aerospace vehicle designed by the method can be completely separated in an unconstrained and uncontrolled manner by aerodynamic force.
Description
Technical Field
The invention relates to the technical field of aerospace vehicles, in particular to a design method for realizing unconstrained and uncontrolled safe parallel separation of aerospace vehicles by completely depending on aerodynamic force.
Background
The aerospace craft usually adopts a two-stage orbit entering mode of a backpack type structure, the first-stage aerospace craft adopts air suction type combined power, the second-stage aerospace craft adopts rocket power, and the aerospace craft can be horizontally lifted and landed at a ground airport and can be repeatedly used for multiple times. After the aerospace craft takes off horizontally in an assembly state, the aerospace craft climbs to a separation window in an accelerating mode by means of combined power, and the first-stage aerospace craft and the second-stage aerospace craft are separated in a parallel mode. The first-stage aircraft automatically returns and horizontally lands, and the second-stage aircraft starts the power of the rocket to climb into the target track in an accelerating way.
The parallel separation of the two-stage orbit-entering aircrafts has the characteristics of high Mach number, high dynamic pressure, high dynamic, strong interference and the like, and the main influence on the parallel separation design comprises the following steps: firstly, the high flying pressure makes aerodynamic force play a leading role in the separation process, and the aerodynamic shape plays a decisive role in the feasibility of the separation scheme; secondly, the parallel two-stage aircrafts have equivalent dimensions, a very complex dynamic shock wave interference is generated by a first-stage appearance and a second-stage appearance in the separation process, the action time of strong interference is in the order of hundred milliseconds, and the attitude is difficult to control in real time in the separation process. Therefore, for the parallel separation problem of the two-stage in-orbit aerospace vehicles, the optimal separation scheme is to realize the safe and uncontrolled separation of the two-stage aerospace vehicles by completely relying on aerodynamic force. Similar research reports are not seen in the prior art.
Disclosure of Invention
The invention aims to provide a parallel separation design method for aerospace vehicles, which solves the problem of safe parallel separation of two-stage in-orbit aerospace vehicles and realizes the unconstrained and uncontrolled safe separation completely depending on aerodynamic force by establishing a secondary aircraft head shape design method matched with a shock wave system.
In order to solve the technical problem, the invention provides a parallel separation design method of an aerospace vehicle, which comprises the following steps
According to the separation Mach number and the separation height determined by the overall scheme design, carrying out CFD numerical calculation of the first-stage aircraft, and selecting an attack angle corresponding to the negative lift force of the first-stage aircraft as a separation attack angle;
designing the relative positions of the primary aircraft and the secondary aircraft to minimize the influence of the gravity of the secondary aircraft on the moment characteristic of the primary aircraft and the pneumatic interference in the separation process;
according to the shock wave systems of the primary aircraft and the secondary aircraft within the range of the separation attack angle and the Mach number, the design of the head shape of the secondary aircraft matched with the shock wave systems is developed;
the numerical calculation method is adopted to determine the preset rudder deflection angles of the elevators of the primary aircraft and the secondary aircraft, so that the primary aircraft and the secondary aircraft are safely separated, and the attitude change range is small.
Further, the separation attack angle is-4 to 0 degrees.
Further, the relative position design of the primary aircraft and the secondary aircraft is specifically as follows: the deviation range of the mass center positions of the first-stage aircraft and the second-stage aircraft in the vertical direction is not more than 5% of the total length of the first-stage aircraft, and the installation space of the unlocking mechanism is reserved in the vertical direction for the first-stage aircraft and the second-stage aircraft.
Further, the barycenter position of first order aircraft and second grade aircraft is at vertical direction coincidence, first order aircraft and second grade aircraft are at vertical direction distance 50 ~ 200 mm.
Further, the secondary aircraft head shape design comprises the following steps
Determining the height of the head vertex according to the fact that the back shock wave of the first-stage aircraft is always located below the head vertex of the second-stage aircraft within the separation attack angle and Mach number range;
re-evaluating the thermal environment of the head of the secondary aircraft along the flight profile, and determining the head rounding radius of the secondary aircraft by combining the thermal protection requirement;
the secondary aircraft head bus is designed by adopting an exponential type curve, a von karman type curve or a parabolic type curve.
Further, the head rounding radius determination method is as follows
In the formula, q is the stagnation heat flow,RNis the radius of the stagnation point.
Further, the head bus of the secondary aircraft adopts an exponential curve design
Wherein x is the distance in the length direction; l is the theoretical length of the curve segment; rdIs the maximum radius of the curve segment; n is an index.
Further, RdThe radius of each bus corresponding to the equivalent revolution body, and the value range of the index n is 0.6 to 0.75.
Further, the specific method for determining the preset rudder deflection angles of the elevators of the primary and secondary aircrafts comprises the following steps: and carrying out parallel separation prediction by adopting a CFD numerical calculation method, and repeatedly correcting the primary elevator deflection and the secondary elevator deflection according to a calculation result, so that firstly, the primary aircraft and the secondary aircraft can be safely separated, secondly, the primary attitude and the secondary attitude are stably changed after the safety separation, the attitude change is not more than 10 degrees, and the secondary aircraft is required to be at a positive attack angle.
Compared with the prior art, the invention has the beneficial effects that:
the invention provides a design method for realizing the safe parallel separation of an uncontrolled and unconstrained aerospace craft completely depending on aerodynamic force, which makes full use of the back shock wave of a primary aerospace craft, selects a proper negative attack angle range to enhance the back shock wave strength and enables the primary aerospace craft to bear negative lift force on the one hand, and provides a key parameter design method for matching the head of a secondary aerospace craft with the primary back shock wave on the other hand, so that the secondary aerospace craft bears positive lift force under the action of the primary back shock wave, and the problem of the driving force of parallel separation is solved.
The influence of the interference of a complex shock wave system on the attitude is reduced by reasonably configuring the relative positions of the first-stage aircraft and the second-stage aircraft, and meanwhile, the elevator adopts a reasonable preset rudder deflection angle, so that the change of the attitude angle of the first-stage aircraft and the second-stage aircraft in the separation process is ensured to meet the design requirement.
The parallel separation wind tunnel test verification of the typical aerospace vehicle is completed, the separation scheme has high reliability and low cost, and has higher practical value.
Drawings
The accompanying drawings, which are included to provide a further understanding of the embodiments of the invention and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the principles of the invention. It is obvious that the drawings in the following description are only some embodiments of the invention, and that for a person skilled in the art, other drawings can be derived from them without inventive effort.
FIG. 1 is a schematic diagram of a two-stage orbital aerospace vehicle and primary shock system according to an embodiment of the invention;
FIG. 2 is a schematic structural diagram of a secondary aircraft provided in accordance with an embodiment of the present invention;
FIG. 3 is a schematic diagram of shock wave interference at a typical location during separation of an aerospace vehicle according to an embodiment of the invention;
fig. 4 is a schematic view of a shock system of the aerospace vehicle at the moment when the primary and secondary stages have no shock wave interference with each other and the separation ends according to the embodiment of the invention.
Wherein the figures include the following reference numerals:
1 is a first-level aircraft; 2 is a secondary aircraft; 3 is an unlocking mechanism; 4 is the elevator of the first-level aircraft; 5 is the centroid position of the primary aircraft; 6 is the back shock of the primary aircraft forebody; 7 is a first-class aircraft forebody compression surface shock wave; 8 is the elevator of the second-level aircraft; 9 is the centroid position of the secondary aircraft; 10 is a reflected shock wave system; 11 is the head vertex height; 12 is a head bus; and 13 is the head shock wave of the secondary aircraft.
Detailed Description
It should be noted that the embodiments and features of the embodiments in the present application may be combined with each other without conflict. The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. The following description of at least one exemplary embodiment is merely illustrative in nature and is in no way intended to limit the invention, its application, or uses. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It is noted that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of example embodiments according to the present application. As used herein, the singular forms "a", "an" and "the" are intended to include the plural forms as well, and it should be understood that when the terms "comprises" and/or "comprising" are used in this specification, they specify the presence of stated features, steps, operations, devices, components, and/or combinations thereof, unless the context clearly indicates otherwise.
Aiming at the problem of safe parallel separation of two-stage in-orbit aerospace vehicles, the invention provides a design method for realizing parallel separation of aerospace vehicles by completely relying on aerodynamic force, which comprises the following steps:
s1, according to the separation Mach number and the height determined by the overall scheme design, carrying out CFD numerical calculation of the first-stage aircraft, and selecting an attack angle of which the lift force of the first-stage aircraft is a negative value as a separation attack angle, wherein the attack angle is not too large and is generally selected to be-4-0 degrees.
And S2, determining the relative positions of the primary aircraft and the secondary aircraft, and ensuring that the influence of the gravity of the secondary aircraft on the moment characteristic of the primary aircraft and the aerodynamic interference of the separation process are minimum.
Length direction: the mass center positions of the first-level aircraft and the second-level aircraft are overlapped as much as possible, and the general deviation range does not exceed 5% of the total length of the first-level aircraft; height direction: the installation space of the unlocking mechanism is reserved, and the installation space is generally 50-200 mm.
And S3, according to the shock wave systems of the primary and secondary aircrafts within the separation attack angle and Mach number range, carrying out the design of the head shape of the secondary aircraft matching the shock wave system.
Firstly, according to the back shock wave position of the first-stage aircraft in the range of separation attack angle and Mach number, the back shock wave position is always positioned below the head vertex of the second-stage aircraft, and the height of the head vertex is determined. The height of the primary aircraft, which is determined according to the shock wave position, is generally higher than the height of the unlocking mechanism, and if the height of the primary aircraft is lower than the height of the unlocking mechanism, the original head shape is maintained, so that a high-pressure area is formed on the lower surface of the head of the secondary aircraft, and further, positive lift force is generated.
Secondly, the thermal environment of the head of the secondary aircraft is reevaluated along the flight profile, and the head rounding radius of the secondary aircraft is determined in combination with the thermal protection requirements.
And finally, designing a head bus of the secondary aircraft by adopting an exponential curve.
The head bus can adopt an exponential type curve, a parabolic type curve or a von Karman type curve, and the exponential type curve is generally adopted because the aerodynamic resistance of the exponential type curve is minimum.
And S4, determining the elevator preset rudder deflection angles of the primary aircraft and the secondary aircraft by adopting a numerical calculation method, so that the primary aircraft and the secondary aircraft are safely separated and the attitude change range is small.
In the separation process, complex shock wave system interference is formed between the primary aircraft and the secondary aircraft, pneumatic interference is strong, the change is severe, and the preset rudder deflection angle is difficult to determine simply. And carrying out parallel separation prediction by adopting a CFD numerical calculation method, and repeatedly correcting the primary elevator deflection and the secondary elevator deflection according to the calculation result until the primary aircraft and the secondary aircraft can be safely separated and the attitude change range is less than 10 degrees, so as to obtain a reasonable elevator preset rudder deflection angle.
Specific embodiments of the present invention will be described in detail below with reference to the accompanying drawings.
The aerospace craft parallel separation design method provided by the invention is suitable for a two-stage aerospace craft with parallel separation, as shown in figure 1. The two-stage orbital flight vehicle comprises a first-stage flight vehicle 1 and a second-stage flight vehicle 2. The primary aircraft 1 and the secondary aircraft 2 are connected through the unlocking mechanism 3, and the relative positions of the primary aircraft 1 and the secondary aircraft 2 are fixed under the constraint of the unlocking mechanism 3 during the horizontal takeoff to the separation window. After the aircraft reaches the separation window, the unlocking mechanism 3 releases the restraint of the first-stage aircraft 1 and the second-stage aircraft 2, and the aircraft starts to move by separation under the action of pneumatic force.
And S1, determining the range of the parallel separation attack angle and the position of the key shock wave. And determining the range of the parallel separation attack angle and the key shock wave structure through CFD numerical calculation. And (3) carrying out CFD numerical calculation of the first-stage aircraft 1 according to the conditions of Mach number and altitude separation to obtain the change rule of aerodynamic force of the first-stage aircraft 1 along with the attack angle and the shock wave system structure under each attack angle, wherein the back shock wave 6 of the first-stage aircraft precursor is a key shock wave for influencing parallel separation. An attack angle with a negative lift force of the first-stage aircraft 1 is selected as a separation attack angle, but the attack angle is not too large, and is generally selected to be-4-0 degrees.
And S2, determining the relative positions of the primary aircraft 1 and the secondary aircraft 2. Mainly in the length direction, the mass center position 5 of the primary aircraft and the mass center position 9 of the secondary aircraft are coincident as much as possible in the vertical direction, and the general deviation range does not exceed 5% of the total length of the primary aircraft. The installation space of the unlocking mechanism 3 is reserved in the height direction, and the height direction is generally 50-200 mm.
And S3, developing the shape design of the head of the secondary aircraft of the matched shock wave system.
The profile of the secondary aircraft 2 is substantially determined by the mission. In order to meet the parallel separation design, the appearance characteristics of the secondary aircraft 2 are not suitable to be modified in a large range, so that the invention only designs the head shape of the secondary aircraft 2 to be matched with a shock wave system on the basis of the existing secondary aircraft, and the key parameters comprise the head vertex height, the radius and the bus form.
First, the height of the head apex of the secondary aircraft profile 11, i.e. the height of the head apex of the secondary aircraft from the lower surface of the secondary aircraft, is determined. The principle of selecting the head vertex height 11 is that in the range of separating the attack angle and the Mach number, the back shock wave of the first-stage aircraft 1 is ensured to always hit below the head vertex of the second-stage aircraft 2, and the lower surface of the second-stage aircraft generates a multiple strong reflection shock wave system 10, so that a high-pressure area is formed on the lower surface of the head of the second-stage aircraft 2, and further positive lift force is generated.
Because the length of the primary aircraft 1 is in the order of tens of meters, the height of the back shock wave 6 of the primary aircraft 1 is generally larger than that of the unlocking mechanism 3, and at the moment, the relative height of the secondary aircraft 2 and the primary aircraft 1 is properly increased, so that the back shock wave of the primary aircraft 1 always hits below the head vertex of the secondary aircraft 2. If the determined height of the head vertex is lower than the height of the unlocking mechanism 3, the head vertex of the secondary aircraft is maintained at the original position, and because the back shock wave 6 in the position is definitely hit on the lower surface of the head of the secondary aircraft 2, the design purpose of matching a shock wave system is met.
Secondly, the head rounding radius of the secondary aircraft 2 is determined according to the thermal protection requirements. Since the thermal environment of the head is necessarily changed when the back shock wave 6 of the primary aircraft strikes the head of the secondary aircraft 2, the thermal environment needs to be reevaluated along the flight profile, the stagnation heat flow can be reduced by increasing the radius according to the Fay-Riddell formula shown below, and the radius of the head is redetermined by combining with the thermal protection material.
Wherein q is the stagnation heat flow, RNIs the radius of the stagnation point.
Finally, a head bus 12 of the secondary aircraft 2 is determined. Common aircraft head bus bars are exponential, von karman, parabolic, and the like. The head of the secondary aircraft 2, in addition to meeting the parallel separation design requirements, also requires a minimum of aerodynamic drag. The exponential curve has the best lift-drag characteristics in high-speed flight, and is generally selected in the following specific form:
wherein x is the distance in the length direction; l is the curve segment theoryThe theoretical length; rdFor the maximum radius of the curve segment, R is due to the fact that the head of the secondary aircraft 2 is not axisymmetricdThe radius of the equivalent rotation body corresponding to each bus; n is an index, typically 0.6 to 0.75.
And S4, determining the preset rudder deflection angles of the primary aircraft elevator 4 and the secondary aircraft elevator 8. The preset rudder deflection angle changes the moment characteristics of the first-stage aircraft 1 and the second-stage aircraft 2 to further influence the attitude change, and meanwhile, the lift force is also brought by the control surface to influence the separation speed. Complex shock system interference is formed between the back shock wave 6 of the primary aircraft and the head shock wave 13 of the secondary aircraft in the separation process. As shown in fig. 3, aerodynamic disturbance is strong and varies drastically, and it is difficult to determine the preset rudder deflection angle. The method utilizes a CFD numerical calculation method to carry out parallel separation prediction, and repeatedly corrects the primary elevator deviation and the secondary elevator deviation according to the calculation result. The preset rudder deflection angle firstly ensures that the first-stage aircraft 1 and the second-stage aircraft 2 can be safely separated, as shown in figure 4, no shock wave system interference exists between the first-stage aircraft and the second-stage aircraft, secondly ensures that the attitude change of the first-stage aircraft and the second-stage aircraft is stable after the safe separation, generally not more than 10 degrees, and the second stage should be a positive attack angle.
Features that are described and/or illustrated above with respect to one embodiment may be used in the same way or in a similar way in one or more other embodiments and/or in combination with or instead of the features of the other embodiments.
It should be emphasized that the term "comprises/comprising" when used herein, is taken to specify the presence of stated features, integers, steps or components but does not preclude the presence or addition of one or more other features, integers, steps, components or groups thereof.
The many features and advantages of these embodiments are apparent from the detailed specification, and thus, it is intended by the appended claims to cover all such features and advantages of these embodiments which fall within the true spirit and scope thereof. Further, since numerous modifications and changes will readily occur to those skilled in the art, it is not desired to limit the embodiments of the invention to the exact construction and operation illustrated and described, and accordingly, all suitable modifications and equivalents may be resorted to, falling within the scope thereof.
The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
The invention has not been described in detail and is in part known to those of skill in the art.
Claims (9)
1. A parallel separation design method for aerospace vehicles is characterized by comprising the following steps
According to the separation Mach number and the separation height determined by the overall scheme design, carrying out CFD numerical calculation of the first-stage aircraft, and selecting an attack angle corresponding to the negative lift force of the first-stage aircraft as a separation attack angle;
designing the relative positions of the primary aircraft and the secondary aircraft to minimize the influence of the gravity of the secondary aircraft on the moment characteristic of the primary aircraft and the pneumatic interference in the separation process;
according to the shock wave systems of the primary aircraft and the secondary aircraft within the range of the separation attack angle and the Mach number, the design of the head shape of the secondary aircraft matched with the shock wave systems is developed;
the numerical calculation method is adopted to determine the preset rudder deflection angles of the elevators of the primary aircraft and the secondary aircraft, so that the primary aircraft and the secondary aircraft are safely separated, and the attitude change range is small.
2. The aerospace vehicle parallel separation design method of claim 1, wherein the separation angle of attack is-4 to 0 degrees.
3. The aerospace vehicle parallel separation design method of claim 1, wherein the relative position design of the primary and secondary aircraft is specifically: the deviation range of the mass center positions of the first-stage aircraft and the second-stage aircraft in the vertical direction is not more than 5% of the total length of the first-stage aircraft, and the installation space of the unlocking mechanism is reserved in the vertical direction for the first-stage aircraft and the second-stage aircraft.
4. The aerospace vehicle parallel separation design method of claim 3, wherein the positions of the centers of mass of the primary and secondary aerial vehicles coincide in the vertical direction, and the primary and secondary aerial vehicles are vertically 50mm apart.
5. The aerospace vehicle parallel separation design method of claim 1, wherein the secondary aircraft head profile design comprises the steps of
Determining the height of the head vertex according to the fact that the back shock wave of the first-stage aircraft is always located below the head vertex of the second-stage aircraft within the separation attack angle and Mach number range;
re-evaluating the thermal environment of the head of the secondary aircraft along the flight profile, and determining the head rounding radius of the secondary aircraft by combining the thermal protection requirement;
the secondary aircraft head bus is designed by adopting an exponential type curve, a von karman type curve or a parabolic type curve.
6. The aerospace vehicle parallel separation design method of claim 5, wherein the head blend radius determination method is as follows
Wherein q is the stagnation heat flow, RNIs the radius of the stagnation point.
7. The aerospace vehicle parallel separation design method of claim 5, wherein the secondary aircraft head bus is designed using exponential curves
Wherein x is the distance in the length direction; l is the theoretical length of the curve segment; rdIs the maximum radius of the curve segment; n is an index.
8. The aerospace vehicle parallel separation design method of claim 7, wherein R isdThe radius of each bus corresponding to the equivalent revolution body, and the value range of the index n is 0.6 to 0.75.
9. The aerospace vehicle parallel separation design method according to claim 1, wherein the specific method for determining the preset rudder deflection angles of the elevators of the primary and secondary aircrafts is as follows: and carrying out parallel separation prediction by adopting a CFD numerical calculation method, and repeatedly correcting the primary elevator deflection and the secondary elevator deflection according to a calculation result, so that firstly, the primary aircraft and the secondary aircraft can be safely separated, secondly, the primary attitude and the secondary attitude are stably changed after the safety separation, the attitude change is not more than 10 degrees, and the secondary aircraft is required to be at a positive attack angle.
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|---|---|---|---|---|
| CN116266238A (en) * | 2022-12-28 | 2023-06-20 | 中国航天科工飞航技术研究院（中国航天海鹰机电技术研究院） | Supersonic near-ground parallel interstage separation method with preset rudder deflection characteristics |
| CN116266238B (en) * | 2022-12-28 | 2024-05-03 | 中国航天科工飞航技术研究院（中国航天海鹰机电技术研究院） | Supersonic near-ground parallel interstage separation method with preset rudder deflection characteristics |
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| CN114291292B (en) | 2023-11-03 |
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