The method adopts a back-mounted engine wing-body blended layout

Abstract

The design method of the fly-by-wire integrated design of the back-mounted engine wing-body blended configuration is presented, including the design of the upper surface of the rear fuselage based on the engine position and installation angle, the design of the short inlet and the engine inlet, and the design of the tail nozzle and the fuselage. The difficulty of the coordination design between the circular nozzle and the upper surface of the fuselage is solved by lifting the tail nozzle, which ensures the circular tail nozzle and avoids the ablation of the jet flow to the upper surface of the fuselage. The inlet quality is ensured by the design of the upper surface of the fuselage and the short inlet. The engine is located at the back of the rear fuselage, and the tail nozzle is not exposed. The global parameterized modeling is avoided by dividing the fixed area, the modified area and the transition area, which improves the design efficiency. The three-dimensional integrated design of the engine position, installation angle and other parameters and the fuselage shape control parameters is realized, which effectively overcomes the limitations of the existing design method mainly for the separate design and two-dimensional design of the inlet and short inlet.

Description

Technical Field

[0001] This invention relates to the field of aircraft aerodynamic design, specifically to an integrated flight-engine design method employing a dorsal-mounted engine blended wing-body layout. Background Technology

[0002] Blended wing-body configuration refers to an aircraft shape where the wings and fuselage are highly integrated, forming a full-lift surface. Under the same load requirements, the blended wing-body design can reduce the overall wetted area of ​​the aircraft, thereby reducing frictional drag. Compared with the traditional cylindrical fuselage + wing configuration, the cruise efficiency of the blended wing-body configuration can be 15-20% higher than that of the Boeing B787 and Airbus A350, and it also has the potential to reduce noise, emissions, and structural weight.

[0003] Aircraft-engine integrated design is a crucial component of aircraft design. In blended wing-body configurations, engines are typically located in the upper rear of the fuselage. Depending on whether the engine nacelle is integrated with the fuselage, it can be categorized as either a dorsal-supported or dorsal-mounted type. In a dorsal-supported configuration, the engine nacelle is supported by a bracket on the upper surface of the fuselage. This layout is technically simple and easy to implement, but it is prone to generating shock waves between the nacelle and the fuselage and unfavorable nose-down moment, reducing aerodynamic efficiency. The dorsal-mounted configuration, also known as an embedded or recessed layout, embeds the engine nacelle inside the fuselage. It utilizes boundary layer intake to improve propulsion efficiency and fuel economy. This embedded layout theoretically offers performance benefits, but its technical challenge lies in the integrated design issues arising from flight-engine coupling.

[0004] The existing design method for a dorsal-mounted engine blended wing-body layout has the following shortcomings.

[0005] (i) In most existing blended wing-body configurations with embedded engines, the exhaust nozzles extend beyond the trailing edge of the fuselage (Carter MB, Campbell RL, Pendergraft OC, et al. Designing and Testing a Blended WingBody with Boundary-Layer Ingestion Nacelles[J]. Journal of Aircraft, 2006.). This design approach aims to maintain a circular nozzle to reduce thrust loss. However, it also faces challenges in coordinating the circular nozzle with the near-planar fuselage surface, lacking a design method that ensures the nozzle's shape is compatible with the fuselage surface. This outward-extending nozzle design occupies space that would otherwise be used for the tail rudder, making tail rudder placement difficult and weakening longitudinal trim capability. Furthermore, the fuselage cannot effectively shield engine exhaust noise, increasing takeoff and landing noise levels.

[0006] (ii) Most existing design methods for the dorsal-mounted engine blended wing-body layout involve separate designs of the air intake or nacelle (Fleming J, Anderson J, Ng W, et al. Sensing and Active Flow Control for Advanced BWB Propulsion-Airframe Integration Concepts. 2005.), and no integrated flight-engine design method considering fuselage modification has been found.

[0007] (III) Most existing design methods for hybrid wingbody layouts with dorsal-mounted engines are two-dimensional designs in the vertical plane of the engine (Liou M, Kim H, Liou M. Challenges and Progress in Aerodynamic Design of Hybrid Wingbody Aircraft with Embedded Engines. 2016.), with few three-dimensional design methods.

[0008] To address the limitations of existing dorsal-engine blended wing-body layout design methods, such as difficulties in coordinating the exhaust nozzle with the fuselage upper surface, separate design of the air intake / nacelle, and the constraints of two-dimensional design, a new integrated flight-engine design method for dorsal-engine blended wing-body layouts is proposed. This method is a parametric design approach that comprehensively considers aspects such as shape coordination, aerodynamic constraints, and intake and exhaust quality, achieving a three-dimensional integrated design of the fuselage, air intake, and nacelle. The method solves the problem of maintaining a circular exhaust nozzle by employing an "uplift" strategy with a height difference between the engine exhaust nozzle and the fuselage upper surface, while also ensuring the exhaust nozzle is not exposed, thus guaranteeing noise shielding and providing space for the tail rudder. Through integrated shaping of the fuselage upper surface, the method ensures uniform and unobstructed airflow through the air intake, while maintaining an intake length comparable to that of conventional turbofan engines, thus guaranteeing airflow quality. Furthermore, the strategy of integrating the lower half of the nacelle with the fuselage and rectifying the airflow in the nacelle-fuselage integration area ensures high-quality airflow. Summary of the Invention

[0009] To overcome the difficulties in coordinating the tail nozzle with the shape of the fuselage upper surface, the shortcomings of separate air intake / nacelle design, and the limitations of two-dimensional design in the existing technology, this invention proposes a flight-engine integrated design method for a dorsal engine wing-body blended layout.

[0010] The integrated flight-engine design process of the dorsal engine blended wing-body layout proposed in this invention includes three aspects: design of the upper surface of the rear fuselage based on the engine position and installation angle, integration design of the nacelle / inlet with the fuselage, and coordinated design of the tail nozzle with the fuselage. Taking one side of the fuselage model's symmetrical plane as an example, the integrated flight-engine design method of the dorsal engine blended wing-body layout is described.

[0011] The specific process is as follows:

[0012] Step 1, Design of the upper surface of the rear fuselage based on the engine position and mounting angle:

[0013] For the clean configuration of the blended wing-body layout without an engine, the upper surface of the rear fuselage is designed in conjunction with the turbofan engine power nacelle.

[0014] The clean configuration of the blended wing-body layout mentioned in Step 1 refers to an aerodynamic shape consisting only of the fuselage and wings, excluding components such as engines, tail fins, control surfaces, and landing gear. The turbofan engine nacelle refers to a simplified shape where the engine fan and turbine are reduced to intake and exhaust surfaces. The trailing edge line of the lower fuselage surface in the clean configuration is parallel to the trailing edge line of the upper fuselage surface. The upper fuselage surface trailing edge line is the curve formed by connecting the trailing edge points of the upper surfaces of the fuselage airfoil section; the lower surface trailing edge line is defined in the same way. The surface between the upper and lower trailing edge lines is called the trailing edge surface. The trailing edge surface should have a thickness of 0.003c to 0.015c, where c is the mean aerodynamic chord length.

[0015] The design process for the upper surface of the rear fuselage is as follows:

[0016] The first step is to determine the engine location and installation angle.

[0017] Ⅰ. Take the center of the fan intake surface as the reference point for the engine core, and use it for engine spatial positioning.

[0018] In step 1, the spanwise position of the engine core reference point is 0.08c to 0.18c from the vertical symmetry plane of the fuselage, the chordwise position is 1.10c to 1.50c from the top of the fuselage head, and the vertical position is 0.35d to 0.60d from the upper surface of the fuselage, where d is the diameter of the engine fan.

[0019] II. Select the direction of the engine thrust line. The engine thrust line is a straight line in space perpendicular to the fan intake surface and passing through its center, used to determine the direction of engine thrust. The angle between the direction of the engine thrust line and the x-axis is less than 5°.

[0020] Ⅲ. The vertical plane passing through the thrust line is taken as the engine's vertical symmetry plane. The engine's vertical symmetry plane refers to an imaginary plane passing through the thrust line and parallel to the z-axis. The lowest vertical point of the intersection line between the fan intake surface and the engine's vertical symmetry plane is taken as the lower edge point of the fan intake surface.

[0021] The second step is to determine the area to be modified on the upper surface of the rear fuselage.

[0022] The intersection of the engine's vertical symmetry plane and the clean upper surface is taken as the first control line of the upper surface; the intersection of the fuselage's vertical symmetry plane and the clean upper surface is taken as the second control line of the upper surface; the fuselage's vertical symmetry plane is the zx plane; the intersection of the fuselage's outer vertical plane and the clean upper surface is taken as the third control line of the upper surface. The outer vertical plane of the fuselage refers to an imaginary plane located outside the engine and parallel to the fuselage's vertical symmetry plane. The distance between the outer vertical plane of the fuselage and the fuselage's vertical symmetry plane is 0.08c to 0.18c. Within the engine's vertical symmetry plane, a circle is drawn with the lower edge of the fan intake surface as the center and the upper surface shaping area control radius Rm as the radius. The intersection of this circle and the first control line of the upper surface is taken as the first dividing point of the upper surface. The downstream part of this point is the shaping segment of the first control line of the upper surface, and the upstream part is the fixed segment. The shaping segment refers to the part that needs to change its shape in subsequent designs, and the fixed segment refers to the part that remains unchanged in subsequent designs. After the shape of the shaping segment changes, it still forms the first control line of the upper surface together with the fixed segment. The first boundary point of the upper surface is projected onto the second control line and the third control line of the upper surface, respectively, to obtain two projection points, namely the second boundary point and the third boundary point of the upper surface. The downstream part of the second boundary point of the upper surface is the modified segment of the second control line of the upper surface, and the downstream part of the third boundary point of the upper surface is the modified segment of the third control line of the upper surface; the upstream part of the second boundary point of the upper surface is the fixed segment of the second control line of the upper surface, and the upstream part of the third boundary point of the upper surface is the fixed segment of the third control line of the upper surface.

[0023] The upstream segment is a line segment along the positive x-axis with each dividing point as its origin, and the downstream segment is a line segment along the negative x-axis with each dividing point as its origin.

[0024] The downstream portion of the second boundary point on the upper surface and the downstream portion of the third boundary point on the upper surface are respectively the shaping segments of the second control line and the third control line on the upper surface, while the upstream portion is the fixed segment.

[0025] The third step is to reconstruct the control lines on the upper surface.

[0026] Project the lower edge of the fan intake surface onto the vertical symmetry plane of the fuselage and the vertical plane on the outer side of the fuselage, respectively, to obtain the first control point and the second control point on the upper surface.

[0027] Ⅰ. Reconstruct the modified segment of the first control line on the upper surface. A spline curve is generated using the first boundary point on the upper surface, the lower edge of the fan intake surface, and the downstream endpoint of the initial modified control line segment as the new modified segment of the first control line on the upper surface. The initial modified control line segment refers to the modified control line segment before reconstruction.

[0028] In the design of the rear fuselage upper surface based on the engine position and mounting angle, the new modified segment of the first control line of the upper surface should satisfy:

[0029] First, the tangent direction of the control line at the lower edge of the fan intake surface is the same as the direction of the engine thrust line;

[0030] Second, it is tangent to the fixed segment of the first control line on the upper surface at the first dividing point on the upper surface;

[0031] Third, the rear end of the spline is tangent to the modified section of the initial control line.

[0032] II. Reconstruct the modified segment of the second control line on the upper surface. Use the second boundary point of the upper surface, the first control point of the upper surface, and the downstream endpoint of the modified segment of the initial control line to generate a spline curve as the new modified segment of the second control line on the upper surface.

[0033] The new shaping segment of the second control line on the upper surface should meet the following requirements. ：

[0034] First, the tangent direction of the control line at the first control point on the upper surface is the same as the projection direction of the engine thrust line onto the vertical symmetry plane of the fuselage;

[0035] Second, it is tangent to the fixed segment of the second control line on the upper surface at the second dividing point on the upper surface;

[0036] Third, the rear end of the spline is tangent to the modified section of the initial control line.

[0037] Ⅲ. Reconstruct the modified segment of the third control line on the upper surface. Use the third boundary point of the upper surface, the second control point of the upper surface, and the downstream endpoint of the modified segment of the initial control line to generate a spline curve as the new modified segment of the third control line on the upper surface.

[0038] The new modified segment of the third control line on the upper surface should meet the following requirements:

[0039] First, the tangent direction of the control line at the second control point on the upper surface is the same as the projection direction of the engine thrust line onto the vertical plane outside the fuselage;

[0040] Second, it is tangent to the fixed segment of the third control line on the upper surface at the third dividing point on the upper surface;

[0041] Third, the rear end of the spline is tangent to the modified section of the initial control line.

[0042] The fourth step is to generate the upper surface of the new fuselage.

[0043] Using the modified upper surface first control line, upper surface second control line, upper surface third control line, as well as the fuselage upper surface trailing edge line, fuselage upper surface arc connecting line, fuselage upper surface side edge line, wing-fuselage boundary line, and fuselage leading edge line as control lines, a new fuselage upper surface is generated. The upper surface arc connecting line and the fuselage upper surface side edge line are both upper surface edge lines, with the upper surface arc connecting line connected to both the fuselage upper surface trailing edge line and the fuselage upper surface side edge line. The wing-fuselage boundary line is the boundary line between the fuselage upper surface and the upper wing surface. The fuselage leading edge line is the boundary line between the fuselage upper surface and the lower surface. The upper surface second control line, fuselage upper surface trailing edge line, fuselage upper surface arc connecting line, fuselage upper surface side edge line, wing-fuselage boundary line, and fuselage leading edge line are connected end-to-end to form a closed curve.

[0044] Step 2, nacelle / intake integrated with fuselage design:

[0045] Based on the initial axisymmetric nacelle of the engine and the upper surface of the new fuselage, the nacelle / intake duct and fuselage integration design were carried out.

[0046] The initial axisymmetric nacelle of the engine mentioned in step 2 refers to the outer casing and air intake portion of the turbofan engine's power nacelle. The outer casing portion and the air intake portion are bounded by the leading edge lip circle, and the initial axisymmetric nacelle of the engine is axisymmetric about the engine thrust line.

[0047] The design process for integrating the nacelle / air intake with the fuselage is as follows:

[0048] The first step is to determine the nacelle fixing area.

[0049] Within the forward plane of the nacelle, which is the plane containing the leading edge lip circle of the initial axisymmetric nacelle of the engine, the central angle of the leading edge lip circle is taken as the central angle θe of the fixed section of the nacelle. This central angle θe is symmetrical about the engine's vertical plane of symmetry, and the engine thrust line passes through the vertex of the central angle θe, which is between 150° and 180°. Utilizing the axisymmetric characteristic of the initial axisymmetric nacelle about the engine thrust line, the initial axisymmetric nacelle is divided into an upper and lower half according to the central angle θe. The upper half is the fixed section of the nacelle, and the boundary between the upper and lower half is formed by the upper fixed line of the air intake and the upper fixed line of the nacelle. The upper fixed line of the air intake and the upper fixed line of the nacelle are bounded by the leading edge lip circle. Taking advantage of the symmetry of the fan intake surface about the engine thrust line, the circular edge of the fan intake surface is divided into upper and lower parts according to the central angle θe, and the lower half is taken as the control line of the intake duct.

[0050] The second step is to generate the intake duct reshaping area.

[0051] In the plane of the front part of the nacelle, the intersection line of the filling curved surface formed by the closed leading edge lip and the upper surface of the new fuselage is taken as the bottom edge line of the air intake. The bottom edge line of the air intake is connected to the upstream endpoint of the upper fixed line of the air intake through a spline, which is the side edge line of the air intake.

[0052] In the design of integrating the nacelle / air intake with the fuselage:

[0053] The side line of the air intake satisfies the following conditions: first, it is tangent to the leading edge lip circle; second, it is tangent to the bottom line of the air intake in the plane of the front part of the nacelle.

[0054] The inlet front control line is formed by the side line and bottom line of the inlet. The multi-section curved surface generated by the inlet front control line, the inlet rear control line, and the upper fixed line of the inlet constitutes the inlet shaping area. The inlet shaping area satisfies the following conditions: first, it is tangent to the nacelle fixed area; second, it is tangent to the upper surface of the new fuselage.

[0055] The third step is to generate the nacelle shaping area.

[0056] The plane containing the maximum width of the initial axisymmetric nacelle of the engine is taken as the mid-plane of the nacelle. Within this plane, the intersection of this plane and the fixed line at the upper end of the nacelle is taken as the starting point. A line segment of length L1 is drawn as the nacelle forward taper control line. The angle between the nacelle forward taper control line and the vertical direction is the nacelle taper control angle γz1. L1 is 0.6d to 1.0d, and γz1 is -5° to 5°. It is defined that the nacelle taper control angle γz1 is positive when it deflects towards the vertical symmetry plane of the fuselage, and negative when it deflects away from it. A spline is drawn from the starting and ending points of the nacelle forward taper control line; this spline is the nacelle forward control line, which is tangent to the initial axisymmetric nacelle at its starting point.

[0057] The plane containing the trailing edge of the initial axisymmetric nacelle of the engine is taken as the rear plane of the nacelle. Within this plane, a line segment of length L2 is drawn as the starting point, using the downstream endpoint of the upper fixed line of the nacelle as the starting point, to form the nacelle rear taper control line. The angle between the nacelle rear taper control line and the vertical direction is the nacelle taper control angle γz2, where L2 is 0.6d to 1.0d and γz2 is -5° to 35°. It is defined that the nacelle taper control angle γz1 is positive when it deflects towards the vertical symmetry plane of the fuselage, and negative otherwise. A spline is drawn using the starting and ending points of the nacelle rear taper control line; this spline is the nacelle rear control line, which is tangent to the initial axisymmetric nacelle at its starting point.

[0058] Using the side line of the air intake, the forward control line of the nacelle, and the aft control line of the nacelle as control lines, a nacelle shaping area is generated. The nacelle shaping area satisfies the following conditions: first, it is tangent to the air intake shaping area; second, it is tangent to the nacelle fixing area.

[0059] The fourth step is to generate the nacelle rectification zone.

[0060] Take the vertical plane containing the upstream endpoint of the upper fixed line of the air intake as the planes on both sides of the nacelle. Take the intersection line of this plane with the nacelle shaping area, and another intersection line of this plane with the upper surface of the new fuselage. Make a first rounding with a rounding radius r1 between these two intersection lines. In the plane of the middle part of the nacelle, take the intersection line of this plane with the nacelle shaping area, and another intersection line of this plane with the upper surface of the new fuselage. Make a second rounding with a rounding radius r2 between these two intersection lines. In the plane of the rear part of the nacelle, take the intersection line of this plane with the nacelle shaping area, and another intersection line of this plane with the upper surface of the new fuselage. Make a third rounding with a rounding radius r3 between these two intersection lines.

[0061] The values ​​of r1, r2, and r3 are 0.05d to 0.15d.

[0062] Taking the lower end of the side line of the air intake as the starting point for rectification, and using the first, second, and third roundings as control lines, a nacelle rectification area is generated. The nacelle rectification area satisfies the following conditions: first, it is tangent to the upper surface of the new fuselage; second, it is tangent to the nacelle shaping area.

[0063] Step 3, Design of the tail nozzle in harmony with the fuselage:

[0064] The specific process is as follows:

[0065] The first step is to generate the tail nozzle lift zone.

[0066] The initial control point of the outer duct is taken as the intersection of the engine's vertical symmetry plane and the initial trailing edge line of the outer duct near the fuselage. The initial trailing edge line of the outer duct is the circular trailing edge line of the outer ring of the initial inner surface of the outer duct. Within the engine's vertical symmetry plane, the initial control point of the outer duct is moved vertically downwards by the nozzle lift height hex to form the inner lift control point, where the nozzle lift height hex is 0.01d to 0.04d. The intersection of the third rounded edges on both sides of the nacelle and the upper surface of the new fuselage is taken as the two lift control points. Using the inner lift control point and the two lift control points as control points, a lower control line for the lift zone is generated. This lower control line is tangent to the third rounded edges. The initial trailing edge line of the outer duct is divided by the intersection of the third rounded edges on both sides of the nacelle and the initial trailing edge line of the outer duct. The lower half of the division is taken as the upper control line for the lift zone. The nozzle lift zone is generated using the lower control line, the upper control line, and the third rounded edges on both sides of the nacelle as edge lines.

[0067] The second step is to generate the tail nozzle expansion zone.

[0068] The intersection of the engine's vertical plane of symmetry and the trailing edge line of the fuselage's upper surface is taken as the initial control point for the trailing edge. Within the engine's vertical plane of symmetry, the initial control point is moved downwards by a depth *dex* to form an inner control point for the extension. The depth *dex* should be less than the thickness of the fuselage's rear edge surface. On the trailing edge line of the fuselage's upper surface, the initial control point is moved inwards and outwards by 0.5 times the extension width *wex* to form control points at both ends of the extension. *wex* ranges from 0 to *d*. Using the inner control point and the control points at both ends of the extension, a rear control line for the extension area is generated within the fuselage's rear edge surface. This rear control line is tangent to the trailing edge line of the fuselage's upper surface at both ends. Connecting the lifting control points at both ends and the extension control points at both ends, a curved surface is drawn within the new fuselage's upper surface to form the control lines on both sides of the extension area.

[0069] The control lines on both sides of the extended area satisfy the following conditions: First, the upstream line is tangent to the intersection of the nacelle rectification area and the upper surface of the new fuselage; second, the downstream line is perpendicular to the trailing edge line of the upper surface of the fuselage. The tail nozzle extended area is generated using the lower control line of the lift area, the rear control line of the extended area, and the control lines on both sides of the extended area as edge lines; the tail nozzle extended area is tangent to the upper surface of the new fuselage.

[0070] The third step is to generate the rear edge surface of the new fuselage.

[0071] The extended area rear control line is an inner curve of the rear edge surface of the fuselage, and the two ends of the extended area rear control line are located on the rear edge line of the upper surface of the fuselage. Therefore, the extended area rear control line divides the rear edge surface of the fuselage into upper and lower parts, and the lower half is taken to form a new rear edge surface of the fuselage.

[0072] Through steps 1 to 3, the integrated aerodynamic shape design of the upper surface of the dorsal engine wing-body blended layout is completed.

[0073] The full fuselage is generated by symmetrically processing the semi-molded fuselage.

[0074] With this, the integrated flight-engine design with a dorsal engine and blended wing-body layout was completed.

[0075] Compared with the prior art, the beneficial effects achieved by the present invention are:

[0076] 1. The proposed "uplift" tail nozzle strategy solves the problem of coordinating the design of the circular nozzle and the upper surface of the fuselage in the dorsal engine blended wing-body layout. It ensures the circular tail nozzle while avoiding the ablation of the fuselage upper surface by the jet. Figure 7 and Figure 8 , Figure 9 and Figure 10 The vertical symmetry plane flow characteristics of the back-supported and back-mounted engines under high thrust on the ground at two different speeds were compared. It can be seen that the back-mounted design of the engine does not affect the power jet characteristics, and the "lifted" design also effectively avoids the jet from burning the surface of the fuselage.

[0077] 2. The proposed integrated flight-engine design method employs methods such as fuselage surface modification and short air intake design to ensure air intake quality. From Figure 8 and Figure 10 It can be seen that the dorsal engine blended wing-body layout still maintains unobstructed air intake even with a large air intake volume. Figure 11 This application compares the proposed back-mounted (short intake) design with that in the literature (Plas A, Crichton D, Sargeant M, et al.).

[0078] Performance of a boundary layer ingesting(BLI)propulsion system[C].USA:

[0079] In the Proceedings of the 45th AIAA Aerospace Sciences Meeting and Exhibit, 2007:8-11., the total pressure distribution of the intake surface of a medium-length inlet-type boundary layer intake layout engine shows that the design proposed in this application can significantly reduce the proportion of the distortion region, which helps to improve intake efficiency.

[0080] 3. The proposed integrated flight-engine design method places the entire engine on the rear fuselage, with the exhaust nozzle not exposed. From Figure 8 and Figure 10 It can be seen that there is ample space at the rear edge of the fuselage to accommodate the tail rudder, and the fuselage can also shield the engine exhaust noise.

[0081] 4. The proposed integrated flight-engine design method is based on the physical principle that the dorsal engine mainly affects the rear fuselage. By dividing the area into fixed area, shaping area and transition area, it avoids global parametric modeling and improves design efficiency.

[0082] 5. The proposed integrated flight-engine design method achieves three-dimensional integrated design of parameters such as engine position and installation angle with fuselage shape control parameters. This overcomes the limitations of existing design methods, which mainly focus on separate design and two-dimensional design of the air intake / nacelle. Figures 12 to 15 A comparison of the aerodynamic performance of the back-support and dome-mounted designs at a 1.5° angle of attack in high-altitude, high-speed conditions is presented. The cruise design point in the figure represents the operating point where thrust and drag are balanced, i.e., the main flight state. It can be seen that the dome-mounted design improves the lift-to-drag ratio by 1.0 without significant changes in lift and torque characteristics, demonstrating the feasibility and rationality of the proposed design method. Attached Figure Description

[0083] Figure 1It is a clean configuration with a blended wing-body layout.

[0084] Figure 2 It is the power nacelle of a turbofan engine.

[0085] Figure 3 This is a schematic diagram of the design of the upper surface of the rear fuselage.

[0086] Figure 4 It is a cross-sectional view of the engine's vertical symmetry plane in the design of the upper surface of the rear fuselage.

[0087] Figure 5 It is a top view showing the relative positions of the engine and fuselage.

[0088] Figure 6 This is a schematic diagram of the nacelle / air intake integrated with the fuselage design.

[0089] Figure 7 This is a schematic diagram of the nacelle and fuselage integration design within the plane where the engine's initial axisymmetric nacelle has its maximum width.

[0090] Figure 8 It is a schematic diagram of the nacelle integrated with the fuselage in the vertical plane where the lower end of the side line of the air intake is located.

[0091] Figure 9 This is a schematic diagram illustrating the coordinated design of the tail nozzle and fuselage.

[0092] Figure 10 The flow characteristics of the back-supported layout with vertical symmetry are as follows: flight altitude 0, incoming Mach number 0.2, flight angle of attack 8°, and low-pressure compressor relative speed 72.9%.

[0093] Figure 11 The flow characteristics of the vertically symmetrical plane in the back-mounted configuration are as follows: flight altitude is 0, incoming Mach number is 0.2, flight angle of attack is 8°, and low-pressure compressor relative speed is 72.9%.

[0094] Figure 12 The flow characteristics of the back-supported layout with vertical symmetry are as follows: flight altitude 0, incoming Mach number 0.2, flight angle of attack 8°, and low-pressure compressor relative speed 92.3%.

[0095] Figure 13 The flow characteristics of the vertical symmetry plane in the back-mounted layout are as follows: flight altitude is 0, incoming Mach number is 0.2, flight angle of attack is 8°, and low-pressure compressor relative speed is 92.3%.

[0096] Figure 14 It refers to the total pressure distribution of the fan intake surface in both backpack and long-intake boundary layer intake configurations, where... Figure 14 A is a backpack style. Figure 14 b is the long intake type.

[0097] Figure 15 This is a comparison of the lift coefficients of the back-support and back-mounted designs at different speeds under the following conditions: flight altitude of 11.5 km, incoming Mach number of 0.85, and flight angle of attack of 1.5°.

[0098] Figure 16 This is a comparison of the pitching moment coefficients of the back-support and back-mounted design schemes at different speeds under the following conditions: flight altitude of 11.5 km, incoming Mach number of 0.85, and flight angle of attack of 1.5°.

[0099] Figure 17 This is a comparison of the drag coefficients of the back-support and back-mounted design schemes at different speeds under the following conditions: flight altitude of 11.5 km, incoming Mach number of 0.85, and flight angle of attack of 1.5°.

[0100] Figure 18 This is a comparison of the lift-to-drag ratio of the back-support and back-mounted design schemes at different speeds under the following conditions: flight altitude of 11.5 km, incoming Mach number of 0.85, and flight angle of attack of 1.5°.

[0101] Figure 19 This is a flowchart of the present invention.

[0102] In the diagram: 1. Clean configuration with blended wing-body layout; 2. Turbofan engine nacelle; 3. Trailing edge line of upper fuselage surface; 4. Trailing edge surface of fuselage surface; 5. Engine core reference point; 6. Engine thrust line; 7. Engine vertical symmetry plane; 8. Lower edge of fan intake surface; 9. First control line of upper surface; 10. Vertical symmetry plane of fuselage surface; 11. Second control line of upper surface; 12. Vertical plane of outer fuselage surface; 13. Third control line of upper surface; 14. First boundary point of upper surface; 15. Second boundary point of upper surface; 16. 17. Third boundary point of the upper surface; 18. First control point of the upper surface; 19. Second control point of the upper surface; 20. Arc-shaped connecting line of the upper fuselage surface; 21. Side edge line of the upper fuselage surface; 22. Wing-fuselage boundary line; 23. Leading edge line of the fuselage; 24. Upper surface of the new fuselage; 25. Initial axisymmetric nacelle; 26. Leading edge lip circle; 27. Forward plane of the nacelle; 28. Nacelle fixing area; 29. ​​Upper fixing line of the air intake; 30. Aft control line of the air intake; 31. Bottom edge line of the air intake; 2. Inlet side edge line; 33. Inlet forward control line; 34. Inlet shaping area; 35. Nacelle center plane; 36. Nacelle forward taper control line; 37. Nacelle forward control line; 38. Nacelle rear plane; 39. Nacelle rear taper control line; 40. Nacelle rear control line; 41. Nacelle shaping area; 42. Nacelle side planes; 43. First rounding; 44. Second rounding; 45. Third rounding; 46. Rectification starting point; 47. Nacelle rectification area; 48. Initial trailing edge line of bypass; 49. Initial control line of bypass. Points; 50. Lift inner control point; 51. Lift end control points; 52. Lower control line of the lift zone; 53. Upper control line of the lift zone; 54. Tail nozzle lift zone; 55. Trailing edge initial control point; 56. Extended inner control point; 57. Extended end control points; 58. Extended zone rear control line; 59. Extended zone side control lines; 60. Tail nozzle extended zone; 61. New fuselage rear edge surface; 62. Intake distortion area; 63. Cruise design point; 64. Back-supported scheme data curve; 65. Back-mounted scheme data curve.

[0103] c: Mean aerodynamic chord length;

[0104] d: Engine fan diameter;

[0105] Rm: Control radius of the upper surface shaping area;

[0106] θe: Central angle of the fixed section of the nacelle;

[0107] L: Length of the nacelle taper control line;

[0108] γz: Nacelle taper control angle;

[0109] r: radius of the fillet;

[0110] hex: height of the tail nozzle lift;

[0111] dex: extends depth;

[0112] wex: Expands the width;

[0113] NL: Relative speed of the low-pressure compressor;

[0114] C L : Lift coefficient;

[0115] C D Drag coefficient;

[0116] C m Pitch moment coefficient;

[0117] K: Rise-to-drag ratio. Detailed Implementation

[0118] This embodiment is a flight-engine integrated design method that adopts a dorsal engine blended wing-body layout, including three aspects: the design of the upper surface of the rear fuselage based on the engine position and installation angle, the nacelle / inlet blended design with the fuselage, and the coordinated design of the tail nozzle with the fuselage.

[0119] Due to the overall geometric symmetry of the aircraft, this invention uses a one-side half-model as an example to illustrate the specific steps of the integrated flight-engine design method for a dorsal-engine blended wing-body layout:

[0120] Before introducing the specific implementation method, the coordinate system and direction definitions used in this embodiment will be explained first. In this embodiment, the nose apex of the fuselage is defined as the origin; the chord direction of the fuselage is the x-axis direction, with the direction pointing from the origin to the tail of the fuselage as the positive x-axis direction, or downstream direction, and vice versa as the negative x-axis direction, or upstream direction; the spanwise direction of the fuselage is the y-axis direction, with the direction pointing from the origin to the wingtip as the positive y-axis direction, or outward direction, and vice versa as inward direction. The direction passing through the origin and perpendicular to the xy plane is the z-axis direction, or vertical direction, with upward being defined as positive.

[0121] Step 1, Design of the upper surface of the rear fuselage based on the engine position and mounting angle:

[0122] For the clean configuration 1 of the blended wing-body layout without engine, combined with the turbofan engine power nacelle 2, the upper surface of the rear fuselage is designed.

[0123] The clean configuration 1 of the blended wing-body layout mentioned in step 1 refers to an aerodynamic shape consisting only of the fuselage and wings, excluding components such as engines, tail fins, control surfaces, and landing gear. The turbofan engine nacelle 2 refers to a simplified shape that simplifies the engine fan and turbine into intake and exhaust surfaces. In this embodiment, the trailing edge line of the lower fuselage surface of the clean configuration 1 is parallel to the trailing edge line 3 of the upper fuselage surface. The trailing edge line of the upper fuselage surface is a curve formed by connecting the trailing edge points of the upper surface of the fuselage section airfoil, and the trailing edge line of the lower surface is defined in the same way. The surface between the trailing edge lines of the upper and lower surfaces is the trailing edge surface 4. In this embodiment, the trailing edge surface 4 should have a thickness of 0.003c to 0.015c, where c is the mean aerodynamic chord length.

[0124] The specific process for designing the upper surface of the rear fuselage is as follows:

[0125] The first step is to determine the engine location and installation angle.

[0126] I. The center of the fan intake surface is taken as the engine core reference point 5 for engine spatial positioning. In this embodiment, the spanwise position of the engine core reference point 5 is 0.08c to 0.18c from the vertical symmetry plane 10 of the fuselage, the chordwise position is 1.10c to 1.50c from the top vertex of the fuselage head, and the vertical position is 0.35d to 0.60d from the upper surface of the fuselage, where d is the diameter of the engine fan.

[0127] II. The direction of the engine thrust line 6 is selected. The engine thrust line refers to an imaginary spatial straight line perpendicular to the fan intake surface and passing through its center, used to determine the thrust direction of the engine. In this embodiment, the angle formed between the direction of the engine thrust line 6 and the x-axis direction is less than 5°.

[0128] Ⅲ. The vertical plane passing through the thrust line 6 is taken as the engine's vertical symmetry plane 7. The engine's vertical symmetry plane 7 refers to an imaginary plane passing through the thrust line 6 and parallel to the z-axis. The lowest vertical point of the intersection line between the fan intake surface and the engine's vertical symmetry plane 7 is taken as the lower edge point 8 of the fan intake surface.

[0129] The second step is to determine the area to be modified on the upper surface of the rear fuselage.

[0130] The intersection of the engine's vertical symmetry plane 7 and the clean upper surface is designated as the first control line 9 of the upper surface; the intersection of the fuselage's vertical symmetry plane 10 and the clean upper surface is designated as the second control line 11 of the upper surface, where the fuselage's vertical symmetry plane 10 is the zx plane; and the intersection of the fuselage's outer vertical plane 12 and the clean upper surface is designated as the third control line 13 of the upper surface, where the outer vertical plane 12 refers to an imaginary plane located outside the engine and parallel to the fuselage's vertical symmetry plane 10. In this embodiment, the distance between the outer vertical plane 12 and the fuselage's vertical symmetry plane 10 is 0.08c to 0.18c. Within the vertical symmetry plane 7 of the engine, a circle is drawn with the lower edge point 8 of the fan intake surface as the center and the upper surface shaping area control radius Rm as the radius. The intersection of this circle and the first control line 9 on the upper surface is taken as the first boundary point 14 of the upper surface. The downstream part of this point is the shaping segment of the first control line 9 on the upper surface, and the upstream part is the fixed segment. The shaping segment refers to the part whose shape needs to be changed in subsequent designs, and the fixed segment refers to the part that remains unchanged in subsequent designs. After the shape of the shaping segment is changed, it still forms the first control line 9 on the upper surface together with the fixed segment. The subsequent descriptions of "shaping segment" and "fixed segment" in this embodiment are consistent with the above description. The first boundary point 14 on the upper surface is projected onto the second control line 11 and the third control line 13 on the upper surface, respectively, resulting in two projection points: the second boundary point 15 and the third boundary point 16 on the upper surface. The downstream portions of the second boundary point 15 and the third boundary point 16 are the modified sections of the second control line 11 and the third control line 13 on the upper surface, respectively, while the upstream portion is a fixed section. In this embodiment, the control radius Rm of the upper surface modification area can be taken as 2d to 5d, where d is the diameter of the engine fan.

[0131] The third step is to reconstruct the control lines on the upper surface.

[0132] Project the lower edge point 8 of the fan intake surface onto the vertical symmetry plane 10 of the fuselage and the vertical plane 12 on the outer side of the fuselage, respectively, to obtain the first control point 17 and the second control point 18 on the upper surface.

[0133] Ⅰ. Reconstruct the modified segment of the first control line 9 on the upper surface. Use the first dividing point 14 on the upper surface, the lower edge point 8 of the fan air intake surface, and the downstream endpoint of the initial control line modified segment to generate a spline curve as the new modified segment of the first control line 9 on the upper surface. The initial control line modified segment refers to the control line modified segment before reconstruction.

[0134] In the design of the rear fuselage upper surface based on the engine position and mounting angle, the new modified segment of the first control line 9 on the upper surface should meet the following requirements:

[0135] First, the tangent direction of the control line at point 8, the lower edge of the fan intake surface, is the same as the direction of the engine thrust line 6.

[0136] Second, it is tangent to the fixed segment of the first control line 9 on the upper surface at the first dividing point 14 on the upper surface;

[0137] Third, the rear end of the spline is tangent to the modified section of the initial control line.

[0138] II. Reconstruct the modified segment of the second control line 11 on the upper surface. Use the second boundary point 15 on the upper surface, the first control point 17 on the upper surface, and the downstream endpoint of the modified segment of the initial control line to generate a spline curve as the new modified segment of the second control line 11 on the upper surface.

[0139] The new shaping segment of the second control line 11 on the upper surface should meet the following requirements:

[0140] First, the tangent direction of the control line at the first control point 17 on the upper surface is the same as the projection direction of the engine thrust line 6 onto the vertical symmetry plane 10 of the fuselage.

[0141] Second, at the second dividing point 15 on the upper surface, it is tangent to the fixed segment of the second control line 11 on the upper surface;

[0142] Third, the rear end of the spline is tangent to the modified section of the initial control line.

[0143] III. Reconstruct the modified segment of the third control line 13 on the upper surface. Generate a spline curve as the new modified segment of the third control line 13 on the upper surface using the third boundary point 16 on the upper surface, the second control point 18 on the upper surface, and the downstream endpoint of the modified segment of the initial control line.

[0144] The new shaping segment of the third control line 13 on the upper surface should meet the following requirements:

[0145] First, the tangent direction of the control line at the second control point 18 on the upper surface is the same as the projection direction of the engine thrust line 6 onto the vertical plane 12 on the outer side of the fuselage.

[0146] Second, at the third dividing point 16 on the upper surface, it is tangent to the fixed segment of the third control line 13 on the upper surface;

[0147] Third, the rear end of the spline is tangent to the modified section of the initial control line.

[0148] The fourth step is to generate the upper surface of the new fuselage.

[0149] Using the modified upper surface control lines 9, 11, and 13, as well as the fuselage upper surface trailing edge line 3, arc-shaped connecting line 19, side edge line 20, wing-fuselage boundary line 21, and leading edge line 22 as control lines, a new fuselage upper surface 23 is generated. The arc-shaped connecting line 19 and the side edge line 20 are both upper surface edge lines, with the arc-shaped connecting line 19 connecting to both the upper surface trailing edge line 3 and the side edge line 20. The wing-fuselage boundary line 21 is the boundary between the upper surface of the fuselage and the upper surface of the wing. The leading edge line 22 is the boundary between the upper and lower surfaces of the fuselage. The upper surface second control line 11, the fuselage upper surface trailing edge line 3, the fuselage upper surface arc connecting line 19, the fuselage upper surface side edge line 20, the wing-fuselage dividing line 21, and the fuselage leading edge line 22 are connected end to end to form a closed curve.

[0150] Step 2, nacelle / intake integrated with fuselage design:

[0151] Based on the initial axisymmetric nacelle 24 of the engine and the upper surface 23 of the new fuselage, the nacelle / intake duct and fuselage integration design is carried out.

[0152] The initial axisymmetric nacelle 24 of the engine mentioned in step 2 refers to the outer casing and air intake portion of the turbofan engine power nacelle 2. The outer casing portion and the air intake portion are bounded by the leading edge lip circle 25. The initial axisymmetric nacelle 24 of the engine is axisymmetric about the engine thrust line 6.

[0153] The design process for integrating the nacelle / air intake with the fuselage is as follows:

[0154] The first step is to determine the nacelle fixing area.

[0155] Within the front plane 26 of the nacelle, which is the plane containing the leading edge lip circle 25 of the initial axisymmetric nacelle 24 of the engine, the central angle of the leading edge lip circle 25 is taken as the central angle θe of the nacelle fixed section. The central angle θe of the nacelle fixed section is symmetrical about the vertical symmetry plane 7 of the engine, and the engine thrust line 6 passes through the vertex of the central angle θe. In this embodiment, θe can be 150° to 180°. Utilizing the characteristic of the initial axisymmetric nacelle 24 of the engine being symmetrical about the engine thrust line 6, the initial axisymmetric nacelle 24 of the engine is divided into upper and lower parts according to the central angle θe. The upper part is the nacelle fixed area 27, and the dividing line between the upper and lower parts is formed by the upper end fixed line 28 of the air intake and the upper end fixed line 29 of the nacelle. The upper end fixed line 28 of the air intake and the upper end fixed line 29 of the nacelle are bounded by the leading edge lip circle 25. Taking advantage of the symmetry of the fan intake surface about the engine thrust line, the circular edge of the fan intake surface is divided into upper and lower parts according to the central angle θe, and the lower half is taken as the rear control line 30 of the intake duct.

[0156] The second step is to generate the intake duct reshaping area.

[0157] Within the forward plane 26 of the nacelle, the intersection of the filled curved surface formed by the closed leading edge lip circle 25 and the upper surface 23 of the new fuselage is taken as the bottom edge line 31 of the air intake. The bottom edge line 31 of the air intake is connected to the upstream endpoint of the upper fixed line 28 of the air intake through a spline, which is the side edge line 32 of the air intake.

[0158] In the design of integrating the nacelle / air intake with the fuselage:

[0159] The side line 32 of the air intake satisfies the following conditions: first, it is tangent to the leading edge lip circle 25; second, it is tangent to the bottom line 31 of the air intake within the front plane 26 of the nacelle.

[0160] The inlet front control line 33 is formed by the inlet side line 32 and the inlet bottom line 31. The inlet front control line 33, the inlet rear control line 30 and the inlet upper fixed line 28 form a multi-section curved surface as the inlet shaping area 34. The inlet shaping area 34 satisfies the following conditions: first, it is tangent to the nacelle fixed area 27; second, it is tangent to the upper surface 23 of the new fuselage.

[0161] The third step is to generate the nacelle shaping area.

[0162] The plane containing the maximum width of the initial axisymmetric nacelle 24 of the engine is taken as the nacelle mid-plane 35. Within this plane, the intersection of this plane and the upper fixed line 29 of the nacelle is taken as the starting point, and a line segment of length L1 is drawn as the nacelle forward taper control line 36. The angle between the nacelle forward taper control line 36 and the vertical direction is the nacelle taper control angle γz1. In this embodiment, L1 can be 0.6d to 1.0d, and γz1 can be -5° to 5°. It is defined that the nacelle taper control angle γz1 is positive when it deflects towards the vertical symmetry plane of the fuselage, and negative when it deflects away from it. A spline is drawn from the starting point and the ending point of the nacelle forward taper control line 36 as the nacelle forward control line 37. The nacelle forward control line 37 is tangent to the initial axisymmetric nacelle 24 at its starting point.

[0163] The plane containing the trailing edge of the initial axisymmetric nacelle of the engine is taken as the rear plane 38 of the nacelle. Within this plane, the downstream endpoint of the upper fixed line 29 of the nacelle is taken as the starting point, and a line segment of length L2 is drawn as the rear taper control line 39 of the nacelle. The angle between the rear taper control line 39 of the nacelle and the vertical direction is the nacelle taper control angle γz2. In this embodiment, L2 can be 0.6d to 1.0d, and γz2 can be -5° to 35°. It is defined that the nacelle taper control angle γz1 is positive when it deflects towards the vertical symmetry plane of the fuselage, and negative when it deflects away from it. A spline is drawn from the starting point and the ending point of the rear taper control line 39 of the nacelle as the rear control line 40 of the nacelle. The rear control line 40 of the nacelle is tangent to the initial axisymmetric nacelle 24 at its starting point.

[0164] Using the side line 32 of the air intake, the front control line 37 of the nacelle, and the rear control line 40 of the nacelle as control lines, a nacelle shaping area 41 is generated. The nacelle shaping area 41 satisfies the following conditions: first, it is tangent to the air intake shaping area 34; second, it is tangent to the nacelle fixing area 27.

[0165] The fourth step is to generate the nacelle rectification zone.

[0166] Take the vertical plane containing the upstream endpoint of the upper fixed line 28 of the air intake as the planes 42 on both sides of the nacelle. Take the intersection line of this plane with the nacelle shaping area 41, and the intersection line of this plane with the upper surface 23 of the new fuselage. Between these two intersection lines, make a first rounding radius 43 with a rounding radius r1. In the middle plane 35 of the nacelle, take the intersection line of this plane with the nacelle shaping area 41, and the intersection line of this plane with the upper surface 23 of the new fuselage. Between these two intersection lines, make a second rounding radius 44 with a rounding radius r2. In the rear plane 38 of the nacelle, take the intersection line of this plane with the nacelle shaping area 41, and the intersection line of this plane with the upper surface 23 of the new fuselage. Between these two intersection lines, make a third rounding radius 45 with a rounding radius r3.

[0167] In this embodiment, r1 can be 0.05d to 0.15d; r2 can be 0.05d to 0.15d; and r3 can be 0.05d to 0.15d.

[0168] Taking the lower end point of the side line 32 of the air intake as the rectification starting point 46, and using the first rounded corner 43, the second rounded corner 44 and the third rounded corner 45 as control lines, a nacelle rectification area 47 is generated. The nacelle rectification area 47 satisfies the following: first, it is tangent to the upper surface 23 of the new fuselage; second, it is tangent to the nacelle shaping area 41.

[0169] Step 3, Design of the tail nozzle in harmony with the fuselage:

[0170] The specific process is as follows:

[0171] The first step is to generate the tail nozzle lift zone.

[0172] The intersection of the engine's vertical symmetry plane 7 and the initial trailing edge line 48 of the duct near the fuselage is taken as the initial control point 49 of the duct. The initial trailing edge line 48 of the duct is the circular trailing edge line of the outer ring of the initial inner surface of the duct. Within the engine's vertical symmetry plane 7, the initial control point 49 of the duct is moved vertically downwards by the nozzle lift height hex to form the inner lift control point 50. In this embodiment, the nozzle lift height hex can be 0.01d to 0.04d. The intersection of the third rounded edges 45 on both sides of the nacelle and the upper surface 23 of the new fuselage is taken as the lift end control point 51. Using the inner lift control point 50 and the lift end control point 51 as control points, the lower control line 52 of the lift area is generated. The lower control line 52 of the lift area is tangent to the third rounded edges 45. The initial trailing edge line 48 of the duct is divided by the intersection of the third rounded edges 45 on both sides of the nacelle and the initial trailing edge line 48 of the duct. The lower half of the division is taken as the upper control line 53 of the lift area. Using the lower control line 52 of the lift zone, the upper control line 53 of the lift zone, and the third rounded edges 45 on both sides of the nacelle as the boundary lines, the tail nozzle lift zone 54 is generated.

[0173] The second step is to generate the tail nozzle expansion zone.

[0174] The intersection of the engine's vertical symmetry plane 7 and the trailing edge line 3 of the fuselage upper surface is taken as the initial trailing edge control point 55. Within the engine's vertical symmetry plane 7, the initial trailing edge control point 55 is moved downwards by an extension depth *dex* to form an extended inner control point 56. The extension depth *dex* should be less than the thickness of the fuselage's rear edge surface 4. On the trailing edge line 3 of the fuselage upper surface, the initial trailing edge control point 55 is moved inwards and outwards by 0.5 times the extension width *wex* to form extended end control points 57. In this embodiment, *wex* can be 0 to *d*. Using the extended inner control point 56 and the extended end control points 57 as control points, an extended area rear control line 58 is generated within the fuselage's rear edge surface 4. The extended area rear control line 58 satisfies the condition that both ends are tangent to the fuselage upper surface trailing edge line 3. Connecting the lifting end control point 51 and the extended end control points 57, a curved surface is drawn within the new fuselage upper surface 23 to form the extended area side control lines 59.

[0175] The control lines 59 on both sides of the extended area satisfy the following conditions: first, the upstream line is tangent to the intersection of the nacelle rectification area 47 and the upper surface 23 of the new fuselage; second, the downstream line is perpendicular to the trailing edge line 3 of the upper surface of the fuselage. Using the lower control line 52 of the lift area, the rear control line 58 of the extended area, and the control lines 59 on both sides of the extended area as edge lines, a tail nozzle extended area 60 is generated, which satisfies the condition of being tangent to the upper surface 23 of the new fuselage.

[0176] The third step is to generate the rear edge surface of the new fuselage.

[0177] Since the extended area rear control line 58 is an inner curve of the rear edge surface 4 of the fuselage, and the two endpoints of the extended area rear control line 58 are located on the rear edge line 3 of the upper surface of the fuselage, the extended area rear control line 58 can divide the rear edge surface 4 of the fuselage into upper and lower parts, and take the lower half to form a new rear edge surface 61 of the fuselage.

[0178] Through steps 1 to 3, the integrated aerodynamic shape design of the upper surface of the dorsal engine wing-body blended layout is completed.

[0179] The full fuselage can be generated by symmetrically processing the half-molded fuselage.

[0180] This completes the integrated flight-engine design of the dorsal engine blended wing-body layout.

[0181] To verify the rationality of the proposed integrated flight-engine design method, a numerical simulation comparative study was conducted on the blended wing-body layout schemes of dorsal-supported and dorsal-mounted engines under high-speed and low-speed conditions. Both schemes used the same engine type and had the same forward fuselage and wing shape, differing only in engine arrangement. The low-speed calculation conditions were: flight altitude 0, incoming Mach number 0.2, flight angle of attack 8°, and low-pressure compressor relative speeds of 72.9% and 92.3%. The high-speed calculation conditions were: flight altitude 11.5 km, incoming Mach number 0.85, flight angle of attack 1.5°, and low-pressure compressor relative speeds ranging from 73% to 95%. Figures 7 to 10 The vertical symmetry plane flow characteristics of the engine under low speed and high thrust are presented for two different configurations. From the intake perspective, it can be seen that the dorsal-mounted engine with blended wing-body configuration still maintains unobstructed airflow channels even with large air intake volumes. From the exhaust perspective, it can be seen that the exhaust channels of the dorsal-mounted engine with blended wing-body configuration are unobstructed, and the exhaust is uniform and symmetrical, with no significant difference from the dorsal-supported configuration. Figures 12 to 15 A comparison of the aerodynamic performance of the two schemes at high speed is presented. It can be seen that the back-mounted design scheme improves the lift-to-drag ratio by 1.0 without significant changes in lift and torque characteristics, which demonstrates the rationality of the proposed integrated flight-engine design method.

Claims

1. A flight-engine integrated design method for a dorsal-mounted engine blended wing-body layout, characterized in that, It includes three aspects: the design of the upper surface of the rear fuselage based on the engine position and installation angle, the design of the nacelle / inlet integrated with the fuselage, and the design of the tail nozzle coordinated with the fuselage. Taking one side of the longitudinal axis of the fuselage model as an example, the integrated flight-engine design method of the dorsal engine wing-body layout is described. The specific process is as follows: Step 1, Design of the upper surface of the rear fuselage based on the engine position and mounting angle: For the clean configuration of the blended wing-body layout without an engine, the upper surface of the rear fuselage is designed in conjunction with the turbofan engine power nacelle. The design process for the upper surface of the rear fuselage is as follows: The first step is to determine the engine location and installation angle: Ⅰ. Take the center of the fan intake surface as the reference point for the engine core, and use it for engine spatial positioning; II. Select the direction of the engine thrust line. The engine thrust line is a straight line in space that is perpendicular to the fan intake surface and passes through its center. It is used to determine the thrust direction of the engine. The angle between the direction of the engine thrust line and the x-axis direction is less than 5°. Ⅲ. Take the vertical plane passing through the thrust line as the engine vertical symmetry plane. The engine vertical symmetry plane refers to the imaginary plane passing through the thrust line and parallel to the z-axis; take the lowest vertical point of the intersection line between the fan intake surface and the engine vertical symmetry plane as the lower edge point of the fan intake surface. The second step is to determine the area to be modified on the upper surface of the rear fuselage: The intersection of the engine's vertical symmetry plane and the clean upper surface is taken as the first control line of the upper surface; the intersection of the fuselage's vertical symmetry plane and the clean upper surface is taken as the second control line of the upper surface; the fuselage's vertical symmetry plane is the zx plane; the intersection of the fuselage's outer vertical plane and the clean upper surface is taken as the third control line of the upper surface, the outer vertical plane of the fuselage being an imaginary plane located outside the engine and parallel to the fuselage's vertical symmetry plane; the distance between the outer vertical plane of the fuselage and the fuselage's vertical symmetry plane is 0.08c~0.18c, where c is the average aerodynamic chord length; within the engine's vertical symmetry plane, a circle is drawn with the lower edge of the fan intake surface as the center and the upper surface shaping area control radius Rm as the radius, the intersection of this circle and the first control line of the upper surface is taken as the first boundary point of the upper surface, and the downstream part of this point is the first boundary point of the upper surface. The control line has a modified segment upstream of a fixed segment. The modified segment refers to the part whose shape needs to be changed in subsequent design, while the fixed segment refers to the part that remains unchanged in subsequent design. After the shape is changed, the modified segment still forms the first control line on the upper surface together with the fixed segment. The first boundary point of the upper surface is projected onto the second and third control lines of the upper surface, respectively, to obtain two projection points, which are the second and third boundary points of the upper surface. The downstream part of the second boundary point of the upper surface is the modified segment of the second control line of the upper surface, and the downstream part of the third boundary point of the upper surface is the modified segment of the third control line of the upper surface. The upstream part of the second boundary point of the upper surface is the fixed segment of the second control line of the upper surface, and the upstream part of the third boundary point of the upper surface is the fixed segment of the third control line of the upper surface. The upstream is a line segment in the positive x-axis direction with each dividing point as the origin, and the downstream is a line segment in the negative x-axis direction with each dividing point as the origin. The control radius Rm of the upper surface shaping area is 2d~5d; where d is the diameter of the engine fan. Third step, reconstruct the upper surface control lines: By reconstructing the modified sections of the first control line, the second control line, and the third control line on the upper surface, the lower edge of the fan intake surface is projected onto the vertical symmetry plane of the fuselage and the vertical plane on the outer side of the fuselage, respectively, and the projection points are the first control point and the second control point on the upper surface, respectively. Ⅰ Reconstruct the modified segment of the first control line on the upper surface; use the first dividing point of the upper surface, the lower edge point of the fan air intake surface and the downstream endpoint of the modified segment of the initial control line to generate a spline curve as the new modified segment of the first control line on the upper surface. The modified segment of the initial control line refers to the modified segment of the control line before reconstruction. II. Reconstruct the modified segment of the second control line on the upper surface; use the second boundary point of the upper surface, the first control point of the upper surface, and the downstream endpoint of the modified segment of the initial control line to generate a spline curve as a new modified segment of the second control line on the upper surface; Ⅲ Reconstruct the modified segment of the third control line on the upper surface; use the third boundary point of the upper surface, the second control point of the upper surface, and the downstream endpoint of the modified segment of the initial control line to generate a spline curve as a new modified segment of the third control line on the upper surface; Step 4: Generate the upper surface of the new fuselage: Using the modified upper surface first control line, upper surface second control line, upper surface third control line, as well as the fuselage upper surface trailing edge line, fuselage upper surface arc connecting line, fuselage upper surface side edge line, wing-fuselage boundary line, and fuselage leading edge line as control lines, a new fuselage upper surface is generated; the upper surface arc connecting line and the fuselage upper surface side edge line are both upper surface edge lines, wherein the upper surface arc connecting line is connected to the fuselage upper surface trailing edge line and the fuselage upper surface side edge line respectively; the wing-fuselage boundary line refers to the boundary line between the fuselage upper surface and the wing upper surface; the fuselage leading edge line is the boundary line between the fuselage upper surface and the lower surface; the upper surface second control line, fuselage upper surface trailing edge line, fuselage upper surface arc connecting line, fuselage upper surface side edge line, wing-fuselage boundary line, and fuselage leading edge line are connected end to end to form a closed curve; Step 2, nacelle / intake integrated with fuselage design: Based on the initial axisymmetric engine nacelle and the aforementioned upper surface of the new fuselage, a design integrating the nacelle / intake with the fuselage is carried out; the specific process is as follows: The first step is to determine the nacelle anchoring area: Within the forward plane of the nacelle, the forward plane of the nacelle is the plane containing the leading edge lip circle of the initial axisymmetric nacelle of the engine. The central angle of the leading edge lip circle is taken as the central angle θe of the fixed section of the nacelle. The central angle θe of the fixed section of the nacelle is symmetrical about the vertical symmetry plane of the engine. The engine thrust line passes through the vertex of the central angle θe, which is 150°~180°. Utilizing the characteristic of the initial axisymmetric nacelle of the engine being symmetrical about the engine thrust line, the initial axisymmetric nacelle of the engine is divided into an upper half and a lower half according to the central angle θe. The upper half is the fixed area of ​​the nacelle. The dividing line between the upper half and the lower half is formed by the upper fixed line of the air intake and the upper fixed line of the nacelle. The upper fixed line of the air intake and the upper fixed line of the nacelle are bounded by the leading edge lip circle. Utilizing the characteristic of the fan intake surface being symmetrical about the engine thrust line, the circular edge of the fan intake surface is divided into upper and lower parts according to the central angle θe. The lower half is taken as the rear control line of the air intake. The second step is to generate the intake duct reshaping area: In the plane of the front part of the nacelle, the intersection of the filled curved surface formed by the closed leading edge lip and the upper surface of the new fuselage is taken as the bottom edge line of the air intake. The bottom edge line of the air intake is connected to the upstream endpoint of the upper fixed line of the air intake through a spline, which is the side edge line of the air intake. The front control line of the air intake is composed of the side line and the bottom line of the air intake. The multi-section curved surface generated by the front control line, the rear control line and the upper fixed line of the air intake is the air intake shaping area. The third step is to generate the nacelle shaping area: The plane containing the maximum width of the initial axisymmetric nacelle of the engine is taken as the mid-plane of the nacelle. Within this plane, the intersection of this plane and the fixed line at the upper end of the nacelle is taken as the starting point. A line segment of length L1 is drawn as the nacelle forward taper control line. The angle between the nacelle forward taper control line and the vertical direction is the nacelle taper control angle γz1. L1 is 0.6d~1.0d, and γz1 is -5°~5°. It is defined that the nacelle taper control angle γz1 is positive when it deflects towards the vertical symmetry plane of the fuselage, and negative when it deflects away from it. A spline is drawn from the starting point and the ending point of the nacelle forward taper control line; this spline is the nacelle forward control line; the nacelle forward control line is tangent to the initial axisymmetric nacelle at its starting point. Take the plane where the rear edge of the initial axisymmetric nacelle of the engine is located as the rear plane of the nacelle. In this plane, take the downstream end point of the upper fixed line of the nacelle as the starting point, and take the line segment of length L2 as the rear taper control line of the nacelle. The angle between the rear taper control line of the nacelle and the vertical direction is the nacelle taper control angle γz2, where L2 is 0.6d~1.0d and γz2 is -5°~35°. It is defined that the nacelle taper control angle γz1 is positive when it deflects towards the vertical symmetry plane of the fuselage, and negative when it deflects away from it. A spline is drawn with the starting point and the ending point of the rear taper control line of the nacelle. This spline is the rear control line of the nacelle, and the rear control line of the nacelle is tangent to the initial axisymmetric nacelle at its starting point. The nacelle shaping area is generated using the side line of the air intake, the front control line of the nacelle, and the rear control line of the nacelle as control lines. Step 4: Generate the nacelle rectification zone: Take the vertical plane containing the upstream endpoint of the upper fixed line of the air intake as the planes on both sides of the nacelle. Take the intersection line of this plane with the nacelle shaping area and the intersection line of this plane with the upper surface of the new fuselage. Make a first rounding with a rounding radius r1 between the two intersection lines. In the plane of the middle part of the nacelle, take the intersection line of this plane with the nacelle shaping area and the intersection line of this plane with the upper surface of the new fuselage. Make a second rounding with a rounding radius r2 between the two intersection lines. In the plane of the rear part of the nacelle, take the intersection line of this plane with the nacelle shaping area and the intersection line of this plane with the upper surface of the new fuselage. Make a third rounding with a rounding radius r3 between the two intersection lines. The values ​​of r1, r2, and r3 are all between 0.05d and 0.15d. Taking the lower end of the side line of the air intake as the starting point of the rectification, and starting from this point, the first rounding, the second rounding and the third rounding are used as control lines to generate the nacelle rectification area; Step 3, coordinating the tail nozzle design with the fuselage; The specific process is as follows: The first step is to generate the tail nozzle lift zone: The initial control point of the outer duct is taken as the intersection of the engine's vertical symmetry plane and the initial trailing edge line of the outer duct near the fuselage. The initial trailing edge line of the outer duct is the circular trailing edge line of the outer ring of the initial inner surface of the outer duct. The initial control point of the outer duct is moved vertically downward within the engine's vertical symmetry plane, and the nozzle lift height hex is used to form the inner lift control point, where the nozzle lift height hex is 0.01d to 0.04d. The intersection of the third rounded edges on both sides of the nacelle and the upper surface of the new fuselage is taken as the two lift control points. The lower control line of the lift zone is generated using the inner lift control point and the two lift control points as control points. The lower lift control line is tangent to the third rounded edges. The initial trailing edge line of the outer duct is divided by the intersection of the third rounded edges on both sides of the nacelle and the initial trailing edge line of the outer duct. The lower half of the division is taken as the upper lift control line. The nozzle lift zone is generated using the lower lift control line, the upper lift control line, and the third rounded edges on both sides of the nacelle as edge lines. The second step is to generate the tail nozzle expansion zone: The intersection of the engine's vertical symmetry plane and the trailing edge line of the fuselage's upper surface is taken as the initial control point of the trailing edge. Within the engine's vertical symmetry plane, the initial control point of the trailing edge is moved down by an extension depth dex to form an inner control point of the extension. The extension depth dex should be less than the thickness of the rear edge surface of the fuselage. On the trailing edge line of the upper surface of the fuselage, the initial control point of the trailing edge is moved inward and outward by 0.5 times the extension width wex to form two control points of the extension. Wex is taken as 0~d. Using the inner control point of the extension and the two control points of the extension, a rear control line of the extension area is generated within the rear edge surface of the fuselage. The rear control line of the extension area satisfies that both ends are tangent to the trailing edge line of the upper surface of the fuselage. Connect the lifting two control points and the extension two control points, and draw an inner curve within the curved surface of the new upper surface of the fuselage as the control lines on both sides of the extension area. The third step is to generate the new rear edge surface of the machine: The extended area rear control line is an inner curve of the rear edge surface of the fuselage, and the two ends of the extended area rear control line are located on the rear edge line of the upper surface of the fuselage. Therefore, the extended area rear control line divides the rear edge surface of the fuselage into upper and lower parts, and the lower half is taken to form a new rear edge surface of the fuselage. Through steps 1 to 3, complete the integrated aerodynamic shape design of the upper surface of the dorsal engine wing-body blended layout; The full fuselage is generated by symmetrically processing the semi-molded fuselage; With this, the integrated flight-engine design with a dorsal engine and blended wing-body layout was completed.

2. The integrated flight-engine design method for a dorsal-mounted engine with blended wing-body layout as described in claim 1, characterized in that, The clean configuration of the blended wing-body layout mentioned in step 1 refers to an aerodynamic shape consisting only of the fuselage and wings, excluding the engine, tail, control surfaces, and landing gear components; the turbofan engine nacelle refers to a simplified shape that simplifies the engine fan and turbine into intake and exhaust surfaces; the trailing edge line of the lower surface of the clean configuration is parallel to the trailing edge line of the upper surface of the fuselage; the trailing edge line of the upper surface of the fuselage refers to the curve formed by connecting the trailing edge points of the upper surface of the airfoil section, and the trailing edge line of the lower surface is defined in the same way; the surface between the trailing edge lines of the upper and lower surfaces is the trailing edge surface; the trailing edge surface should have a thickness of 0.003c~0.015c, where c is the mean aerodynamic chord length.

3. The flight-engine integrated design method for a dorsal-mounted engine with blended wing-body layout as described in claim 1, characterized in that, In step 1, the spanwise position of the engine core reference point is 0.08c~0.18c from the vertical symmetry plane of the fuselage, the chordwise position is 1.10c~1.50c from the top of the fuselage head, and the vertical position is 0.35d~0.60d from the upper surface of the fuselage, where d is the diameter of the engine fan.

4. The integrated flight-engine design method for a dorsal-mounted engine with blended wing-body layout as described in claim 1, characterized in that, In the design of the rear fuselage upper surface based on the engine position and mounting angle: The new shaping segment of the first control line on the upper surface should satisfy: First, the tangent direction of the control line at the lower edge of the fan intake surface is the same as the direction of the engine thrust line; second, it is tangent to the fixed segment of the first control line on the upper surface at the first dividing point; third, the rear end of the spline is tangent to the modified segment of the initial control line. The new shaping segment of the second control line on the upper surface should meet the following requirements. ： First, the tangent direction of the control line at the first control point on the upper surface is the same as the projection direction of the engine thrust line onto the vertical symmetry plane of the fuselage; second, it is tangent to the fixed segment of the second control line on the upper surface at the second dividing point; third, the rear end of the spline is tangent to the shaping segment of the initial control line. The new modified segment of the third control line on the upper surface should meet the following requirements: First, the tangent direction of the control line at the second control point on the upper surface is the same as the projection direction of the engine thrust line on the vertical plane outside the fuselage; second, it is tangent to the fixed segment of the third control line on the upper surface at the third dividing point on the upper surface; third, the rear end of the spline is tangent to the shaping segment of the initial control line.

5. The integrated flight-engine design method for a dorsal-mounted engine with blended wing-body layout as described in claim 1, characterized in that, The initial axisymmetric nacelle of the engine refers to the outer casing and air intake portion of the turbofan engine's power nacelle. The outer casing portion and the air intake portion are bounded by the leading edge lip circle, and the initial axisymmetric nacelle of the engine is axisymmetric about the engine thrust line.

6. The integrated flight-engine design method for a dorsal-mounted engine with blended wing-body layout as described in claim 1, characterized in that, In the design of integrating the nacelle / air intake with the fuselage: The side line of the air intake satisfies the following conditions: first, it is tangent to the leading edge lip circle; second, it is tangent to the bottom line of the air intake in the plane of the front part of the nacelle. The air intake shaping area satisfies the following conditions: first, it is tangent to the nacelle fixing area; second, it is tangent to the upper surface of the new fuselage.

7. The flight-engine integrated design method for a dorsal-mounted engine with blended wing-body layout as described in claim 1, characterized in that, In the design of integrating the nacelle / air intake with the fuselage: The nacelle shaping area satisfies the following conditions: first, it is tangent to the air intake shaping area; second, it is tangent to the nacelle fixing area. The nacelle rectification area satisfies the following conditions: first, it is tangent to the upper surface of the new fuselage; second, it is tangent to the nacelle shaping area.

8. The integrated flight-engine design method for a dorsal-mounted engine with blended wing-body layout as described in claim 1, characterized in that, The values ​​of r1, r2, and r3 are 0.05d to 0.15d.

9. The integrated flight-engine design method for a dorsal-mounted engine with blended wing-body layout as described in claim 1, characterized in that, In the coordinated design of the tail nozzle and fuselage: The control lines on both sides of the extended area satisfy the following conditions: First, the upstream line is tangent to the intersection of the nacelle rectification area and the upper surface of the new fuselage; Second, the downstream line is perpendicular to the trailing edge line of the upper surface of the fuselage; The tail nozzle extended area is generated with the lower control line of the lift area, the rear control line of the extended area, and the control lines on both sides of the extended area as the side lines; The tail nozzle extended area is tangent to the upper surface of the new fuselage.

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