Design method of shock wave cone for reducing heat and drag of plane-symmetric hypersonic aircraft

Abstract

The invention provides a design method of a shock wave cone for heat reduction and resistance reduction of a plane-symmetric hypersonic aircraft, the method is used for the plane-symmetric hypersonic aircraft with the flight Mach number Ma being equal to or more than 3, the shock wave cone is installed on the head of the aircraft, and the method is used for meeting the design requirement for heat reduction and resistance reduction of the aircraft and comprises the following steps that 1, features of the shape of the head of the aircraft are extracted; 2, a supporting rod and a precursor form of the shock wave cone are designed; 3, shock wave cone heat reduction and resistance reduction effect evaluation is carried out based on a CFD tool, if heat reduction and resistance reduction requirements are met, a final form of the heat reduction and resistance reduction shock wave cone is formed, and if not, the step 2 is executed to continue optimization design, and design parameters of the shock wave cone are adjusted. The method has the advantages of simplicity, easiness in use, easiness in engineering realization and remarkable improvement effect, and realizes effective reduction of aerodynamic resistance of the plane-symmetric hypersonic aircraft and effective reduction of peak heat flow load by introducing a geometric device to adjust a local flow structure, reduce shock wave intensity and destroy a heat flow peak generation mechanism.

Description

Technical Field

[0001] The present invention relates to the technical field of plane-symmetric hypersonic aircraft design, and in particular to a shock cone design method for reducing heat and drag of a plane-symmetric hypersonic aircraft. Background Art

[0002] Symmetrical hypersonic aircraft is a common aerodynamic layout design for aircraft, such as HTV-2 and HiFIRE-5 aircraft, which are commonly seen in hypersonic aircraft. Heat reduction and drag reduction are important design methods for aircraft to improve range and thermal protection design capabilities.

[0003] When an aircraft flies at hypersonic speed, a bow-shaped detached shock wave is formed at the head due to the interaction between the aircraft and the incoming flow. The strong mutual friction between the incoming flow and the wall of the aircraft can cause an increase in resistance and heat flux load. The resistance of a hypersonic aircraft mainly comes from two parts, including pressure difference resistance and friction resistance. The areas where the heat flux load of a hypersonic aircraft is prominent include the stagnation area and the interference between components, among which the head stagnation area is the key area of ​​thermal protection. Generally speaking, in order to reduce the resistance and heat flux load of the aircraft, it is mainly achieved through the optimization design of the aerodynamic shape. However, due to the constraints of engineering design such as load constraints and aerodynamic performance constraints, the design space based on the optimization of the head shape is very limited. For this reason, it is urgent to form an effective means of reducing heat and drag for hypersonic aircraft.

[0004] There are some shock cone design methods for missiles with rotational bodies or axisymmetric shapes. However, for plane-symmetric hypersonic vehicles, there is no effective and mature shock cone design method. Summary of the invention

[0005] The object of the present invention is to provide a shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic aircraft, so as to solve the above-mentioned technical problems.

[0006] The present invention provides a shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic aircraft, which is used for a plane-symmetrical hypersonic aircraft with a flying Mach number of more than Ma=3. The shock cone is installed at the head of the aircraft to meet the design requirements of reducing heat and drag of the aircraft. The method comprises the following steps:

[0007] Step 1: Feature extraction of the aircraft head shape;

[0008] Step 2: Design of the support rod and the front body of the shock cone;

[0009] Step 3: Based on the CFD tool, evaluate the heat reduction and drag reduction effect of the shock cone. If it meets the requirements, the final form of the heat reduction and drag reduction shock cone is formed. If it does not meet the requirements, go to step 2 to continue optimizing the design and adjust the design parameters of the shock cone.

[0010] Furthermore, in step 1, the head features of the aircraft are extracted, including the bluntness R and characteristic length L of the head; for the ellipsoidal head, the major axis Lx and the minor axis Ly are extracted, and for the spherical head, the diameter D is extracted, and the characteristic length L is defined as the major axis Lx or the diameter D.

[0011] Furthermore, in step 2, the dimensions of the support rod of the shock cone are designed, including the length l and the diameter d; and the form of the shock cone precursor is designed, including the precursor bluntness r and the geometric form.

[0012] Furthermore, the length l of the support rod is 1.5-2.5 times the characteristic length L of the aircraft head, and the diameter d of the support rod is 0.05-0.2 times the characteristic length L of the aircraft head; the forebody bluntness r of the shock cone is greater than the bluntness R of the aircraft head.

[0013] Furthermore, in step 2, the shape of the shock cone is optimized and designed, and an orthogonal matrix design is carried out for the forebody bluntness r, geometric form, support rod length l and diameter d, and local transition form to obtain the influence of the shock cone under different parameters on the heat reduction and drag reduction effects for optimization design.

[0014] Furthermore, in step 3, based on the CFD numerical calculation tool, the flow field calculation with and without the shock cone is carried out to obtain the influence of the shock cone on the flow field structure and physical quantities of the aircraft head, and the flow control effect of heat reduction and drag reduction is statistically calculated. If the engineering design requirements are met, the final heat reduction and drag reduction shock cone geometry is formed. If not, go to step 2 to further optimize the design parameters of the shock cone.

[0015] Furthermore, the effect of heat reduction and drag reduction in step 3 is determined by the following method:

[0016]

[0017] Where ΔQ max is the reduction of the peak value of the heat flux load on the entire surface of the aircraft, Q max,control is the peak value of the heat flux load on the entire surface of the vehicle with shock cone, Q max,no_contrl is the peak value of the wall heat flux load on the entire surface of the original aircraft;

[0018]

[0019] Where ΔD is the reduction of the overall aerodynamic drag of the aircraft, D control is the aerodynamic drag of the vehicle with shock cone, D no_contrl is the aerodynamic drag of the original aircraft.

[0020] Furthermore, the shock cone is composed of a forebody and a support rod structure, one end of the support rod is connected to the forebody, and the other end is connected to the head of the aircraft.

[0021] Furthermore, the forebody bluntness r, geometric form, length l of the support rod, and diameter d of the support rod of the shock cone are determined based on the flow field regulation and heat reduction and drag reduction effects of the aircraft head area.

[0022] Furthermore, by introducing the shock cone, the flow field structure and the distribution of physical quantities at the head of the aircraft are adjusted, thereby achieving the flow control effect of reducing the heat flux peak in the head area of ​​the aircraft and reducing the pressure difference resistance and friction resistance of the aircraft.

[0023] The present invention is based on flow characteristics and physical mechanisms, and has the advantages of being simple to use, easy to implement in engineering, and having significant improvement effects. By introducing geometric devices to adjust the local flow structure, the shock wave intensity can be reduced, the heat flux peak generation mechanism can be destroyed, and the aerodynamic drag of the plane-symmetric hypersonic aircraft and the peak heat flux load can be effectively reduced. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] In order to more clearly illustrate the specific implementation methods of the present invention or the technical solutions in the prior art, the drawings required for use in the specific implementation methods or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some implementation methods of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0025] Figure 1 A flow chart for running the design method of the present invention;

[0026] Figure 2 It is a schematic diagram of the geometric parameter design of the shock cone of the present invention;

[0027] Figure 3 is a side view of the shock cone of the present invention;

[0028] Figure 4 is a top view of the shock cone of the present invention;

[0029] Figure 5 is a front view of the shock cone of the present invention;

[0030] Figure 6 A three-dimensional diagram of the shock cone of the present invention;

[0031] Figure 7 The schematic diagram of the calculation and investigation of the head and shock cone of the background aircraft of the present invention with a length of 0.12m;

[0032] Figure 8 The computational grid of the background aircraft and shock cone of the present invention;

[0033] Fig. 9This is a schematic diagram of heat flow distribution at the head of the original aircraft of the present invention;

[0034] Fig.10 It is a schematic diagram of heat flow distribution of the aircraft head with shock cone of the present invention;

[0035] Fig.11 It is a schematic diagram of the heat flow profile comparison result of the present invention;

[0036] Description of reference numerals:

[0037] In the figure: 1-aircraft, 2-support rod, 3-front body; DETAILED DESCRIPTION

[0038] The technical solution of the present invention will be clearly and completely described below in conjunction with the embodiments. Obviously, the described embodiments are part of the embodiments of the present invention, rather than 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.

[0039] In the description of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "length", "width", "thickness", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", "clockwise", "counterclockwise" and the like indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, and do not indicate or imply that the referred device or element must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as limiting the present invention.

[0040] In addition, the terms "first" and "second" are only used for descriptive purposes and cannot be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Therefore, the features defined as "first" and "second" may explicitly or implicitly include one or more of the features. In the description of the present invention, "multiple" means two or more, unless otherwise clearly and specifically defined. In addition, the terms "installed", "connected" and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, or it can be indirectly connected through an intermediate medium, or it can be the internal connection of two components. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0041] Example 1

[0042] like Figure 1-Figure 11 As shown:

[0043] A shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic aircraft is provided, which is used for a plane-symmetrical hypersonic aircraft 1 flying at a Mach number of Ma=3 or above. The shock cone is installed at the head of the aircraft 1 to meet the design requirements of reducing heat and drag of the aircraft 1.

[0044] The shock cone is composed of a front body 3 and a support rod 2 structure. One end of the support rod 2 is connected to the front body 3, and the other end is connected to the head of the aircraft 1. It is a passive heat reduction and drag reduction flow control technology.

[0045] A shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic vehicle 1 specifically comprises the following steps:

[0046] Step 1: Feature extraction of the head shape of aircraft 1.

[0047] The head features of the aircraft 1 are extracted, including the bluntness R of the head and the characteristic length L of the head; for the ellipsoidal head, the major axis Lx and the minor axis Ly are extracted, and for the spherical head, the diameter D is extracted. The characteristic length L of the head is defined as the major axis Lx or the diameter D.

[0048] Step 2: Design of the support rod 2 and the front body 3 of the shock cone.

[0049] The dimensions of the support rod 2 of the shock cone are designed, including the length l and the diameter d. The form of the shock cone forebody 3 is designed, including the bluntness r and the geometric form of the forebody 3.

[0050] The length l of the support rod 2 is 1.5-2.5 times the characteristic length L of the aircraft 1 head, and the diameter d of the support rod 2 is 0.05-0.2 times the characteristic length L of the aircraft 1 head; the bluntness r of the forebody 3 of the shock cone is greater than the bluntness R of the aircraft 1 head.

[0051] The shock cone shape is optimized by carrying out orthogonal matrix design for the bluntness r of the forebody 3, the geometric form, the length l and diameter d of the support rod 2, and the local transition form, and obtaining the influence of the shock cone under different parameters on the heat reduction and drag reduction effect for optimization design.

[0052] Step 3: Based on the CFD tool, evaluate the heat reduction and drag reduction effect of the shock cone. If it meets the requirements, the final form of the heat reduction and drag reduction shock cone is formed. If it does not meet the requirements, go to step 2 to continue optimizing the design and adjust the design parameters of the shock cone.

[0053] Based on CFD numerical calculation tools, flow field calculations with and without shock cones are carried out to obtain the influence of shock cones on the flow field structure and physical quantities of the head of the aircraft 1, and the flow control effect of heat reduction and drag reduction is statistically analyzed. If the engineering design requirements are met, the final heat reduction and drag reduction shock cone geometry is formed. If not, go to step 2 to further optimize the design parameters of the shock cone.

[0054] The effect of heat reduction and drag reduction is determined by the following methods:

[0055]

[0056] Where ΔQ max is the peak reduction of the heat flux load on the entire surface of the aircraft 1, Q max,control is the peak value of the heat flux load on the entire surface of the vehicle 1 with the shock cone, Q max,no_contrl is the peak value of the wall heat flux load on the entire surface of the original aircraft 1;

[0057]

[0058] Where ΔD is the reduction of the overall aerodynamic drag of the aircraft 1, D control is the aerodynamic drag of the vehicle 1 with shock cone, D no_contrl is the aerodynamic drag of the original aircraft 1.

[0059] The bluntness r, geometric form, length l and diameter d of the shock cone's forebody 3 are determined based on the flow field regulation and heat and drag reduction effects of the aircraft's 1 head region.

[0060] The geometrical form of the precursor 3 of the shock cone is semi-ellipsoidal or hemispherical.

[0061] By introducing the shock cone, the flow field structure and the distribution of physical quantities at the head of the aircraft 1 are adjusted to reduce the peak value Q of the heat flux in the head area of ​​the aircraft 1. max , reduce the pressure difference resistance D of aircraft 1 p and friction resistance D f flow control effect.

[0062] In the present invention, when the plane-symmetric hypersonic aircraft 1 degenerates into an axisymmetric aircraft, the design parameters are still applicable.

[0063] Fig. 9 The heat flux distribution on the symmetric surface of the head of the original aircraft 1; Fig.10 The heat flux distribution on the symmetric surface of the head of the aircraft 1 with the shock cone.

[0064] like Fig.11 As shown, the comparison results of the heat flow profiles on the symmetric surface of the head of the plane-symmetric hypersonic vehicle 1 before and after the introduction of the shock cone.

[0065] Comparing the curves, we can see that the peak heat flux in the head area of ​​aircraft 1 is significantly reduced. As can be seen from the figure, after control, the peak heat flux in the head of aircraft 1 is reduced by 24.58%, and the maximum peak heat flux load on the entire surface under the corresponding state is reduced by 12.87%. It should be pointed out that the reduction in resistance of aircraft 1 is related to the scale of aircraft 1. The corresponding resistance reduction of aircraft 1 of the current scale is between 4.83% and 6.07%.

[0066] The detailed parameters of the incoming flow conditions of the numerical simulation of the present invention are shown in the following table:

[0067] Ma Re / m Pt(MPa) Tt(K) 8 1.14E7 5 748

[0068] In the above table, the present invention is applied to the incoming flow state of Ma=8, and the actual effect of the shock cone formed by the present invention is verified and evaluated by the CFD numerical simulation method.

[0069] The comparison of the resistance and the peak value of the heat flux under the conditions investigated by the present invention is shown in the following table:

[0070] Ma Re / m T_inf(K) 7.96 1.11E+07 54.7 Case D Dp Df ΔD ΔDp ΔDf Qmax ΔQmax smooth 5.431 4.777 0.653 481.190 JBZ-11.5-d0.1 5.423 4.808 0.615 0.15％ -0.63％ 5.89％ 441.098 8.33％ JBZ-12.0-d0.1 5.381 4.768 0.614 0.92％ 0.21％ 6.07％ 441.013 8.35％ JBZ-12.5-d0.1 5.339 4.725 0.614 1.70％ 1.11％ 5.97％ 419.283 12.87％ JBZ-13.0-d0.1 5.315 4.695 0.619 2.15％ 1.73％ 5.22％ 629.072 -30.73％ JBZ-13.5-d0.1 5.296 4.674 0.622 2.49％ 2.17％ 4.83％ 664.828 -38.16％ JBZ-12.5-d0.2 5.346 4.726 0.619 1.57％ 1.07％ 5.23％ 419.482 12.82％ JBZ-12.5-d0.05 5.344 4.725 0.619 1.60％ 1.10％ 5.31％ 422.136 12.27％

[0071] As shown in the above table, Smooth represents the original aircraft 1, and JBZ-la-db represents that the length of the support rod 2 is a times the characteristic length L of the head of the aircraft 1, and the diameter is b times the characteristic length L of the head of the aircraft 1.

[0072] From the calculation results, it can be seen that when the support rod 2 with a diameter of 0.1 times the characteristic length L of the head of the aircraft 1 is used, when the length l of the support rod 2 changes from 1.5L to 3.5L, the overall aerodynamic resistance reduction range ΔD of the aircraft 1 changes from 0.15% to 2.49%, where the friction resistance D f The reduction is more obvious, with a maximum reduction of 6.07%. When the length of the support rod 2 changes from 2.0L to 3.5L, the pressure difference resistance D p and friction resistance D f All have decreased.

[0073] When the length of the support rod 2 changes, the heat flux peak value starts to increase at a length of l=3.0L and beyond.

[0074] When the length of the fixed support rod 2 is l=2.5L and the diameter of the support rod 2 is d=0.05l-0.2l, the aerodynamic resistance and the heat flux peak reduction effects are basically the same, indicating that the influence of the diameter of the support rod 2 on the control effect is smaller than that of the length of the support rod 2.

[0075] The present invention realizes the effective reduction of aerodynamic drag and peak heat flux load of a plane-symmetrical hypersonic aircraft; the method of the present invention is based on flow characteristics and physical mechanisms, and has the advantages of being simple and easy to use, easy to implement in engineering and having significant improvement effects; by introducing a geometric device to adjust the local flow structure, the shock wave intensity is reduced, the heat flux peak generation mechanism is destroyed, and the design goal of reducing heat and drag of a plane-symmetrical hypersonic aircraft is achieved; by adjusting the bow shock wave at the head to an oblique shock wave, and combining the backflow of the shock cone, the pressure and heat flux peak in the head area of ​​the aircraft are effectively reduced, thereby achieving the control goal of reducing heat and drag.

[0076] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or replace some or all of the technical features therein with equivalents. However, these modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. A shock cone design method for reducing heat and drag of a symmetrical hypersonic vehicle, characterized in that: For a plane-symmetrical hypersonic aircraft with a flying Mach number of more than Ma=3, the shock cone is installed at the head of the aircraft to meet the heat reduction and drag reduction design requirements of the aircraft, and the method includes the following steps: Step 1: Feature extraction of the aircraft head shape; Step 2: Design of the support rod and the front body of the shock cone; Step 3: Based on the CFD tool, evaluate the heat reduction and drag reduction effect of the shock cone. If it meets the requirements, the final form of the heat reduction and drag reduction shock cone is formed. If it does not meet the requirements, go to step 2 to continue optimizing the design and adjust the design parameters of the shock cone.

2. The shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic vehicle according to claim 1, characterized in that: In step 1, the head features of the aircraft are extracted, including the bluntness R and characteristic length L of the head; for the ellipsoidal head, the major axis Lx and the minor axis Ly are extracted, and for the spherical head, the diameter D is extracted. The characteristic length L is defined as the major axis Lx or the diameter D.

3. The shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic vehicle according to claim 2 is characterized in that: In step 2, the supporting rod dimensions of the shock cone are designed, including the length l and the diameter d; the forebody form of the shock cone is designed, including the forebody bluntness r and the geometric form.

4. The shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic vehicle according to claim 3 is characterized in that: The length l of the support rod is 1.5-2.5 times the characteristic length L of the aircraft head, and the diameter d of the support rod is 0.05-0.2 times the characteristic length L of the aircraft head; the forebody bluntness r of the shock cone is greater than the bluntness R of the aircraft head.

5. The shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic vehicle according to claim 3 is characterized in that: In step 2, the shock cone’s shape is optimized and an orthogonal matrix design is performed on the forebody bluntness r, geometric form, support rod length l and diameter d, and local transition form to obtain the influence of shock cone under different parameters on heat reduction and drag reduction effects for optimization design.

6. The shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic vehicle according to claim 1, characterized in that: In step 3, based on the CFD numerical calculation tool, the flow field calculation with and without shock cone is carried out to obtain the influence of shock cone on the flow field structure and physical quantity of the aircraft head, and the flow control effect of heat reduction and drag reduction is statistically analyzed. If the engineering design requirements are met, the final heat reduction and drag reduction shock cone geometry is formed. If not, go to step 2 to further optimize the design parameters of the shock cone.

7. The shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic vehicle according to claim 1, characterized in that: The effect of heat reduction and drag reduction in step 3 is determined by the following method: Where ΔQ max is the reduction of the peak value of the heat flux load on the entire surface of the aircraft, Q max,control is the peak value of the heat flux load on the entire surface of the vehicle with shock cone, Q max,no_contrl is the peak value of the wall heat flux load on the entire surface of the original aircraft; Where ΔD is the reduction of the overall aerodynamic drag of the aircraft, D control is the aerodynamic drag of the vehicle with shock cone, D no_contrl is the aerodynamic drag of the original aircraft.

8. The shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic vehicle according to claim 1, characterized in that: The shock cone is composed of a forebody and a support rod structure, one end of the support rod is connected to the forebody, and the other end is connected to the head of the aircraft.

9. The shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic vehicle according to claim 1, characterized in that: The forebody bluntness r, geometric form, support rod length l, and support rod diameter d of the shock cone are determined based on the flow field regulation and heat reduction and drag reduction effects in the head area of ​​the aircraft.

10. The shock cone design method for reducing heat and drag of a plane-symmetrical hypersonic vehicle according to claim 5, characterized in that: By introducing the shock cone, the flow field structure and the distribution of physical quantities at the head of the aircraft are adjusted, so as to achieve the flow control effect of reducing the peak heat flux in the head area of ​​the aircraft and reducing the pressure difference resistance and friction resistance of the aircraft.

Images