CN105667811A - Design method for multi-stage coupling integrated structure of front body and air inflow channel of hypersonic aircraft - Google Patents

Abstract

The invention discloses a design method for the multi-stage coupling integrated configuration of the precursor of a hypersonic vehicle and the inlet, which couples two layout design technologies of the multi-stage compression waverider precursor and the truncated Busemann inlet, and can generate any Precursor inlet integrated waverider configuration for given flow capture curve and inlet capture curve at flight altitude and Mach number. Aiming at the problem that the singular point angle of the Busemann reference flow field cannot be coupled with the multi-stage compression waverider, a transitional compression section is added between the two reference flow fields, and the discrete points of the leading edge curve flow around the reference flow around the zero-attack angle cone respectively. In the reference flow field around the conical flow field and the transitional compression section, the streamline tracking is carried out until the angle of the first Mach wave in the inner conical flow field is smaller than the singularity angle, so as to solve the coupling problem. Combining the good off-design state start-up performance of the multi-stage compression waverider precursor and the advantages of the isentropic compression of the truncated Busemann inlet, it provides a new technical approach for the integrated layout design of the inlet of the hypersonic vehicle precursor.

Description

Design method for multi-stage coupling integrated configuration of hypersonic vehicle precursor and inlet

Technical field:

The invention relates to a design method for a multi-stage coupling integrated configuration of a hypersonic vehicle precursor and an air inlet, and belongs to the field of aviation system design.

Background technique:

For hypersonic aircraft, the precursor acts as a compression surface to decelerate and pressurize the incoming flow, which plays a decisive role in the performance of the inlet. Due to its excellent performance and the compression effect on the incoming flow, the waverider is an ideal precursor solution for the air-breathing hypersonic vehicle. In order to push the waverider into practical engineering, it is necessary to develop the integrated design technology of the waverider-inlet-tail nozzle. The technical obstacle that hinders the further engineering application of the waverider lies in the integration of the waverider and the inlet technical bottleneck.

Both multi-stage compression waveriders and Busemann inlets can decelerate and boost the incoming flow, but the compression principles of the two are different, so each has advantages and disadvantages. Multi-stage compression waveriders (Lv Zhenjun, Wang Xudong, Ji Weidong, Wang Jiangfeng. Design technology and experimental analysis of three-stage compression cone-guided waveriders[J]. Experimental Fluid Mechanics, 2015, 05:38-44.) is through multi-stage The shock wave compresses the incoming flow, the compression process is direct and efficient, and it still has excellent performance when it deviates from the design state, and it is not sensitive to the change of flight conditions. However, because it is compressed by shock waves, each shock wave will cause a certain total pressure loss. The more compression stages, the greater the total pressure loss. The Busemann inlet (RamasubramanianV, StarkeyR, LewisM.AnEulerNumericalStudyofBusemannandQuasi-BusemannHypersonicInlets at On-andOff-DesignSpeeds[J].AIAA2008,2008,66.) is composed of a series of compressed Mach waves and terminal shock waves to compress the incoming flow, except for the terminal shock The entire compression process of the wave is isentropic, and the total pressure remains constant during the compression process. However, one disadvantage of isentropic compression is that the compression process is slow, resulting in a long Busemann inlet, which is not suitable for engineering applications. In addition, the wave system structure of the Busemann inlet is extremely complex, and small changes in flight conditions will cause the Busemann inlet to deviate from the design state, and the start-up performance is poor under low Mach number conditions.

Therefore, a new technology that combines the advantages of multi-stage compression waveriders and Busemann inlets is proposed to obtain an integrated layout of precursor inlets suitable for hypersonic vehicle propulsion systems, which has very high academic significance and engineering practical value. .

Invention content:

The purpose of the present invention is to combine the good non-design state startup performance of the existing multi-stage compression waverider precursor and the characteristics of the isentropic compression of the truncated Busemann inlet to propose a brand-new hypersonic aircraft precursor and inlet multi-stage The design method of the coupling integrated configuration can generate the integrated waverider configuration of the precursor inlet with a given leading edge curve and shock wave curve at any flight height and Mach number, which is an integrated configuration for the inlet of the hypersonic aircraft precursor. Modernized layout design provides a new technical approach.

The present invention adopts the following technical scheme: a design method for the multi-stage coupling integrated configuration of the precursor of the hypersonic vehicle and the inlet, which includes the following steps:

Step 1. Given the inlet capture curve and the flow capture curve, the three-dimensional problem is transformed into a two-dimensional problem within the second-order precision according to the kissing cone method, that is, the streamline tracking is performed on the leading edge points in each kissing plane;

Step 2. Perform streamline tracing on the multi-stage compression waverider precursor, and the obtained streamline is the compression surface of the multi-stage compression waverider precursor;

Step 3. Construct the truncated Busemann inner conical reference flow field and perform streamline tracking. The streamline is traced in the last stage reference flow field of the waverider precursor until the first Mach wave of the Busemann inlet is cut off. The Mach wave The relationship between the compression angle and the iterative end angle of the Busemann inner conical reference flow field is a supplementary angle, and the iterative end angle must be smaller than the singular point angle of the Taylor-Maccoll equation of the flow field, otherwise modify the last-stage compression angle of the integrated configuration and repeat steps 2- 3 Until the integration coupling condition is met, step 3 needs to be iterated repeatedly until the flow field parameters at the iteration end angle of the Busemann inner conical reference flow field are satisfied, and the Mach number and angle of attack of the first Mach wave front flow field of the truncated Busemann inlet are the same, And the direction of the flow field after the shock wave reflected by the lip of the inlet inlet at the iteration start angle is horizontal, and the streamline tracing is continued in the Busemann inner cone reference flow field to obtain the truncated Busemann inlet compression surface;

Step 4. Perform three-dimensional fitting on the streamlines tracked in each kissing plane, and finally generate the integrated coupling configuration of the multi-stage compression waverider precursor and the truncated Busemann inlet.

Further, step 2 specifically includes:

Step 2-1. Construct the first-level reference flow field and perform streamline tracking. The first-level compression reference flow field can be constructed by using the zero-attack angle cone flow reference flow field. The flow field is solved by the Taylor-Maccoll equation, and the incoming flow passes through After compression, the first-stage compression Mach wave is tracked in the reference flow field of the zero-attack angle conical flow until reaching the position of the second-stage compression Mach wave;

Step 2-2. Construct the second-stage reference flow field and perform streamline tracking. When the streamline traces to the position of the second-stage compression Mach wave in the reference flow field of zero-attack angle conical flow, the parameters of the flow field before the second-stage shock wave are At a certain angle of attack, at this time, the axis of the reference flow field of the second-stage cone with an inclination angle needs to be rotated by a corresponding angle around the cone point to make it parallel to the incoming flow direction of the second-stage compression Mach wave. Streamline tracking in the reference flow field around the inclined cone;

Step 2-3. The construction method of the reference flow field above the second level is the same as the second level method, and the last level of compression surface of the precursor is a coupling connection transition section combined with the truncated Busemann inlet. After the airflow is compressed by the shock wave Continue to trace streamlines in its reference flow field until the first Mach wave compression angle of the truncated Busemann inlet. If the multi-stage compression waverider has only two stages, this step is the same as step 2-2.

Further, the geometric design parameters of the integrated configuration of the precursor intake port include: the intake port capture curve and the flow capture curve, taking the intersection point of the flow capture curve on the symmetry plane as the coordinate origin, wherein the flow capture curve includes a straight line segment and a curve segment, The equation form of the straight line segment is: 0≤x≤L _u , y=0, the equation form of the curved segment is: x≥L _u , y=B(xL _u ) ^m ; The segment equation form is: 0≤x≤L _s , y=-H, the curve segment equation form is: x≥L _s , y=-H+A(xL _s ) ⁿ , the intersection point of the flow capture curve and the inlet capture curve The coordinates are (X _coj , Y _coj ), the relative position ratio of the intersection of the flow capture curve and the inlet capture curve in the y direction is s, the height of the entire inlet is H, and |Y _coj |=sH is satisfied.

Furthermore, the design parameters of the reference flow field of the multi-stage compression waverider precursor in each kissing plane are as follows: the incoming flow of the conical shock wave in each stage of the reference flow field is parallel to the symmetry axis of the reference flow field, so the rotation of each stage of the reference flow field Make the axis deflection angle α _axis equal to the incoming flow deflection angle α _str , that is, α _axis = α _str , obtain the post-shock parameters of each stage through the oblique shock wave relation as the initial condition of the reference flow field of each stage, and the reference flow field of each stage The field solution equation is the dimensionless Taylor-Maccoll equation, and the equation form is

$\left\{\begin{array}{c}\frac{{\mathrm{<span\; class="notranslate">\; dV</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span>}}{\mathrm{<span\; class="notranslate">\; d</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}<span\; class="notranslate">\; =</span><span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; V</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span>}<span\; class="notranslate">\; +</span>{<span\; class="notranslate">\; (</span>\mathrm{<span\; class="notranslate">\; a</span>}<span\; class="notranslate">\; /</span>{\mathrm{<span\; class="notranslate">\; a</span>}}_{\mathrm{<span\; class="notranslate">\; \&infin;</span>}}<span\; class="notranslate">\; )</span>}^{\mathrm{<span\; class="notranslate">\; 2</span>}}\frac{{\mathrm{<span\; class="notranslate">\; V</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span>}<span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; V</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span>}\mathrm{<span\; class="notranslate">\; cot</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}{{\mathrm{<span\; class="notranslate">\; V</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span>\mathrm{<span\; class="notranslate">\; 2</span>}}<span\; class="notranslate">\; -</span>{<span\; class="notranslate">\; (</span>\mathrm{<span\; class="notranslate">\; a</span>}<span\; class="notranslate">\; /</span>{\mathrm{<span\; class="notranslate">\; a</span>}}_{\mathrm{<span\; class="notranslate">\; \&infin;</span>}}<span\; class="notranslate">\; )</span>}^{\mathrm{<span\; class="notranslate">\; 2</span>}}}\\ \frac{{\mathrm{<span\; class="notranslate">\; dV</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span>}}{\mathrm{<span\; class="notranslate">\; d</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}<span\; class="notranslate">\; =</span>{\mathrm{<span\; class="notranslate">\; V</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span>}\end{array}\right.$

in $V_{\mathrm{\&theta;}}^{*}=\frac{V_{\mathrm{\&theta;}}}{a_{\mathrm{\&infin;}}},V_r^{*}=\frac{V_r}{a_{\mathrm{\&infin;}}},{\left(\frac{a}{a_{\mathrm{\&infin;}}}\right)}^2=1+\frac{\mathrm{\&gamma;}-1}{2}(m_{\mathrm{\&infin;}}^2-V_r^{*2}-V_{\mathrm{\&theta;}}^{*2}),$ θ is the reference flow field iteration angle (the size is the angle with the flow field axis), V _θ is the circumferential velocity component, V _r is the radial velocity component, the asterisk is the dimensionless quantity, a is the speed of sound, and M is the Mach number , and the subscript ∞ represents the incoming flow parameter. According to the flow function formula, the discrete points of the flow capture curve are traced by the streamline in each level of the reference flow field, and the obtained streamline is the partial compression surface of the multi-stage compression waverider precursor.

Furthermore, the design parameters of the Busemann inner conical reference flow field in each kissing plane are as follows: the truncated Busemann inlet is the last stage of compression in the integrated configuration, so the iteration end angle θ _end of the reference flow field must be the same as that of the integrated configuration. The compression angle β _b of the last stage of the model satisfies the supplementary angle relationship, that is, θ _end = 180°-β _b , and when the Mach number of the incoming flow in the reference flow field is M _b1 , the Taylor-Maccoll equation is at the singular point angle θ _sp = 180°-arcsin The flow field parameters at (1/M _b1 ) have an inflection point. In order to solve the problem of the integrated coupling singularity between the truncated Busemann inlet and the multi-stage compression waverider precursor, the iterative end angle θ _end of the Busemann inner conical reference flow field must be smaller than the singularity Point angle θ _sp , when the inflow angle of the Busemann inner cone reference flow field is α _str , θ _end must be <180°-arcsin(1/M _b1 )-α _str , the solution equation of the Busemann inner cone reference flow field is Dimensional Taylor-Maccoll equation, the equation form is

$\left\{\begin{array}{c}\frac{{\mathrm{<span\; class="notranslate">\; dv</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}}{\mathrm{<span\; class="notranslate">\; d</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}<span\; class="notranslate">\; =</span><span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; v</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}<span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; a</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>\mathrm{<span\; class="notranslate">\; 2</span>}}\frac{{\mathrm{<span\; class="notranslate">\; v</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}<span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; v</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}\mathrm{<span\; class="notranslate">\; cot</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}{{\mathrm{<span\; class="notranslate">\; v</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>\mathrm{<span\; class="notranslate">\; 2</span>}}<span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; a</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>\mathrm{<span\; class="notranslate">\; 2</span>}}}\\ \frac{{\mathrm{<span\; class="notranslate">\; dv</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}}{\mathrm{<span\; class="notranslate">\; d</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}<span\; class="notranslate">\; =</span>{\mathrm{<span\; class="notranslate">\; v</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}\end{array}\right.$

in $v_{\mathrm{\&theta;}}^{**}=v_{\mathrm{\&theta;}}/V_{max},v_r^{**}=v_r/V_{max},a^{**2}={(a/V_{max})}^2,{\left(\frac{a}{V_{max}}\right)}^2=\frac{\mathrm{\&gamma;}-1}{2}(1-v_r^{**2}-v_{\mathrm{\&theta;}}^{**2}),$ θ is the iteration angle of the reference flow field, v _θ is the circumferential velocity component, v _r is the radial velocity component, double asterisks are dimensionless quantities, a is the speed of sound, T ₀ is the total temperature of incoming flow, γ is the specific heat ratio of gas, R is the gas constant.

The invention has the following beneficial effects: on the technical level, it solves the interference of the Taylor-Maccoll flow singularity problem in the coupling process of the Busemann inlet and the multi-stage compression waverider precursor; it is different from the existing design technology of the multi-stage compression waverider precursor Compared with the multi-stage coupling technology method of the present invention, the lift-to-drag ratio of the traditional multi-stage compression waverider and the total pressure recovery coefficient of the inlet of the intake port have been significantly improved; the low-speed start-up performance of the intake port has been improved; the intake port has been improved Superb anti-clogging performance.

Description of drawings:

Fig. 1 is a schematic diagram of the coupling configuration of the multi-stage compression waverider precursor and the truncated Busemann inlet in a single kissing plane.

Figure 2 is a schematic diagram of the flow capture curve and the inlet capture curve.

Fig. 3 is a schematic diagram of the reference flow field at each level of the multi-stage compression waverider precursor.

Fig. 4 is a schematic diagram of the reference flow field of the truncated Busemann inlet.

Fig. 5 is a three-dimensional diagram of the coupled configuration of the multi-stage compression waverider precursor and the truncated Busemann inlet.

Fig. 6 is a side view of the integrated configuration generated by the method of the present invention compared with the traditional multi-stage compression waverider precursor.

Fig. 7 is a front view of the comparison between the integrated configuration generated by the method of the present invention and the traditional multi-stage compression waverider precursor.

Fig. 8 is a bottom view of the integrated configuration generated by the method of the present invention compared with the traditional multi-stage compression waverider precursor.

Fig. 9 is a perspective view comparing the integrated configuration generated by the method of the present invention and the traditional multi-stage compression waverider precursor.

Fig. 10 is a dimensionless density nephogram of the integrated configuration flow field generated by the method of the present invention.

Fig. 11 is a dimensionless density cloud map of the traditional multi-stage compression waverider precursor flow field.

Label names in the figure: 1-multistage compression waverider precursor, 2-truncated Busemann inlet, 3-coupling transition section, 4-flow capture curve (FCC), 5-inlet capture curve (ICC), 6-reference flow field around cone with zero angle of attack, 7-reference flow field within Busemann inner cone, 8-reference flow field around cone with inclination angle, 9-reflex shock wave at inlet lip, 10-truncated Busemann inlet The first Mach wave of the road, 11-two-stage compression Mach wave, 12-one-stage compression Mach wave, 13-kiss cut plane, 14-traditional three-stage compression waverider precursor, 15-two-stage compression waverider precursor and truncation Busemann integrated coupling configuration, 16-inlet isolation section, 17-incoming flow capture section, 18-inlet inlet capture section, 19-isolation section capture section.

detailed description:

The present invention will be further described below in conjunction with the accompanying drawings and embodiments.

An integrated design method for multi-stage coupling of hypersonic vehicle precursor and inlet, combined with the integrated three-dimensional layout of multi-stage compression waverider precursor and truncated Busemann inlet, focusing on multi-stage compression waverider precursor and truncated Busemann inlet integrated coupling technology will be described.

The multi-stage coupling integrated configuration of the precursor of the hypersonic vehicle and the inlet of the present invention is composed of a multi-stage compression waverider precursor 1 and a truncated Busemann inlet 2, wherein the multi-stage compression waverider precursor 1 is the last compression section Connect the transition piece 3 for the coupling. At any flight altitude and Mach number, given the variable parameter flow capture curve (FCC, leading edge curve)4 and inlet capture curve (ICC)5, the three-dimensional multi-stage coupling integrated structure The model generation is transformed into the two-dimensional streamline generation problem in each kissing surface 13 within the second-order precision, and the given flow capture curve 4 is traced in the reference flow field at all levels to obtain the multi-level compression surface and the truncated Busemann Integrated configuration of the compression surface; the reference flow field 6 used for the design of the multi-stage compression waverider in each kissing surface 13 and the reference flow field 6 used for the conical flow around the Busemann inner cone are axisymmetric cones The flow field is described by the unquantified Taylor-Maccoll flow equation; aiming at the problem that the singular point angle of the existing Busemann benchmark flow field 7 cannot be coupled with the existing multi-stage compression waverider precursor 1, the existing design method is improved, A coupling transition section 3 is added between the two reference flow fields, and the discrete points of the flow capture curve (leading edge curve) 4 are respectively in the reference flow field 6 of the zero-attack angle cone flow and the reference flow flow around the cone with an inclination angle in the transition compression section. Streamline tracing is carried out in field 8 until the iteration end angle θ _end of the Busemann inner cone reference flow field 7 is smaller than the singular point angle θ _sp , then the reference flow field is transformed into a truncated Busemann inner cone reference flow field 7 and the streamline tracing continues to The inlet lip reflects the shock wave 9; fitting all the streamlines in the kissing plane into a multi-stage coupling compression surface to obtain an integrated configuration.

The specific description parameters of the multi-stage coupling integrated configuration of the precursor of the hypersonic vehicle and the inlet and its reference flow field are as follows. Due to the symmetry, for the convenience of description, the following half-model configuration is used for description:

2-1. The geometric design parameters of the integrated configuration of the precursor inlet include: inlet capture curve (ICC) 5 and flow capture curve (FCC) 4 . Take the intersection point of the flow capture curve (FCC) 4 on the symmetry plane as the coordinate origin, wherein the flow capture curve (FCC) 4 includes a straight line segment and a curved line segment, and the equation form of the straight line segment is: 0≤x≤L _u , y=0, the curved line segment The equation form is: x≥L _u , y=B(xL _u ) ^m ; the inlet capture curve (ICC) 5 includes a straight line segment and a curved line segment, and the equation form of the straight line segment is: 0≤x≤L _s , y=- H, the equation form of the curve segment is: x≥L _s , y=-H+A(xL _s ) ⁿ . The coordinates of the intersection of the flow capture curve (FCC) 4 and the inlet capture curve (ICC) 5 are (X _coj , Y _coj ), and the intersection points of the flow capture curve (FCC) 4 and the inlet capture curve (ICC) 5 are opposite in the y direction The position ratio is s, the height of the whole inlet is H, satisfying |Y _coj |=sH.

2-2. The design parameters of the reference flow field of the multi-stage compression waverider precursor in each kiss section 13 are as follows. The incoming flow before the conical shock wave of each level of reference flow field is parallel to the symmetry axis of the reference flow field, so the reference flow field of each level is rotated so that the axis deflection angle α _axis is equal to the incoming flow deflection angle α _str , that is, α _axis = α _str . The parameters after each stage of compression shock wave are obtained through the oblique shock wave relation as the initial conditions of each stage of reference flow field. The solution equation for each level of reference flow field is the dimensionless Taylor-Maccoll equation, and the equation form is

$\left\{\begin{array}{c}\frac{{\mathrm{<span\; class="notranslate">\; dV</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span>}}{\mathrm{<span\; class="notranslate">\; d</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}<span\; class="notranslate">\; =</span><span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; V</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span>}<span\; class="notranslate">\; +</span>{<span\; class="notranslate">\; (</span>\mathrm{<span\; class="notranslate">\; a</span>}<span\; class="notranslate">\; /</span>{\mathrm{<span\; class="notranslate">\; a</span>}}_{\mathrm{<span\; class="notranslate">\; \&infin;</span>}}<span\; class="notranslate">\; )</span>}^{\mathrm{<span\; class="notranslate">\; 2</span>}}\frac{{\mathrm{<span\; class="notranslate">\; V</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span>}<span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; V</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span>}\mathrm{<span\; class="notranslate">\; cot</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}{{\mathrm{<span\; class="notranslate">\; V</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span>\mathrm{<span\; class="notranslate">\; 2</span>}}<span\; class="notranslate">\; -</span>{<span\; class="notranslate">\; (</span>\mathrm{<span\; class="notranslate">\; a</span>}<span\; class="notranslate">\; /</span>{\mathrm{<span\; class="notranslate">\; a</span>}}_{\mathrm{<span\; class="notranslate">\; \&infin;</span>}}<span\; class="notranslate">\; )</span>}^{\mathrm{<span\; class="notranslate">\; 2</span>}}}\\ \frac{{\mathrm{<span\; class="notranslate">\; dV</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span>}}{\mathrm{<span\; class="notranslate">\; d</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}<span\; class="notranslate">\; =</span>{\mathrm{<span\; class="notranslate">\; V</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span>}\end{array}\right.$

in $V_{\mathrm{\&theta;}}^{*}=\frac{V_{\mathrm{\&theta;}}}{a_{\mathrm{\&infin;}}},V_r^{*}=\frac{V_r}{a_{\mathrm{\&infin;}}},{\left(\frac{a}{a_{\mathrm{\&infin;}}}\right)}^2=1+\frac{\mathrm{\&gamma;}-1}{2}(m_{\mathrm{\&infin;}}^2-V_r^{*2}-V_{\mathrm{\&theta;}}^{*2}),$ θ is the reference flow field iteration angle (the size is the angle with the flow field axis), V _θ is the circumferential velocity component, V _r is the radial velocity component, the asterisk is the dimensionless quantity, a is the speed of sound, and M is the Mach number , and the subscript ∞ represents the incoming flow parameter. According to the flow function formula, the four discrete points of the flow capture curve (FCC) trace the streamline in each level of the reference flow field, and the obtained streamline is the partial compression surface of the multi-level compression waverider precursor.

2-3. The design parameters of the Busemann inner conical reference flow field 7 in each kissing plane are as follows. The truncated Busemann inlet 2 is the last-stage compression section of the integrated configuration, so its reference flow field iteration end angle θ _end must satisfy the supplementary angle relationship with the last-stage compression angle β _b of the integrated configuration, that is, θ _end =180 ° _-βb . When the Mach number of the reference flow field is M _b1 , the Taylor-Maccoll equation has an inflection point at the singular point angle θ _sp =180°-arcsin(1/M _b1 ). Therefore, in order to solve the problem of integral coupling singularity between the truncated Busemann inlet 2 and the multi-stage compression waverider 1, the iteration end angle θ _end of the Busemann inner conical reference flow field 7 must be smaller than the singularity angle θ _sp , and the Busemann inner cone When the inflow deflection angle of the reference flow field 7 is α _str , θ _end must be <180°-arcsin(1/M _b1 )-α _str . The solution equation of the Busemann inner cone reference flow field 7 is the dimension Taylor-Maccoll equation, and the equation form is

$\left\{\begin{array}{c}\frac{{\mathrm{<span\; class="notranslate">\; dv</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}}{\mathrm{<span\; class="notranslate">\; d</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}<span\; class="notranslate">\; =</span><span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; v</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}<span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; a</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>\mathrm{<span\; class="notranslate">\; 2</span>}}\frac{{\mathrm{<span\; class="notranslate">\; v</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}<span\; class="notranslate">\; +</span>{\mathrm{<span\; class="notranslate">\; v</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}\mathrm{<span\; class="notranslate">\; cot</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}{{\mathrm{<span\; class="notranslate">\; v</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>\mathrm{<span\; class="notranslate">\; 2</span>}}<span\; class="notranslate">\; -</span>{\mathrm{<span\; class="notranslate">\; a</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>\mathrm{<span\; class="notranslate">\; 2</span>}}}\\ \frac{{\mathrm{<span\; class="notranslate">\; dv</span>}}_{\mathrm{<span\; class="notranslate">\; r</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}}{\mathrm{<span\; class="notranslate">\; d</span>}\mathrm{<span\; class="notranslate">\; \&theta;</span>}}<span\; class="notranslate">\; =</span>{\mathrm{<span\; class="notranslate">\; v</span>}}_{\mathrm{<span\; class="notranslate">\; \&theta;</span>}}^{<span\; class="notranslate">\; *</span><span\; class="notranslate">\; *</span>}\end{array}\right.$

in $v_{\mathrm{\&theta;}}^{**}=v_{\mathrm{\&theta;}}/V_{max},v_r^{**}=v_r/V_{max},a^{**2}={(a/V_{max})}^2,{\left(\frac{a}{V_{max}}\right)}^2=\frac{\mathrm{\&gamma;}-1}{2}(1-v_r^{**2}-v_{\mathrm{\&theta;}}^{**2}),$ θ is the reference flow field iteration angle (the size is the angle with the flow field axis), v _θ is the circumferential velocity component, v _r is the radial velocity component, double asterisks are dimensionless quantities, a is the speed of sound, T ₀ is the The total flow temperature, γ is the specific heat ratio of the gas, and R is the gas constant. According to the flow function formula, continue to trace the streamline in the Busemann reference flow field 7, and the obtained streamline is the partial compression surface of the Busemann inlet.

The multi-stage coupling integrated configuration and design method of the hypersonic vehicle precursor and the air inlet of the present invention comprises the following steps:

Step 1. Given the inlet capture curve (ICC) 5 and the flow capture curve (FCC) 4, the three-dimensional problem is transformed into a two-dimensional problem within the second-order precision according to the kissing cone method, that is, in each kissing plane 13, the Streamline tracking at leading edge points;

Step 2. Perform streamline tracing on the part of the multi-stage compression waverider precursor, and the obtained streamline is the compression surface of the multi-stage compression waverider precursor 1 .

Step 2-1. Construct the first-level reference flow field and perform streamline tracking. The first-stage compression reference flow field can be constructed by using the reference flow field 6 of zero-attack angle cone flow, and the flow field is solved by the Taylor-Maccoll equation. After the incoming flow is compressed by the first-stage compression Mach wave Tracking is carried out in the field 6 until reaching the position of the second compression Mach wave 11 .

Step 2-2. Construct the second-level reference flow field and perform streamline tracking. When the streamline traces to the position of the secondary compression Mach wave 11 in the reference flow field 6 of the flow around the cone with zero angle of attack, the parameters of the flow field in front of the secondary shock wave have a certain angle of attack. The axis 8 rotates around the cone point by a corresponding angle to make it parallel to the direction of the incoming flow of the two-stage compression Mach wave 11. After being compressed by the second-stage compression Mach wave 11 , the airflow continues to track streamlines in the reference flow field 8 of the second-stage oblique conical flow.

Step 2-3, the construction method of the reference flow field above the second level is the same as the second level method, and the last level of compression surface of the precursor is the coupling connection transition section 3 combined with the truncated Busemann inlet 2, and the airflow passes through the shock wave After compression, continue to trace the streamlines in its reference flow field until the compression angle of the first Mach wave 10 of the Busemann inlet is cut off. -2 is the same.

Step 3. Construct the truncated Busemann inner cone reference flow field 7 and perform streamline tracing. The streamline is traced in the last stage of the reference flow field of the waverider precursor until the first Mach wave 10 of the Busemann inlet is cut off. relationship, and the iteration termination angle must be smaller than the singularity angle of the Taylor-Maccoll equation of the flow field, otherwise modify the last-stage compression angle of the integrated configuration and repeat steps 2-3 until the integrated coupling condition is met. Step 3 needs to iterate repeatedly until the Busemann inner conical reference flow field is satisfied. 7 The flow field parameters at the end angle of the iteration are the same as the flow field Mach number and the angle of attack before the first Mach wave of the truncated Busemann inlet. The direction of the flow field after the shock wave 9 is reflected by the lip of the inlet inlet is horizontal. Streamline tracing is continued in the Busemann inner cone reference flow field 7 to obtain the compression surface of the truncated Busemann inlet channel 2 .

Step 4: Perform three-dimensional fitting on the streamlines tracked in each kissing plane 13, and finally generate an integrated coupling configuration of the multi-stage compression waverider precursor 1 and the truncated Busemann inlet 2.

Applications

In order to verify the effectiveness of the design method, the integrated coupling configuration15 of the two-stage compression waverider precursor and the truncated Busemann is taken as an example.

1. Technical parameters

The selected technical parameters are shown in the table below

Table 1. Design parameters of two-stage compression waverider precursor and truncated Busemann integrated coupling configuration 15

Using the design method of the present invention, the integrated configuration 15 of the two-stage compression waverider and the truncated Busemann inlet is finally generated, and the traditional three-stage compression waverider 14 is generated under the same design parameters.

3. Numerical simulation results

Compared with the traditional three-stage compression waverider precursor 14, the integrated coupling configuration 15 of the two-stage compression waverider and the truncated Busemann has greatly improved the lift-to-drag ratio. In the inviscid state, the maximum increase of the lift-drag ratio at Mach numbers 5, 6, and 7 is 74.29%, 75.96%, and 77.47%, respectively, and they all appear at the design angle of attack of 0°. In viscous state, the maximum increase of lift-to-drag ratio is 48.79%, 52.71%, and 55.34% at Mach numbers 5, 6, and 7, and they all appear at the design angle of attack of 0°. The integrated design method for the multi-stage coupling of the hypersonic vehicle precursor and the air inlet proposed by the present invention greatly improves the lift-drag characteristics of the hypersonic waverider precursor.

In the inviscid design state, the flow coefficient of the two-stage compression waverider and the truncated Busemann integrated coupling configuration 15 is basically equal to that of the traditional three-stage compression waverider 14, and the two are close to 100%. In the viscous state, the two-stage compression The flow coefficient of waverider precursor and truncated Busemann integrated coupling configuration 15 is slightly larger, indicating that compared with the traditional three-stage compression waverider precursor 14, there is less overflow. At the same time, regardless of whether it is viscous or non-viscous, the total pressure recovery coefficient of the inlet capture section 18 of the inlet inlet capture section 18 of the two-stage compression waverider integrated coupling configuration 15 and the truncated Busemann has increased by more than 11%. Although the two configurations have approximately equal cross-sectional areas of the inlet isolation section 16, the integrated coupling configuration 15 of the two-stage compression waverider and cut-off Busemann has a more uniform flow field and a smaller flow field at the inlet inlet. Low pressure area, so there is a small total pressure loss in the inlet isolation section 16. The two-stage compression waverider precursor and truncated Busemann integrated coupling configuration 15 has a nearly 34.7% increase in the total pressure recovery coefficient in the non-viscous state at the capture section 19 of the isolation section 3m away from the entrance of the inlet, and has a 34.7% increase in the viscous state. A lift of almost 20.2%. It shows that the integrated design method of the hypersonic vehicle precursor and the inlet multi-stage coupling proposed by the present invention can significantly improve the compression performance of the hypersonic waverider precursor.

The above is only a preferred embodiment of the present invention, it should be pointed out that for those of ordinary skill in the art, some improvements can also be made without departing from the principle of the present invention, and these improvements should also be regarded as the invention. protected range.

Claims (5)

1. the method for designing of a hypersonic aircraft precursor and the multistage coupling integrated configuration of air intake duct, it is characterised in that: comprise the steps

Step 1, given air intake duct catches curve (5) and curve (4) is caught in flowing, bore method according to osculating and be converted in second order accuracy by three-dimensional problem two-dimensional problems, namely in each osculating face (13), leading edge point is carried out streamlined impeller;

Step 2, multi-stage compression waverider forebody derived part being carried out streamlined impeller, gained streamline is multi-stage compression waverider forebody derived (1) compressing surface;

Step 3, structure blocks Busemann inner conical benchmark flow field (7) and carries out streamlined impeller, streamline is tracked until blocking Busemann air intake duct first Mach wave (10) place in waverider forebody derived afterbody benchmark flow field, this Mach wave compression angle and Busemann inner conical benchmark flow field (7) iteration ends angle are supplementary angle relation, and iteration ends angle is necessarily less than Taylor-Maccoll equation singular point angle, flow field, otherwise revise integration configuration afterbody compression angle and repeat step 2-3 until meeting integration coupling condition, step 3 need iterate until meet Busemann inner conical benchmark flow field (7) iteration ends angle place flow field parameter with block Busemann air intake duct first Mach wave (10) front flow field Mach number with the angle of attack identical, and inlet induction road, iteration initial angle place lip reflected shock wave (9) flow field direction level afterwards, this Busemann inner conical benchmark flow field (7) proceeds streamlined impeller acquisition and blocks Busemann air intake duct (2) compressing surface,

Step 4, carry out three-dimensional matching following the trail of the streamline obtained in each osculating face (13), the coupling arrangement integrated with blocking Busemann air intake duct that ultimately generate multi-stage compression waverider forebody derived.

2. the method for designing of hypersonic aircraft precursor as claimed in claim 1 and the multistage coupling integrated configuration of air intake duct, it is characterised in that: step 2 specifically includes

Step 2-1, structure first order benchmark flow field also carry out streamlined impeller, one stage of compression benchmark flow field can be streamed benchmark flow field (6) with zero-incidence circular cone and construct, Taylor-Maccoll equation solution is passed through in flow field, and incoming flow streams at zero-incidence circular cone after one stage of compression Mach wave (12) compresses and is tracked until arriving two-stage compression Mach wave (11) position in benchmark flow field (6);

Step 2-2, structure benchmark flow field, the second level also carry out streamlined impeller, streamline is when zero-incidence circular cone streams and tracks two-stage compression Mach wave (11) position in benchmark flow field (6), two grades of shock wave front flow field parameter have certain angle of attack, now needing that around cone point, two grades of circular cone benchmark flow field (8) axis with angle are rotated the respective angles flow path direction that makes to come with two-stage compression Mach wave (11) parallel, air-flow continues to stream at two grades of circular cones with angle to carry out streamlined impeller in benchmark flow field (8) after two-stage compression Mach wave (11) compresses;

Step 2-3, benchmark flow field building method more than more than two grades are identical with two stage approach, and the afterbody compressing surface of precursor is and blocks the changeover portion that is of coupled connections (3) that Busemann air intake duct (2) combines, air-flow continues to carry out streamlined impeller in its benchmark flow field after shock wave compression, until blocking position, Busemann air intake duct first Mach wave (10) compression angle place, if multi-stage compression waverider forebody derived (1) only has two-stage, then this step is identical with step 2-2.

3. the method for designing of hypersonic aircraft precursor as claimed in claim 2 and the multistage coupling integrated configuration of air intake duct, it is characterised in that:

Forebody and inlet integration configuration geometry design parameter includes: air intake duct catches curve (5) and curve (4) is caught in flowing, curve (4) is caught at plane of symmetry intersection point for zero with flowing, wherein flowing is caught curve (4) and is included straightway and curved section, and straightway equation form is: 0≤x≤L_u, y=0, curved section equation form is: x >=L_u, y=B (x-L_u)^m; Air intake duct is caught curve (5) and is included straightway and curved section, and straightway equation form is: 0≤x≤L_s, y=-H, curved section equation form is: x >=L_s, y=-H+A (x-L_s)ⁿ, curve (4) and air intake duct are caught in flowing, and to catch curve (5) intersecting point coordinate be (X_coj,Y_coj), flowing catch curve (4) and air intake duct to catch curve (5) intersection point at y direction relative position ratio be s, whole air intake duct height is H, satisfied | Y_coj|=sH.

4. the method for designing of hypersonic aircraft precursor as claimed in claim 3 and the multistage coupling integrated configuration of air intake duct, it is characterised in that:

Each osculating face (13) interior multi-stage compression waverider forebody derived benchmark flow field design parameter is as follows: before every grade of benchmark flow field cone shock, incoming flow is parallel relative to benchmark flow field axis of symmetry, therefore rotates every grade of benchmark flow field and makes axis drift angle α_axisWith incoming flow drift angle α_strEqual, i.e. α_axis=α_str, after trying to achieve every grade of compression shock by oblique shock wave relational expression, parameter is as every grade of benchmark flow field initial condition, and every grade of benchmark flow field solving equation is dimensionless Taylor-Maccoll equation, and equation form is

$\left\{\begin{array}{c}\frac{{\mathrm{dV}}_{\mathrm{\&theta;}}^{*}}{d\mathrm{\&theta;}}=-V_r^{*}+{(a/a_{\mathrm{\&infin;}})}^2\frac{V_r^{*}+V_{\mathrm{\&theta;}}^{*}\cot \mathrm{\&theta;}}{V_{\mathrm{\&theta;}}^{*2}-{(a/a_{\mathrm{\&infin;}})}^2}\\ \frac{{\mathrm{dV}}_r^{*}}{d\mathrm{\&theta;}}=V_{\mathrm{\&theta;}}^{*}\end{array}\right.$

Wherein $V_{\mathrm{\&theta;}}^{*}=\frac{V_{\mathrm{\&theta;}}}{a_{\mathrm{\&infin;}}},V_r^{*}=\frac{V_r}{a_{\mathrm{\&infin;}}},{\left(\frac{a}{a_{\mathrm{\&infin;}}}\right)}^2=1+\frac{\mathrm{\&gamma;}-1}{2}(M_{\mathrm{\&infin;}}^2-V_r^{*2}-V_{\mathrm{\&theta;}}^{*2}),$ θ is iteration angle, benchmark flow field (being sized to and the angle of flow field axis), V_θFor circumferential speed component, V_rFor radial velocity component, asterisk is characteristic, a is the velocity of sound, M is Mach number, footmark ∞ is expressed as incoming flow parameter, according to stream function formula, flowing is caught curve (4) discrete point and is carried out streamlined impeller in every grade of benchmark flow field, and gained streamline is multi-stage compression waverider forebody derived Partial shrinkage face.

5. the method for designing of hypersonic aircraft precursor as claimed in claim 4 and the multistage coupling integrated configuration of air intake duct, it is characterised in that:

In each osculating face, Busemann inner conical benchmark flow field (7) design parameter is as follows: block Busemann air intake duct (2) integrated configuration afterbody compression section, therefore its iteration ends angle, benchmark flow field θ_endMust with integrated configuration afterbody compression angle β_bMeet supplementary angle relation, i.e. θ_end=180 ° of-β_b, benchmark flow field free stream Mach number is M_b1Time, Taylor-Maccoll equation is at singular point angle θ_sp=180 ° of-arcsin (1/M_b1) place's flow field parameter has flex point, in order to solve to block, Busemann air intake duct (2) is integrated with multi-stage compression waverider forebody derived (1) couples singular point problem, Busemann inner conical benchmark flow field (7) iteration ends angle θ_endIt is necessarily less than singular point angle θ_sp, Busemann inner conical benchmark flow field (7) incoming flow drift angle is α_strTime, it is necessary to make θ_end180 ° of-arcsin (1/M of <_b1)-α_str, Busemann inner conical benchmark flow field (7) solving equation is dimension Taylor-Maccoll equation, and equation form is

$\left\{\begin{array}{c}\frac{{\mathrm{dv}}_{\mathrm{\&theta;}}^{**}}{d\mathrm{\&theta;}}=-v_r^{**}+a^{**2}\frac{v_r^{**}+v_{\mathrm{\&theta;}}^{**}\cot \mathrm{\&theta;}}{v_{\mathrm{\&theta;}}^{**2}-a^{**2}}\\ \frac{{\mathrm{dv}}_r^{**}}{d\mathrm{\&theta;}}=v_{\mathrm{\&theta;}}^{**}\end{array}\right.$

Wherein $v_{\mathrm{\&theta;}}^{**}=v_{\mathrm{\&theta;}}/V_{max},v_r^{**}=v_r/V_{max},a^{**2}={(a/V_{max})}^2,{\left(\frac{a}{V_{max}}\right)}^2=\frac{\mathrm{\&gamma;}-1}{2}(1-v_r^{**2}-v_{\mathrm{\&theta;}}^{**2}),$ θ is iteration angle, benchmark flow field, v_θFor circumferential speed component, v_rFor radial velocity component, double asterisk is characteristic, and a is the velocity of sound, T₀For incoming flow stagnation temperature, γ is specific heats of gases ratios, and R is gas constant.