CN114313253A - Aerodynamic layout and design method of high lift-drag ratio air-breathing hypersonic aircraft - Google Patents
Aerodynamic layout and design method of high lift-drag ratio air-breathing hypersonic aircraft Download PDFInfo
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Abstract
The invention discloses a pneumatic layout and a design method of a high lift-drag ratio air-breathing hypersonic aircraft, belonging to the field of pneumatic layout design of aircrafts and comprising an aircraft body/inner runner pneumatic layout; the engine body/inner runner pneumatic layout is respectively provided with a precursor pre-compression surface, a first-stage outer compression surface, a second-stage outer compression surface, a third-stage outer compression surface, a stamping runner, a first tail nozzle and a second tail nozzle from front to back, the air inlet splitter plate is positioned inside the second-stage outer compression surface, and the turbine runner is connected above the stamping runner in parallel; the invention has high-efficiency high-speed cruising flight capability, meets the high lift requirement of the low-speed section of the aircraft, simultaneously improves the flow separation problem of the high-speed wing in the low-speed takeoff state, and can ensure the wide-speed-range flight performance of the aircraft.
Description
Technical Field
The invention relates to the field of aircraft aerodynamic layout design, in particular to an aerodynamic layout and a design method of a high-lift-drag-ratio air-breathing hypersonic aircraft.
Background
Hypersonic aircraft are leading-edge sites in the current aerospace field. The hypersonic aircraft is one of typical applications of the hypersonic aircraft, the hypersonic aircraft needs to take off and land from the ground horizontally and cruise for a long time under high-altitude and high-speed conditions, and has large flying airspace, wide speed range and high pneumatic layout design difficulty. Furthermore, aircraft are exposed to complex force/heat environments during flight, and the technical challenges of aerodynamic layout design of such aircraft are further exacerbated by the unclear knowledge of force-heat effects.
The air-breathing combined power is one of the preferred power systems of the hypersonic aircraft. The turbine-based combined cycle engine (TBCC) is suitable for a hypersonic aircraft remotely flying in the atmosphere in the whole course due to low speed ratio rush. The engine net thrust is small and the matching of the thrust resistance of the aircraft is difficult under the hypersonic speed condition, so that the aircraft generally adopts an integrated layout form of the engine body/the propulsion, and the layout design of the aircraft needs to comprehensively consider the coupling matching problem of the air suction type combined power system. The chinese patent CN106321283A, "integrated layout method for aerodynamic propulsion of hypersonic aircraft based on combined power", proposes an integrated layout design method from the perspective of improving TBCC performance, but does not consider the comprehensive thermal problem of the aircraft. Chinese patent CN107368661A, namely a coupling calculation method for the thermoaeroelastic characteristics of a hypersonic aerocraft, provides a coupling calculation method suitable for the thermoaeroelastic problem of the hypersonic aerocraft, but application research on the whole aerocraft is not carried out based on the method.
The requirement for technical development is the primary index of the aerodynamic layout design of the hypersonic aircraft, and in order to ensure that the aerodynamic layout has engineering application value, the aircraft needs to comprehensively consider and consider the aerodynamic performance under high and low speed conditions, and simultaneously considers the comprehensive force/heat problem in the whole flight process. The above-mentioned existing design requirements present a significant challenge to the aerodynamic layout design of this type of aircraft.
Disclosure of Invention
The invention aims to overcome the defects of the prior art, provides a pneumatic layout and a design method of a high-lift-drag-ratio air-breathing hypersonic aircraft, has high-efficiency high-speed cruising flight capability, meets the high-lift requirement of the low-speed section of the aircraft, simultaneously improves the flow separation problem of high-speed wings in a low-speed takeoff state, and can ensure the wide-speed-range flight performance of the aircraft.
The purpose of the invention is realized by the following scheme:
a high lift-drag ratio air-breathing hypersonic aircraft aerodynamic layout comprises an aircraft body/inner runner aerodynamic layout; the engine body/inner runner pneumatic layout is respectively provided with a precursor pre-compression surface, a first-stage outer compression surface, a second-stage outer compression surface, a third-stage outer compression surface, a stamping runner, a first tail nozzle and a second tail nozzle from front to back, the air inlet splitter plate is positioned inside the second-stage outer compression surface, and the turbine runner is connected above the stamping runner in parallel.
Further, the method comprises a windward pneumatic layout; the most front end of the airplane body in the windward side pneumatic layout is provided with a head rounding, and the rear end of the airplane body is provided with a precursor pre-compression surface, a first-stage external compression surface, a second-stage external compression surface, a third-stage external compression surface, a TBCC engine shell and a tail nozzle in sequence, an air inlet splitter plate is positioned inside the second-stage external compression surface, and wings are positioned on two sides of the rear-end airplane body.
Further, a leeward side pneumatic layout is included; in the pneumatic layout of the leeward side, the shape of a longitudinal section is controlled by a leeward side body ridge line, the shape of a transverse section is controlled by a cross section control curve of the body, the two sides of the body are an edge wing and a wing which are connected in front and back, and a connecting section is in smooth transition; the strake wing, the wing transition section front edge and the main wing front edge respectively correspond to the strake wing, the fairing transition area and the front edge rounding of the wing, and the combined wing of the strake wing and the wing is positioned on two sides of the fuselage and is in fairing connection with the fuselage through the wing body fusion curved surface.
Further, including a precursor/inlet integrated aerodynamic layout; in the precursor/inlet integrated aerodynamic layout: and the airflow passes through the integrated aerodynamic layout of the precursor/air inlet from front to back, and the precursor pre-compression shock wave, the precursor expansion wave system and the air inlet external compression surface are sequentially generated.
Further, the integrated pneumatic layout of the rear body/tail nozzle is included; in the integrated pneumatic layout of the rear body/tail nozzle, the outlet air of the engine expands after flowing through the expansion surface of the tail nozzle to generate thrust, an outer nozzle plume shear layer is formed at the lower end of the oblique cutting nozzle, the shear layer interacts with the free incoming flow in front to generate compression shock waves, and the expansion surface of the tail nozzle and the outer nozzle plume shear layer are respectively equivalent to an equivalent expansion surface of the tail nozzle and an equivalent plume shear layer of the outer nozzle.
A design method based on the aerodynamic layout of the high-lift-drag-ratio air-breathing hypersonic aircraft comprises the following steps:
s1, arranging a flow dividing guide plate at a secondary compression surface of an air inlet channel layout of three-stage external compression to form a parallel double-channel TBCC engine layout of turbine runner secondary compression and punching runner tertiary compression, and performing smooth curved surface design and optimization on a precursor/air inlet channel integrated precursor pre-compression surface part to ensure the homogenization of incoming flow;
s2, based on the quadratic curve
Constructing a parameterized tip precursor pre-compression curved surface, wherein an outer air inlet channel compression surface is designed behind a precursor pre-compression shock wave and a precursor expansion wave system; wherein x is the parameter variable of the longitudinal position of the machine body, y is the parameter variable of the position in the height direction, and a, b, c, d and e are the systems of the corresponding terms of the equationCounting;
s3, expanding the tail nozzle
Is equivalent to a line segment
Definition of the equivalent expansion angle
(ii) a Shearing the flow field of the free flow surface of the outer nozzle
Equivalent to the expansion surface of the outer nozzle
Let the expansion angle be
(ii) a Based on formula (1), by engine thrust
Equivalent pressure of thrust reverser nozzle
；
Wherein W is the width of the body;
based on the shock wave relation (2), the wave-rear pressure of the compression shock wave (30) generated by the shear layer is equal to the equivalent pressure
Then calculating the expansion angle by combining the formula (1)
The size of (d);
wherein the content of the first and second substances,
in order to be the angle of the shock wave,
the ratio of specific heat is shown as the ratio,
is Mach number;
setting the flying center of mass as
The pitching moment generated by the fore body and the rear body/tail nozzleM zComprises the following steps:
wherein the content of the first and second substances,
is the curve between the points O and a,
is a curve
The distance of any point in the axial direction from the aircraft head,
in correspondence with
The pressure at the point of location is,
is the distance of the equivalent center point X in the axial direction from the aircraft head,
is the distance of point X from the aircraft head in a normal direction perpendicular to the axial direction;
ensure the pitching moment to be in the designated rangeM z,min，M z,max]Inner, calculate
And
in combination with the equivalent expansion angle
And
to obtain the length of the inner spray pipe
Length of outer nozzle
The lengths are respectively as follows:
s4, integrally fusing the front body side edge, the strake wing and the main wing front edge, carrying out parametric design on the cross section of the wing body to be in a sharp side edge shape by using a CST curve through the same fillet design, enabling the front body side edge to be equivalent to a convex front edge strake, and constructing a main lifting surface layout of a curved edge/straight edge two-stage strake + large sweepback trapezoidal wing through smooth transition, wherein the wing type adopts a double-arc wing with the front and back fillets;
and S5, carrying out parameterization on the whole body by utilizing a CST curve and a spline curve, and carrying out transonic and hypersonic area law design by combining the engine size and the fuel volume parameter.
The beneficial effects of the invention include:
(1) according to the aerodynamic layout of the hypersonic aircraft, through the high ridge back curve configuration and the longitudinal supersonic velocity area law design, the front body/air inlet channel design and the rear body/tail nozzle design of equivalent expansion of the wave system matching design are combined, on the premise that the volume and the pitching stability of the aircraft are guaranteed, the lift-drag ratio of the hypersonic aircraft (about 6 Mach) can reach 4.8, and the hypersonic aircraft has high-efficiency high-speed cruising flight capability.
(2) According to the embodiment of the invention, by combining the combined design of the multi-section strake and the large-sweepback trapezoidal double-arc wing, on the premise of ensuring the aerodynamic performance of the hypersonic section, the stall attack angle of the aircraft exceeds 25 degrees by utilizing the attachment and energy compensation of the shedding vortex of the multi-section strake wing on the main wing surface, so that the high-lift requirement of the low-speed section of the aircraft is met.
(3) The embodiment of the invention develops a pneumatic layout of a high-lift-drag-ratio air-breathing hypersonic aircraft, and solves the problems in the background technology. The aircraft under the aerodynamic layout has the capability of flying at the speed range of M0-6 + efficiently. On the premise of meeting the overall requirement and the comprehensive force heat effect, the requirement of high lift-drag ratio (not lower than 4.5) under the condition of hypersonic speed is met through the area law design and the integrated fusion of a front body/air inlet channel and a rear body/tail nozzle on the premise of ensuring the operation/stability characteristics of the aircraft; in addition, through the multi-section integration design of the fuselage/the wings, the problem of flow separation of the high-speed wings in a low-speed takeoff state is solved, and the wide-speed-range flight performance of the airplane is ensured.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to these drawings without creative efforts.
FIG. 1 is a typical flight trajectory diagram for a hypersonic aircraft;
FIG. 2 is a layout view of the fuselage/internal flow passage of a hypersonic aircraft;
FIG. 3 is a layout view of the windward side of a hypersonic aircraft;
FIG. 4 is a layout view of the leeward side of a hypersonic aircraft;
FIG. 5 is a schematic view of a precursor/inlet integration;
FIG. 6 is a schematic view of the aft body/jet nozzle integration;
in the figure, 1-ascent section trajectory, 2-cruise section trajectory, 3-return section trajectory, 4-turbine flow channel, 5-stamping flow channel, 6-primary external compression surface, 7-secondary external compression surface, 8-tertiary external compression surface, 9-air inlet splitter plate, 10-first tail nozzle, 11-second tail nozzle, 12-precursor pre-compression surface, 13-TBCC engine shell, 14-wing, 15-tail wing, 16-strake wing, 17-wing transition section leading edge, 18-main wing leading edge, 19-leeward side fuselage ridge line, 20-fuselage section, 21-head radius, 22-wing body fusion curved surface, 23-precursor pre-compression shock wave, 24-precursor expansion wave system, 25-air inlet external compression surface, 26-expansion surface of the tail nozzle, 27-equivalent expansion surface of the tail nozzle, 28-plume shear layer of the outer nozzle, 29-equivalent plume shear layer of the outer nozzle and 30-compression shock wave.
Detailed Description
All features disclosed in all embodiments in this specification, or all methods or process steps implicitly disclosed, may be combined and/or expanded, or substituted, in any way, except for mutually exclusive features and/or steps.
The invention at least solves the following technical problems:
the embodiment of the invention develops a pneumatic layout of a high-lift-drag-ratio air-breathing hypersonic aircraft, so that the aircraft has the capability of flying at high efficiency in an M0-6 + speed range. On the premise of meeting the overall requirement and the comprehensive force heat effect, the requirement of high lift-drag ratio (not lower than 4.5) under the condition of hypersonic speed is met through the area law design and the integrated fusion of the front body/air inlet channel and the rear body/tail nozzle on the premise of ensuring the operation/stability characteristics of the aircraft. In addition, the embodiment of the invention provides a fuselage/wing multi-section fusion integrated design scheme, which can improve the flow separation problem of high-speed wings in a low-speed takeoff state and ensure the wide-speed-domain flight performance of an airplane.
In the embodiment of the invention, the tip precursor micro-convex precompression surface parameterized by a quadratic curve is also taken from the viewpoints of overall demand and comprehensive force/thermal effect
The generated shock wave and the expansion wave system are optimized by using the shape parameters, the customized distribution of the precursor wave system and the pressure is realized, the relative position relation between the precursor pre-compression wave system and the compression shock wave outside the air inlet channel is further determined, and the overflow lift force and the air inflow are reasonably distributed and used as the design basis for determining the relative position of the air inlet channel on the body.
Furthermore, the embodiment of the invention also develops an equivalent expansion design method of the single-wall jet nozzle, which can guide the expansion ratio of the jet nozzle
Length of inner nozzle
And outer nozzle length
The pitch trim problem of the long fuselage is solved by designing and matching the pressure distribution of the pre-compression surface of the forebody.
Furthermore, the embodiment of the invention also considers the problems of high flying lift requirement, low resistance requirement at high speed and internal volume ratio, and realizes the reattachment of the strake wing leading edge vortex on the leeward side of the large-sweepback main wing surface by a multi-section fusion type strake design means on the premise of ensuring the comprehensive force heat effect, thereby improving the separation and the stall of the high-speed wing in the state of low speed and large attack angle. In a specific application, the specific technical solution of the embodiment of the present invention may perform the following steps:
the method comprises the following steps: according to flight range R and payload mass
Defining cruise sections on the basis of mission requirementsLift to drag ratio
Thrust of engine
And fuel quality
Simultaneously, combines the trajectory analysis of the ascending section to determine the thrust requirement of the ascending section
And fuel consumption
Calculating to obtain the volume of the aircraft
And take-off lift
。
Step two: according to the thrust requirement of an engine, the air intake amount and the length of the engine are preliminarily determined, parameters such as the width of an air inlet channel, the length of a precursor and the air inlet height are obtained through analysis, based on an equi-shock wave intensity theory, the compression angle of an outer compression surface of the air inlet channel is designed, the air inlet channel layout of three-stage outer compression is obtained, a diversion guide plate is designed at a second-stage compression surface, the parallel double-channel TBCC engine layout of two-stage compression of a turbine runner and three-stage compression of a ram runner is formed, smooth curved surface design and optimization are carried out on the precursor pre-compression surface part of the integration of the precursor/the air inlet channel, and the homogenization of incoming flow is ensured.
Step three: based on quadratic curves
Constructing a parameterized pre-compression curved surface of the tip precursor, and analyzing the pressure distribution of the pre-compression surface of the precursor by combining with a CFD (computational fluid dynamics) means
And shock/expansion wave system shape, and the integral layout size (including the length of the precursor, the overflow height and the like) of the air inlet and the aircraft precursor is determined by taking the wave system envelope of the precursor (namely the precursor pre-compression shock wave 23 and the rear of the precursor expansion wave system 24 in the figure 5) as a design basis and combining the boundary layer thickness development characteristics, wherein the outer-inlet compression shock wave (namely the outer-inlet compression surface 25 in the figure 5) must be positioned at the front body.
Step four: expanding surface of tail nozzle
Is equivalent to a line segment
Definition of the equivalent expansion angle
. Shearing the flow field of the free flow surface of the outer nozzle
Equivalent to the expansion surface of the outer nozzle
Assuming an expansion angle of
. Based on formula (1), by engine thrust
Equivalent pressure of thrust reverser nozzle
。
Based on the shock wave relation (2), the wave of the shock wave 30 is formedThe back pressure being equal to the equivalent pressure
Then calculating the expansion angle by combining the formula (1)
The size of (2).
Assuming a flight centroid position of
The pitching moment generated by the front body and the rear body/tail nozzle is:
wherein the content of the first and second substances,
is the curve between the points O and a,
is a curve
The distance of any point in the axial direction from the aircraft head,
in correspondence with
The pressure at the point of location is,
is the distance of the equivalent center point X in the axial direction from the aircraft head,
is the distance of point X from the aircraft head in a normal direction perpendicular to the axial direction;
ensure the pitching moment to be in the designated rangeM z,min，M z,max]Inner, calculate
And
in combination with the equivalent expansion angle
And
to obtain the length of the inner spray pipe
Length of outer nozzle
The lengths are respectively as follows:
step five: the wing area is determined by combining the takeoff lift requirement, the high-speed and low-speed lift drag characteristic requirements are considered, the front body side edge, the strake wing front edge and the wing front edge are integrated and fused, the front body side edge, the strake wing front edge and the wing front edge are designed to be sharp side edge shapes in a parameterization mode through the same fillet design, the front body side edge can be equivalent to a convex front edge strake, the main lifting surface layout of a curved edge/straight edge two-stage strake + large sweepback trapezoidal wing is constructed through smooth transition, and the wing type adopts a double-arc wing with the front radius and the back radius.
Step six: the whole body is parameterized by using a CST curve and a spline curve, transonic and hypersonic area law design is carried out by combining the size of an engine and fuel volume parameters, and the reasonable distribution of the cross-sectional area of the body along the axial direction is ensured by parameter design and optimization.
Step seven: the area of the tail 15, the layout and its distribution in the fuselage are defined on the basis of the requirements of the manoeuvre/stability characteristics.
Step eight: and (4) determining the aerodynamic thermal environment of the aircraft based on the flight trajectory, and rounding the head of the aircraft.
In other embodiments of the present invention, as shown in fig. 1, the reference flight trajectory of the aircraft is determined by mission requirements, and the trajectory of the hypersonic aircraft may be divided into an ascent trajectory 1, a cruise trajectory 2, and a return trajectory 3, where the thrust requirement of the full-speed region of the aircraft is determined by the ascent trajectory 1, the lift-drag ratio requirement of the aircraft is determined by the cruise trajectory 2, and the fuel consumption of the aircraft in the whole process is obtained by combining the fuel consumption of the ascent trajectory 1 and the return trajectory 3, and the takeoff mass of the aircraft is calculated by referring to the structural mass ratio (about 40%) of a typical hypersonic aircraft, so as to determine the takeoff lift requirement of the aircraft.
As shown in fig. 2, the intake air amount of the intake passage is calculated based on the thrust demand of the uptake trajectory 1
Further, the intake area of the intake duct is specified
According to the width constraint of the machine body, the width-height ratio of the engine flow passage is reasonably distributed to obtain the air inlet height of the air inlet channel
. The diameter of the turbine runner 4 is determined according to the turbine thrust requirement of the aircraft in the transonic speed and turbine/stamping mode conversion section by referring to the performance of the off-the-shelf turbine
While at the same time, according to the fuel flow at cruising speed
And the burning rate
And inlet exit velocity
The engine length of the ram runner is determined by equation (5):
width of fuselage
And air intake height
Under the constraint of (2), constructing a precursor/inlet integrated layout, wherein the precursor pre-compression angle
And (3) calculating by using a diagonal shock angle formula (2).
As shown in fig. 3, the pre-compression angle and the length of the precursor are confirmed on the principle that the oblique shock wave generated by the precursor must be outside the inlet lip. Meanwhile, Mach number, pressure and the like of the precursor after pre-compression shock wave are obtained through conversion of an oblique shock wave relational expression, and the external compression configuration parameters are calculated by taking the Mach number, the pressure and the like as a reference and combining an isoshock wave strength theoretical formula (6).
M6 is used as a design Mach number to construct a three-stage external compression configuration, compression angles are respectively 4.526 degrees, 4.968 degrees and 5.468 degrees, corresponding shock wave angles are respectively 14.7 degrees, 16.23 degrees and 18 degrees, three external compression shock waves are converged at a lip opening position, and the size envelope and the precursor/air inlet channel integrated layout of the engine are preliminarily determined.
Designing the layout of the afterbody/tail nozzle based on the equivalent expansion methods (1), (2), (3) and (4), wherein the length of the inner nozzle of the tail nozzle is 1200mm, the length of the outer nozzle is 2300mm, and the expansion ratio of the nozzle is 5.4.
As shown in fig. 4, the longitudinal profile of the fuselage is parametrized based on spline lines, while the cross-sectional shape of the principal position of the fuselage is parametrized using the CST curve of equation (7).
Definition of fuel volume considering flight trajectory
And aircraft takeoff mass
Reference to wing-mounted main stream fighter plane
Data to determine reasonably the initial wing area
In order to meet the pneumatic requirements of lift increase and high-speed drag reduction, the layout of the strake wing and the large-sweepback trapezoidal wing is adopted, the sweepback angles are respectively 80 degrees and 55 degrees, the wing section adopts a double-arc wing, and the radius of the front edge of the wing is 35mm and the radius of the rear edge is 25mm in consideration of the heat protection requirement. And combining the parameterized fuselage and the initial wing, and performing fairing treatment at the joint of the wing body.
Example 1
A high lift-drag ratio air-breathing hypersonic aircraft aerodynamic layout comprises an aircraft body/inner runner aerodynamic layout; the engine body/inner runner pneumatic layout is respectively provided with a precursor pre-compression surface 12, a first-stage outer compression surface 6, a second-stage outer compression surface 7, a third-stage outer compression surface 8, a stamping runner 5, a first tail nozzle 10 and a second tail nozzle 11 from front to back, an air inlet splitter plate 9 is positioned inside the second-stage outer compression surface 7, and a turbine runner 4 is connected above the stamping runner 5 in parallel.
Example 2
On the basis of the embodiment 1, the pneumatic layout on the windward side is included; in the windward side pneumatic layout, the foremost end of the fuselage is a head rounding 21, and the rear part of the fuselage is a precursor pre-compression surface 12, a first-stage external compression surface 6, a second-stage external compression surface 7, a third-stage external compression surface 8, a TBCC engine shell 13 and a tail nozzle 11 in sequence, an air inlet splitter plate 9 is positioned inside the second-stage external compression surface 7, and wings 14 are positioned on two sides of the rear-end fuselage.
Example 3
On the basis of the embodiment 1 or 2, the pneumatic layout of the leeward side is included; in the leeward side pneumatic layout, the shape of a longitudinal section is controlled by a leeward side fuselage ridge line 19, the shape of a transverse section is controlled by a fuselage cross section control curve 20, the two sides of the fuselage are an edge wing 16 and a wing 14 which are connected in a front-back manner, and the connecting sections are in smooth transition; the strake wing 16, the wing transition section leading edge 17 and the main wing leading edge 18 respectively correspond to the strake wing 16, the fairing transition area and the leading edge rounding of the wing 14, and the combined wing of the strake wing 16 and the wing 14 is positioned on two sides of the fuselage and is in fairing connection with the fuselage through the wing body fusion curved surface 22.
Example 4
On the basis of embodiment 3, the integrated aerodynamic layout of the precursor/air inlet channel is included; in the precursor/inlet integrated aerodynamic layout: the gas flow from front to back through the precursor/inlet integrated aerodynamic layout will in turn generate a precursor pre-compression shock wave 23, a precursor expansion wave train 24 and an outer inlet compression surface 25.
Example 5
On the basis of the embodiment 4, the integrated aerodynamic layout of the rear body/tail nozzle is included; in the integrated pneumatic layout of the afterbody/tail nozzle, the outlet gas of the engine expands after flowing through the expansion surface 26 of the tail nozzle to generate thrust, and meanwhile, an outer nozzle plume shear layer 28 is formed at the lower end of the oblique cutting nozzle, the shear layer interacts with the free incoming flow in front to generate compression shock waves 30, and the expansion surface 26 of the tail nozzle and the outer nozzle plume shear layer 28 are respectively equivalent to an equivalent expansion surface 27 of the tail nozzle and an equivalent plume shear layer 29 of the outer nozzle.
Example 6
On the basis of any embodiment 1-5, the design method based on the aerodynamic layout of the high-lift-drag-ratio air-breathing hypersonic aircraft comprises the following steps:
s1, arranging a flow dividing guide plate at a secondary compression surface of an air inlet channel layout of three-stage external compression to form a parallel double-channel TBCC engine layout of turbine runner secondary compression and punching runner tertiary compression, and performing smooth curved surface design and optimization on a precursor/air inlet channel integrated precursor pre-compression surface part to ensure the homogenization of incoming flow;
s2, based on the quadratic curve
Constructing a parameterized tip precursor pre-compression curved surface, and designing an outer air inlet channel compression surface 25 behind a precursor pre-compression shock wave 23 and a precursor expansion wave system 24; wherein x is a parameter variable of the longitudinal position of the machine body, y is a parameter variable of the position in the height direction, and a, b, c, d and e are coefficients of all items corresponding to an equation;
s3, expanding the tail nozzle
Is equivalent to a line segment
Definition of the equivalent expansion angle
(ii) a Shearing the flow field of the free flow surface of the outer nozzle
Equivalent to the expansion surface of the outer nozzle
Let the expansion angle be
(ii) a Based on formula (1), by engine thrust
Equivalent pressure of thrust reverser nozzle
；
Wherein W is the width of the body;
based on the shock wave relation (2), the wave-backward pressure of the compression shock wave 30 generated by the shear layer is equal to the equivalent pressure
Then calculating the expansion angle by combining the formula (1)
The size of (d);
wherein the content of the first and second substances,
in order to be the angle of the shock wave,
the ratio of specific heat is shown as the ratio,
is Mach number;
setting the flying center of mass as
The pitching moment generated by the fore body and the rear body/tail nozzleM zComprises the following steps:
wherein，
Is the curve between the points O and a,
is a curve
The distance of any point in the axial direction from the aircraft head,
in correspondence with
The pressure at the point of location is,
is the distance of the equivalent center point X in the axial direction from the aircraft head,
is the distance of point X from the aircraft head in a normal direction perpendicular to the axial direction;
ensure the pitching moment to be in the designated rangeM z,min，M z,max]Inner, calculate
And
in combination with the equivalent expansion angle
And
to obtain the length of the inner spray pipe
Exterior and interiorLength of nozzle
The lengths are respectively as follows:
s4, integrally fusing the front body side edge, the strake wing 16 and the main wing front edge 18, carrying out the same fillet design, parametrization designing the fuselage section 20 into a sharp side edge shape by utilizing a CST curve, enabling the front body side edge to be equivalent to a convex front edge strake, and constructing the main lifting surface layout of a curved edge/straight edge two-stage strake + large sweepback trapezoidal wing through smooth transition, wherein the wing type adopts a double-arc wing with the front and back fillets;
and S5, carrying out parameterization on the whole body by utilizing a CST curve and a spline curve, and carrying out transonic and hypersonic area law design by combining the engine size and the fuel volume parameter.
The parts not involved in the present invention are the same as or can be implemented using the prior art.
The above-described embodiment is only one embodiment of the present invention, and it will be apparent to those skilled in the art that various modifications and variations can be easily made based on the application and principle of the present invention disclosed in the present application, and the present invention is not limited to the method described in the above-described embodiment of the present invention, so that the above-described embodiment is only preferred, and not restrictive.
Other embodiments than the above examples may be devised by those skilled in the art based on the foregoing disclosure, or by adapting and using knowledge or techniques of the relevant art, and features of various embodiments may be interchanged or substituted and such modifications and variations that may be made by those skilled in the art without departing from the spirit and scope of the present invention are intended to be within the scope of the following claims.
Claims (6)
1. A high lift-drag ratio air-breathing hypersonic aircraft aerodynamic layout is characterized by comprising an aircraft body/inner runner aerodynamic layout; the engine body/inner runner pneumatic layout is respectively provided with a precursor pre-compression surface (12), a first-stage outer compression surface (6), a second-stage outer compression surface (7), a third-stage outer compression surface (8), a stamping runner (5), a first tail nozzle (10) and a second tail nozzle (11) from front to back, an air inlet channel flow distribution plate (9) is located inside the second-stage outer compression surface (7), and a turbine runner (4) is connected above the stamping runner (5) in parallel.
2. The aerodynamic configuration of a high lift-to-drag ratio air-breathing hypersonic aircraft as claimed in claim 1, comprising a windward aerodynamic configuration; the most front end of the airplane body in the windward side pneumatic layout is a head rounding (21), a precursor pre-compression surface (12), a first-stage outer compression surface (6), a second-stage outer compression surface (7), a third-stage outer compression surface (8), a TBCC engine shell (13) and a tail nozzle (11) are sequentially arranged behind the head rounding, an air inlet splitter plate (9) is located inside the second-stage outer compression surface (7), and wings (14) are located on two sides of the rear-end airplane body.
3. The aerodynamic configuration of a high lift-to-drag ratio air-breathing hypersonic aircraft as claimed in any one of claims 1 or 2, comprising a leeward aerodynamic configuration; in the leeward side pneumatic layout, the shape of a longitudinal section is controlled by a leeward side body ridge line (19), the shape of a transverse section is controlled by a body cross section control curve (20), an edge wing (16) and a wing (14) are arranged on two sides of the body and are connected in a front-back mode, and a connecting section is in smooth transition; the strake wing (16), the wing transition section front edge (17) and the main wing front edge (18) respectively correspond to the strake wing (16), the fairing transition area and the front edge rounding of the wing (14), and the combined wing of the strake wing (16) and the wing (14) is positioned on two sides of the fuselage and is in fairing connection with the fuselage through the wing body fusion curved surface (22).
4. The aerodynamic configuration of a high lift-to-drag ratio air-breathing hypersonic aircraft as claimed in claim 3, comprising a precursor/inlet integrated aerodynamic configuration; in the precursor/inlet integrated aerodynamic layout: the gas flow passes through the precursor/inlet integrated aerodynamic layout from front to back, and the precursor pre-compression shock wave (23), the precursor expansion wave system (24) and the inlet external compression surface (25) are generated in sequence.
5. The aerodynamic configuration of a high lift-to-drag ratio air-breathing hypersonic aircraft as claimed in claim 4, characterized by comprising an aft body/jet nozzle integrated aerodynamic configuration; in the integrated pneumatic layout of the afterbody/tail nozzle, the outlet gas of the engine expands after flowing through the expansion surface (26) of the tail nozzle to generate thrust, an outer nozzle plume shear layer (28) is formed at the lower end of the oblique cutting nozzle, the shear layer interacts with the front free incoming flow to generate compression shock waves (30), and the expansion surface (26) of the tail nozzle and the plume shear layer (28) of the outer nozzle are respectively equivalent to the equivalent expansion surface (27) of the tail nozzle and the equivalent plume shear layer (29) of the outer nozzle.
6. A design method for aerodynamic layout of high lift-drag ratio air-breathing hypersonic speed airplane based on claim 5 is characterized by comprising the following steps:
s1, arranging a flow dividing guide plate at a secondary compression surface of an air inlet channel layout of three-stage external compression to form a parallel double-channel TBCC engine layout of turbine runner secondary compression and punching runner tertiary compression, and performing smooth curved surface design and optimization on a precursor/air inlet channel integrated precursor pre-compression surface part to ensure the homogenization of incoming flow;
s2, based on the quadratic curve
Constructing a parameterized tip precursor pre-compression curved surface, and designing an outer air inlet channel compression surface (25) behind a precursor pre-compression shock wave (23) and a precursor expansion wave system (24); wherein x is a parameter variable of the longitudinal position of the machine body, y is a parameter variable of the position in the height direction, and a, b, c, d and e are coefficients of all items corresponding to an equation;
s3, expanding the tail nozzle
Is equivalent to a line segment
Definition of the equivalent expansion angle
(ii) a Shearing the flow field of the free flow surface of the outer nozzle
Equivalent to the expansion surface of the outer nozzle
Let the expansion angle be
(ii) a Based on formula (1), by engine thrust
Equivalent pressure of thrust reverser nozzle
；
Wherein W is the width of the body;
based on the shock wave relation (2), the wave-rear pressure of the compression shock wave (30) generated by the shear layer is equal to the equivalent pressure
Then calculating the expansion angle by combining the formula (1)
The size of (d);
wherein the content of the first and second substances,
in order to be the angle of the shock wave,
the ratio of specific heat is shown as the ratio,
is Mach number;
setting the flying center of mass as
The pitching moment generated by the fore body and the rear body/tail nozzleM zComprises the following steps:
wherein the content of the first and second substances,
is the curve between the points O and a,
is a curve
The distance of any point in the axial direction from the aircraft head,
in correspondence with
The pressure at the point of location is,
is the distance of the equivalent center point X in the axial direction from the aircraft head,
is the distance of point X from the aircraft head in a normal direction perpendicular to the axial direction;
ensure the pitching moment to be in the designated rangeM z,min，M z,max]Inner, calculate
And
in combination with the equivalent expansion angle
And
to obtain the length of the inner spray pipe
Length of outer nozzle
The lengths are respectively as follows:
s4, integrally fusing the front body side edge, the strake wing (16) and the main wing front edge (18), designing the cross section (20) of the airplane body into a sharp side edge shape by the aid of the same fillet design and utilizing a CST curve in a parameterization mode, enabling the front body side edge to be equivalent to a convex front edge strake, and constructing a main lifting surface layout of a curved edge/straight edge two-stage strake plus a large sweepback trapezoidal wing through smooth transition, wherein the wing type adopts a double-arc wing with the front and back rounded corners;
and S5, carrying out parameterization on the whole body by utilizing a CST curve and a spline curve, and carrying out transonic and hypersonic area law design by combining the engine size and the fuel volume parameter.
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN114912202A (en) * | 2022-05-24 | 2022-08-16 | 大连理工大学 | Integrated coupling control method for propelling of wide-speed-range air-breathing power aircraft body |
| CN115871913A (en) * | 2023-02-28 | 2023-03-31 | 中国空气动力研究与发展中心计算空气动力研究所 | Integrated aerodynamic layout of high lift-drag ratio air-breathing aircraft and construction method thereof |
| CN115871913B (en) * | 2023-02-28 | 2023-06-30 | 中国空气动力研究与发展中心计算空气动力研究所 | Construction method of integrated pneumatic layout of high lift-drag ratio air suction type aircraft |
| CN115946842A (en) * | 2023-03-10 | 2023-04-11 | 中国空气动力研究与发展中心计算空气动力研究所 | Damping device of aircraft and aircraft |
| CN115946842B (en) * | 2023-03-10 | 2023-05-26 | 中国空气动力研究与发展中心计算空气动力研究所 | Drag reduction device of aircraft and aircraft |
| CN117208194A (en) * | 2023-11-09 | 2023-12-12 | 清华大学 | Wing-hair reconfiguration type variant aircraft |
| CN117208194B (en) * | 2023-11-09 | 2024-01-09 | 清华大学 | Wing-hair reconfiguration type variant aircraft |
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