CN110020500B - Design method of common tail nozzle of single-side expansion four-channel combined engine - Google Patents

Design method of common tail nozzle of single-side expansion four-channel combined engine Download PDF

Info

Publication number
CN110020500B
CN110020500B CN201910317706.3A CN201910317706A CN110020500B CN 110020500 B CN110020500 B CN 110020500B CN 201910317706 A CN201910317706 A CN 201910317706A CN 110020500 B CN110020500 B CN 110020500B
Authority
CN
China
Prior art keywords
engine
channel
throat
sub
combustion ramjet
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
CN201910317706.3A
Other languages
Chinese (zh)
Other versions
CN110020500A (en
Inventor
陈荣钱
柳家齐
黄阳灿
吴林宽
龙华强
尤延铖
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Xiamen University
Original Assignee
Xiamen University
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Xiamen University filed Critical Xiamen University
Priority to CN201910317706.3A priority Critical patent/CN110020500B/en
Publication of CN110020500A publication Critical patent/CN110020500A/en
Application granted granted Critical
Publication of CN110020500B publication Critical patent/CN110020500B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/97—Rocket nozzles
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/15—Vehicle, aircraft or watercraft design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • General Engineering & Computer Science (AREA)
  • Geometry (AREA)
  • Theoretical Computer Science (AREA)
  • General Physics & Mathematics (AREA)
  • Evolutionary Computation (AREA)
  • Mechanical Engineering (AREA)
  • Combustion & Propulsion (AREA)
  • Chemical & Material Sciences (AREA)
  • Computer Hardware Design (AREA)
  • Pure & Applied Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Mathematical Optimization (AREA)
  • Mathematical Analysis (AREA)
  • Computational Mathematics (AREA)
  • Testing Of Engines (AREA)

Abstract

A design method of a single-side expansion four-channel combined engine shared tail nozzle relates to a hypersonic aircraft. Designing a three-dimensional model of the tail nozzle of the unilateral expansion asymmetric sub-combustion ramjet engine; designing tail nozzles of a turbine engine channel and a rocket engine channel; designing a combined configuration of a common tail nozzle of the parallel TBCC engine; and designing an adjusting mechanism. The nozzles of a plurality of engines are integrated to intensively exhaust, the problem of asymmetric thrust of multi-channel tail nozzles of the plurality of engines is solved, the resistance of the parallel TBCC engine is reduced, and the throat area can be adjusted through an adjusting plate so as to meet the thrust requirement of the parallel TBCC engine in a wide speed range. The four channels are integrated on the secondary combustion ramjet tail pipe, so that the structure is more compact, the difficulty of the integrated design of the aircraft and the engine is reduced, and the torque trim integrated design of the aircraft and the engine has great advantages.

Description

Design method of common tail nozzle of single-side expansion four-channel combined engine
Technical Field
The invention relates to a hypersonic aircraft, in particular to a design method of a common tail nozzle of a single-side expansion four-channel Combined engine, which is a parallel Turbine Based Combined Cycle (TBCC) engine tail nozzle.
Background
Hypersonic aircrafts are high-tech items of major development in the 21 st century. Supersonic ramjets are considered the "third power revolution" following propeller and jet propulsion and are considered the best power option for hypersonic aircraft due to their advantages of simple structure, no need to carry oxidants, etc. However, the ramjet can only start and normally work under high mach number, in order to make the hypersonic aircraft have the capability of autonomous horizontal take-off and landing and high mach number cruising, researchers have proposed the design concept of Turbine-Based Combined Cycle (TBCC) (pneumatic design, performance research and preliminary optimization of lisk, parallel TBCC exhaust system [ D ], Nanjing aerospace university, 2009), i.e. a power system combining the power of the Turbine engine and the power of rocket engine, ramjet, etc. The aircraft is powered by the turbine engine during low mach number flight and by the ramjet engine during high mach number flight. However, when the flight mach number is greater than 2.5, the thrust of the turbine engine is sharply reduced due to the limited temperature of the turbine gas, and the ram engine cannot provide the thrust in time at the moment, so that the problem of a thrust gap exists, and the thrust requirement of the transition section is required to be met by providing the thrust by the rocket engine. Turbine-based combined power can be divided into parallel type and series type according to the layout arrangement of each engine. For a parallel TBCC combined engine, if a mode of separately exhausting air from each channel is adopted, the problem that an exhaust system occupies a large space exists, and great challenges are brought to the integrated design of a super aircraft. Therefore, the design of the tail nozzle of the parallel TBCC combined engine is a problem to be solved by intensive research.
Disclosure of Invention
The invention aims to provide a design method of a single-side expansion four-channel combined engine shared tail nozzle, which can meet the thrust requirements of a ramjet engine, a turbine engine and a rocket engine coupled parallel TBCC combined engine in different states.
The invention comprises the following steps:
1) designing a three-dimensional model of the tail nozzle of the unilateral expansion asymmetric sub-combustion ramjet engine;
in step 1), the specific method for designing the three-dimensional model of the single-side expansion asymmetric sub-combustion ramjet engine tail pipe can be as follows: firstly, designing upper and lower wall molded lines of a unilateral expansion exhaust nozzle based on a short nozzle theory to obtain a two-dimensional exhaust nozzle molded line, then determining the rectangular size of an inlet of the unilateral expansion asymmetric sub-combustion ramjet nozzle of the sub-combustion ramjet engine according to the flow demand of a sub-combustion ramjet engine channel, and laterally stretching the two-dimensional exhaust nozzle molded line to be as wide as the inlet rectangular to obtain a three-dimensional model of the unilateral expansion asymmetric sub-combustion ramjet engine.
2) Designing tail nozzles of a turbine engine channel and a rocket engine channel;
in step 2), the specific method for designing the exhaust nozzle of the turbine engine passage and the rocket engine passage may be as follows: converging two turbine engine tail nozzles, then combining the two turbine engine tail nozzles with a secondary combustion ramjet engine tail nozzle, calculating the area size of the junction and throat of two turbine engine channels according to the flow requirement of the overall design of a parallel TBCC combined engine, combining the design width of the three-dimensional model of the unilateral expansion asymmetric secondary combustion ramjet engine tail nozzle in the step 1), uniformly transferring the two turbine engine circular inlets to a rectangular converging inlet, and then converging the two turbine engine circular inlets to the upper wall surface of the three-dimensional model of the unilateral expansion asymmetric secondary combustion ramjet engine tail nozzle in the step 1) to form the throat of the turbine engine channel; similarly, according to the flow requirement of the overall design of the parallel TBCC combined engine, calculating the throat area of the rocket engine channel, and then converging the throat area to the upper wall surface of the three-dimensional model of the unilateral expansion asymmetric sub-combustion ramjet nozzle in the step 1) to form the throat of the rocket engine channel.
3) Designing a combined configuration of a common tail nozzle of the parallel TBCC engine;
in step 3), the specific method for designing the combined configuration of the common exhaust nozzle of the parallel TBCC engine may be: the sub-combustion ramjet engine channel is positioned at the lowest part, the turbine engine channel is positioned at the middle part, and the rocket engine channel is positioned at the highest part, so that a parallel TBCC engine shared tail nozzle combined structure is formed.
4) And designing an adjusting mechanism.
In step 4), the specific method for designing the adjusting mechanism may be: adjusting mechanisms are respectively arranged at the throat of the turbine engine channel, the throat of the rocket engine channel and the throat of the sub-combustion ramjet engine channel, the adjusting mechanisms adopt a rotating shaft and an adjusting plate, the throat area of the turbine engine channel and the throat area of the rocket engine channel are changed by rotating the adjusting plate around the rotating shaft, and the throat area of the sub-combustion ramjet engine channel is changed by moving the rotating shaft up and down.
The common tail nozzle of the unilateral expansion four-channel combined engine is provided with a rocket engine channel, a turbine engine channel, a sub-combustion ramjet engine channel, a throat adjusting plate of the rocket engine channel, a throat adjusting plate rotating shaft of the rocket engine channel, a throat adjusting plate of the turbine engine channel, a throat adjusting plate rotating shaft of the turbine engine channel, a throat adjusting rotating shaft of the sub-combustion ramjet engine channel and the unilateral expansion sub-combustion ramjet engine tail nozzle; the secondary combustion ramjet passage is positioned at the lowest part of the tail nozzle and connected with the tail nozzle of the unilateral expansion secondary combustion ramjet, the turbine engine passage is positioned in the middle of the tail nozzle, the rocket engine passage is positioned at the uppermost part of the tail nozzle, and the rocket engine passage and the turbine engine passage are both connected with the upper wall surface of the tail nozzle of the unilateral expansion secondary combustion ramjet; the throat adjusting plate of the rocket engine channel is connected with the rocket engine channel, the turbine engine channel and the sub-combustion ramjet engine channel through a throat adjusting plate rotating shaft of the rocket engine channel; the throat adjusting plate of the turbine engine channel is connected with the throat adjusting rotating shafts of the turbine engine channel, the sub-combustion ramjet engine channel and the sub-combustion ramjet engine channel through the throat adjusting plate rotating shaft of the turbine engine channel; the throat adjusting rotating shaft of the sub-combustion ramjet channel is positioned at the lower part of the tail end of the sub-combustion ramjet channel and is connected with the lower wall surface of the tail nozzle of the unilateral expansion sub-combustion ramjet.
Through the connection method, all the parts form a parallel TBCC engine shared exhaust nozzle combined configuration.
The invention has the following outstanding technical effects:
the unilateral expansion four-channel combined tail nozzle integrates the nozzles of a plurality of engines, concentrates exhaust, solves the problem of asymmetric thrust of the multichannel tail nozzles of the plurality of engines, reduces the resistance of a parallel TBCC engine, and can meet the thrust requirement of the parallel TBCC engine in a wide speed range by adjusting the throat area through the adjusting plate. In addition, the four channels are integrated on the secondary combustion ramjet, so that the structure is more compact, the difficulty of the integrated design of the aircraft and the engine is reduced, and the torque balancing method has great advantages in the torque balancing of the integrated design of the aircraft and the engine.
Drawings
FIG. 1 is a schematic diagram of a design characteristic line of a molded line of a tail pipe of a single-side expansion sub-combustion ramjet engine.
FIG. 2 is a schematic design diagram of a single-side expansion secondary combustion ramjet nozzle. In fig. 2, (a) is a side view of the exhaust nozzle of the single-side expansion sub-combustion ramjet engine, and (b) is a side view of the exhaust nozzle of the single-side expansion sub-combustion ramjet engine.
FIG. 3 is a schematic view of a turbine engine passage design. In fig. 3, (a) is a turbine engine passage side view, and (b) is a turbine engine passage axis view.
FIG. 4 is a schematic view of rocket engine tunnel design. In fig. 4, (a) is a side view of the rocket motor passage, and (b) is an axial view of the rocket motor passage.
FIG. 5 is an axial view of a common jet nozzle of a four-channel combined engine with an adjusting mechanism.
FIG. 6 is a schematic diagram of the operation of a common jet nozzle adjusting mechanism of a four-channel combined engine.
In the figure, each label is: 1 denotes a rocket engine passage, 2 denotes a turbine engine passage, 3 denotes a sub-combustion ramjet engine passage, 4 denotes a throat-adjusting plate of the rocket engine passage, 5 denotes a throat-adjusting plate rotating shaft of the rocket engine passage, 6 denotes a throat-adjusting plate of the turbine engine passage, 7 denotes a throat-adjusting plate rotating shaft of the turbine engine passage, 8 denotes a throat-adjusting rotating shaft of the sub-combustion ramjet engine passage, and 9 denotes a single-side expansion sub-combustion ramjet engine tail pipe.
Detailed Description
The following examples will further illustrate the present invention with reference to the accompanying drawings.
The embodiment of the invention comprises the following steps:
the method comprises the following steps: design is based on unilateral expansion theory's stamping engine exhaust nozzle profile
Designing the molded lines of the upper wall surface and the lower wall surface of the unilateral expansion exhaust Nozzle, and adopting an asymmetric sub-combustion ramjet engine exhaust Nozzle molded line Design Method based on the short Nozzle theory (M.Goeing, Nozle Design Optimization by Method-of-characteristics. AIAA-90-2024): all the expansion sections are compressed to a sharp angle of the throat part, and then the flow field is calculated to obtain the spray pipe with the spray pipe shape. For any internal point in a flow field, firstly determining an airflow turning angle and an airflow direction angle of the point by using a shortened length nozzle theory, then calculating the Mach number of the point according to a Prandtl-Meier relation to obtain a Mach angle, and further determining the coordinate of the point according to a characteristic line equation by using the airflow angle and the Mach angle, wherein the specific process is as follows:
initial expansion angle of sharp point of upper and lower throat
And calculating the Mach number of the outlet according to the Mach number of the inlet and the pressure ratio of the inlet and the outlet. And determining a corresponding Prandtl-Meyer expansion angle according to the Mach number of the inlet and the Mach number of the outlet. According to the short nozzle theory, the initial expansion angle of the upper sharp point is determined to be related to the Mach number of the inlet and the outlet of the nozzle, and after an asymmetric factor F is introduced, the initial expansion angle of the lower sharp point is determined. The asymmetry factor F is the ratio of the initial expansion angles at the upper and lower cusp throats, i.e.:
Figure BDA0002033656490000041
wherein the content of the first and second substances,Uis the initial expansion angle of the upper wall surface,Lis the initial expansion angle of the lower wall:
Figure BDA0002033656490000042
L＝F×U
wherein v iseIs the prandtl-meier expansion angle corresponding to the exit mach number,vinis the prandtl-meier expansion angle corresponding to the inlet mach number.
And (II) solving the flow field of the tail nozzle of the asymmetric unilateral expansion ramjet engine.
Referring to fig. 1, according to the characteristic line theory, the flow field of the asymmetric unilateral expansion ramjet engine tail nozzle is solved, so as to determine the flow parameters of each point of the flow field. When the Mach number of a g point obtained by the intersection of the last characteristic line emitted by the upper sharp point a of the inlet and the characteristic line reflected on the lower wall surface is the solved oroMach number, the intersection point f of the last characteristic line of the upper sharp point a and the lower wall surface is the terminal point of the lower wall surface. d is the lower tip of the inlet, and b is the intersection point of the last characteristic line emitted by the lower tip and the upper wall surface. And the point c at which the lower wall surface reflection characteristic line passing through the point g (i.e. the reflection characteristic line passing through the point e) intersects with the upper wall surface is the upper wall surface terminal point. The wall points are determined based on wall boundary conditions. According to the theory of aerodynamic correlation, the angle of expansion and the angle of directivity of the airflow at each point in the area where the expansion lines intersect is determined by the following equation:
ψ＝ψU+ψL-vin
θ＝ψU-ψL
in the formula, #UIs the angle of inflection to the point of the expansion wave emitted at the apex of the upper wall of the portal, psiLIs the inflection angle to which the expansion wave is emitted at the point of the lower wall face of the inlet, psi is the total inflection angle of the expansion of the airflow from the nozzle inlet to that point, i.e., the airflow expansion angle, and theta is the airflow direction angle at that point. In the process of determining the dispersion of the expansion angles of the upper and lower cusps, the expansion angle of the Mach number of the inlet is considered, so that the expansion angles are overlapped twice in the process of overlapping the expansion angles of the airflow, and v is subtractedin。
Obtaining a two-dimensional tail nozzle molded line, determining the rectangular size of a nozzle inlet according to the flow demand of the design state of the sub-combustion ramjet, and performing spanwise stretching (width L) on the two-dimensional molded line to obtain a three-dimensional unilateral expansion sub-combustion ramjet tail nozzle model of the sub-combustion ramjet, as shown in FIG. 2.
Step two: designing exhaust nozzles for turbine engines and rocket engines
The two turbine engine tail nozzles are firstly converged and then combined with the secondary combustion ramjet engine tail nozzle, and the area of the inlet of a single turbine engine, the junction of the two turbine engines and the area of the throat after convergence are calculated according to the flow requirement and the pneumatic parameters (total temperature, total pressure and flow coefficient) of the overall design of the engine.
A flow formula is needed when determining the area a of the nozzle:
Figure BDA0002033656490000051
wherein k is gas constant, air is 0.04042, and gas is 0.039, PtIs the total inlet pressure, TtIs the total temperature of the inlet, and the temperature of the inlet,
Figure BDA0002033656490000052
to design the flow, q (λ) is the flow coefficient, which takes 1 when calculating the throat area.
The circular import of two turbine engines evenly passes through to the rectangle and joins the import, according to above-mentioned flow formula, the area that the import was joined to the design rectangle is the twice of single turbine engine exit area A, combines the design width L of the sub-combustion ramjet of step one, so can calculate the rectangle and join a height and be:
Figure BDA0002033656490000053
and then the mixed gas is converged to the upper wall surface of the secondary combustion ramjet nozzle in the first step, and a throat of a turbine engine channel is formed according to the area of the throat, as shown in figure 3.
The design method of the rocket engine channel is similar to that of the turbine engine channel, firstly the throat area is calculated according to a flow formula, then the uniform circular-square contraction transition is carried out, and the throat area and the uniform circular-square contraction transition are converged with the upper wall surface of the secondary combustion ramjet tail pipe, as shown in figure 4.
Step three: combined configuration of common tail nozzle of parallel TBCC engine
In full speed range flight of a TBCC combined engine, there are multiple flight regimes involving the starting and shutting down of multiple engines. This requires a composite nozzle which not only has good thrust performance when operating with a single engine, but also needs to maintain thrust stability during mode conversion. The layout mode of the TBCC engine combined nozzle has great influence on the problems. The layout scheme to be adopted by the four-channel combined spray pipe is as follows: the sub-combustion ramjet engine channel is positioned at the lowest part, the turbine engine channel is positioned at the middle part, and the rocket engine channel is positioned at the highest part, so that a combined tail jet pipe structure form of an upper parallel type and a lower parallel type is formed. The final structure is shown in fig. 5.
Step four: design adjustment mechanism
Because the Mach number range of the parallel TBCC engine is very wide, in order to ensure that the combined tail nozzle keeps good thrust performance in the whole flight process, an adjusting mechanism must be added at the throat of each engine channel to adjust the throat area.
Referring to fig. 5 and 6, the common exhaust nozzle of the unilateral expansion four-channel combined engine is provided with a rocket engine channel 1, a turbine engine channel 2, a sublagration ramjet engine channel 3, a throat adjusting plate 4 of the rocket engine channel, a throat adjusting plate rotating shaft 5 of the rocket engine channel, a throat adjusting plate 6 of the turbine engine channel, a throat adjusting plate rotating shaft 7 of the turbine engine channel, a throat adjusting rotating shaft 8 of the sublagration ramjet engine channel and a unilateral expansion sublagration ramjet engine exhaust nozzle 9; the sub-combustion ramjet passage 3 is positioned at the lowest part of the tail nozzle and is connected with the tail nozzle 9 of the single-side expansion sub-combustion ramjet, the turbine engine passage 2 is positioned in the middle of the tail nozzle, the rocket engine passage 1 is positioned at the uppermost part of the tail nozzle, and the rocket engine passage 1 and the turbine engine passage 2 are both connected with the upper wall surface of the tail nozzle 9 of the single-side expansion sub-combustion ramjet; the throat adjusting plate 4 of the rocket engine channel is connected with the rocket engine channel 1, the turbine engine channel 2 and the sub-combustion ramjet engine channel 3 through a throat adjusting plate rotating shaft 5 of the rocket engine channel; the throat adjusting plate 6 of the turbine engine channel is connected with the turbine engine channel 2, the sub-combustion ramjet engine channel 3 and the throat adjusting rotating shaft 8 of the sub-combustion ramjet engine channel through a throat adjusting plate rotating shaft 7 of the turbine engine channel; the throat adjusting rotating shaft 8 of the sub-combustion ramjet channel is positioned at the lower part of the tail end of the sub-combustion ramjet channel 3 and is connected with the lower wall surface of the tail nozzle 9 of the single-side expansion sub-combustion ramjet.
Through the connection method, all the parts form a parallel TBCC engine shared exhaust nozzle combined configuration.
And step two, the maximum throat area of the turbine engine channel and the rocket engine channel is obtained, and the design corresponding to the two tail nozzles is completed on the basis of the maximum throat area. When the working state of each tail nozzle is converted, the area of each throat is only required to be reduced to the required throat area. The adjusting mechanism to be adopted by the invention is shown as a throat adjusting plate 4 of a rocket engine channel, a throat adjusting plate 6 of a turbine engine channel and a throat adjusting rotating shaft 8 of a sub-combustion ramjet engine channel in figures 5 and 6, the adjusting mechanism is of a plate structure, and the upper wall surface and the lower wall surface of the plate are parts of the wall surface of the sub-combustion ramjet engine tail pipe, so that the upper wall surface of the complete sub-combustion ramjet engine tail pipe is formed when the adjusting mechanism is closed. The adjusting plate rotates around the rotating shaft to realize opening and closing of the engine channel and adjustment of the throat area, and the rotating shaft of the adjusting plate is tangent to the wall surface of the tail nozzle and the engine channel.
When the parallel type TBCC engine works under different Mach numbers, the working state of the adjusting mechanism is shown in figure 6. When the Mach number is 0 ~ 2, turbine engine passageway is from the open mode gradually change to the closed state, rocket engine passageway flow crescent, and the scramjet engine passageway flow of subluminescence reduces gradually, promptly: the throat adjusting plate 6 of the turbine engine channel rotates upwards around the rotating shaft 7, the throat adjusting plate 4 of the rocket engine channel rotates downwards around the rotating shaft 5, and the throat adjusting rotating shaft 8 of the scramjet engine channel gradually moves upwards; when mach number is at 2 ~ 4, the turbine engine passageway is closed, and rocket engine passageway flow reduces gradually, and the ramjet passageway flow is constantly adjusted in order to avoid taking place the overexpansion, promptly: the throat adjusting plate 4 of the rocket engine channel rotates upwards around the rotating shaft 5 until the throat adjusting plate is closed, and the throat adjusting rotating shaft 8 of the sub-combustion ramjet engine channel moves up and down to adjust the throat area of the sub-combustion ramjet engine channel; when the Mach number is 4 ~ 4.5, rocket engine passageway is closed, and only the scramjet engine that fires is in operating condition, promptly: the throat-adjusting plate 4 of the rocket motor passage rotates upward around the rotating shaft 5 until the rocket motor passage is closed.

Claims (2)

1. A design method of a common tail nozzle of a single-side expansion four-channel combined engine is characterized by comprising the following steps:
1) designing a three-dimensional model of the tail nozzle of the unilateral expansion asymmetric sub-combustion ramjet engine: firstly, designing upper and lower wall molded lines of a unilateral expansion nozzle based on a short nozzle theory to obtain a two-dimensional tail nozzle molded line, then determining the rectangular size of an inlet of the tail nozzle of the sub-combustion ramjet according to the flow demand of a sub-combustion ramjet channel, and laterally stretching the two-dimensional tail nozzle molded line to be as wide as the inlet rectangle to obtain a three-dimensional model of the unilateral expansion asymmetric sub-combustion ramjet;
2) designing tail nozzles of a turbine engine channel and a rocket engine channel: converging two turbine engine tail nozzles, then combining the two turbine engine tail nozzles with a secondary combustion ramjet engine tail nozzle, then calculating the area size of the junction and throat of the two turbine engines according to the flow requirement of the overall design of a parallel TBCC combined engine, combining the design width of the three-dimensional model of the unilateral expansion asymmetric secondary combustion ramjet engine tail nozzle in the step 1), uniformly transferring the two turbine engine circular inlets to a rectangular converging inlet, then converging the rectangular converging inlet to the upper wall surface of the three-dimensional model of the unilateral expansion asymmetric secondary combustion ramjet engine tail nozzle in the step 1) to form the throat of a turbine engine channel, similarly calculating the throat area of the rocket engine channel according to the flow requirement of the overall design of the parallel TBCC combined engine, and then converging the throat area to the upper wall surface of the three-dimensional model of the unilateral expansion asymmetric secondary combustion ramjet engine tail nozzle in the step 1), forming a throat of the rocket engine passage;
3) designing a combined configuration of a common tail nozzle of a parallel TBCC combined engine: the sub-combustion ramjet engine channel is positioned at the lowest part, the turbine engine channel is positioned at the middle part, and the rocket engine channel is positioned at the highest part, so that a parallel TBCC combined engine shared tail nozzle combined configuration is formed;
4) designing an adjusting mechanism: adjusting mechanisms are respectively arranged at the throat of the turbine engine channel, the throat of the rocket engine channel and the throat of the sub-combustion ramjet engine channel, the adjusting mechanisms adopt a rotating shaft and an adjusting plate, the throat area of the turbine engine channel and the throat area of the rocket engine channel are changed by rotating the adjusting plate around the rotating shaft, and the throat area of the sub-combustion ramjet engine channel is changed by moving the rotating shaft up and down.
2. The common tail nozzle of the unilateral expansion four-channel combined engine is characterized by being provided with a rocket engine channel, a turbine engine channel, a sub-combustion ramjet engine channel, a throat adjusting plate of the rocket engine channel, a throat adjusting plate rotating shaft of the rocket engine channel, a throat adjusting plate of the turbine engine channel, a throat adjusting plate rotating shaft of the turbine engine channel, a throat adjusting rotating shaft of the sub-combustion ramjet engine channel and a unilateral expansion sub-combustion ramjet engine tail nozzle; the secondary combustion ramjet passage is positioned at the lowest part of the tail nozzle and connected with the tail nozzle of the unilateral expansion secondary combustion ramjet, the turbine engine passage is positioned in the middle of the tail nozzle, the rocket engine passage is positioned at the uppermost part of the tail nozzle, and the rocket engine passage and the turbine engine passage are both connected with the upper wall surface of the tail nozzle of the unilateral expansion secondary combustion ramjet; the throat adjusting plate of the rocket engine channel is connected with the rocket engine channel, the turbine engine channel and the sub-combustion ramjet engine channel through a throat adjusting plate rotating shaft of the rocket engine channel; the throat adjusting plate of the turbine engine channel is connected with the throat adjusting rotating shafts of the turbine engine channel, the sub-combustion ramjet engine channel and the sub-combustion ramjet engine channel through the throat adjusting plate rotating shaft of the turbine engine channel; the throat adjusting rotating shaft of the sub-combustion ramjet channel is positioned at the lower part of the tail end of the sub-combustion ramjet channel and is connected with the lower wall surface of the tail nozzle of the unilateral expansion sub-combustion ramjet.
CN201910317706.3A 2019-04-19 2019-04-19 Design method of common tail nozzle of single-side expansion four-channel combined engine Active CN110020500B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201910317706.3A CN110020500B (en) 2019-04-19 2019-04-19 Design method of common tail nozzle of single-side expansion four-channel combined engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201910317706.3A CN110020500B (en) 2019-04-19 2019-04-19 Design method of common tail nozzle of single-side expansion four-channel combined engine

Publications (2)

Publication Number Publication Date
CN110020500A CN110020500A (en) 2019-07-16
CN110020500B true CN110020500B (en) 2021-01-01

Family

ID=67191924

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201910317706.3A Active CN110020500B (en) 2019-04-19 2019-04-19 Design method of common tail nozzle of single-side expansion four-channel combined engine

Country Status (1)

Country Link
CN (1) CN110020500B (en)

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112228243A (en) * 2020-10-16 2021-01-15 中国航发四川燃气涡轮研究院 Hydraulic push type cam mechanism for variable state adjustment of combined spray pipe
CN114000958B (en) * 2021-10-29 2022-11-08 厦门大学 Design method of TBCC tail nozzle adjusting mechanism based on double-plectrum type sliding block principle

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN108561244A (en) * 2017-12-29 2018-09-21 厦门大学 A kind of three power combination engine design methods that super burn coexists with sub- combustion combustion chamber
CN109236496A (en) * 2018-11-15 2019-01-18 厦门大学 The three power combination engine design methods that Asia combustion can be switched with super burn channel
CN109455309A (en) * 2018-11-19 2019-03-12 厦门大学 Rider air intake duct integrated design method in sweepforward based on circular cone precursor shock wave
CN109538377A (en) * 2018-11-15 2019-03-29 厦门大学 Share three power combination engine design methods of sub- combustion combustion chamber

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3974648A (en) * 1968-08-19 1976-08-17 United Technologies Corporation Variable geometry ramjet engine
US7886516B2 (en) * 2006-12-18 2011-02-15 Aerojet-General Corporation Combined cycle integrated combustor and nozzle system
CN106050472A (en) * 2016-07-08 2016-10-26 西北工业大学 Turbo-rocket combined ramjet engine and operating method thereof
CN106168185A (en) * 2016-07-08 2016-11-30 西北工业大学 Air turbine punching press combined engine and method of work thereof
CN108194224B (en) * 2017-12-31 2019-08-27 厦门大学 It is embedded to the TBCC parallel connection jet pipe regulating mechanism design method of diaphragm internal
CN108717487A (en) * 2018-05-17 2018-10-30 中国航空发动机研究院 Air inlet adjustable punching engine integration Optimal Design of Runner System method

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN108561244A (en) * 2017-12-29 2018-09-21 厦门大学 A kind of three power combination engine design methods that super burn coexists with sub- combustion combustion chamber
CN109236496A (en) * 2018-11-15 2019-01-18 厦门大学 The three power combination engine design methods that Asia combustion can be switched with super burn channel
CN109538377A (en) * 2018-11-15 2019-03-29 厦门大学 Share three power combination engine design methods of sub- combustion combustion chamber
CN109455309A (en) * 2018-11-19 2019-03-12 厦门大学 Rider air intake duct integrated design method in sweepforward based on circular cone precursor shock wave

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
"TBCC Engine Inlet Design and Ramp Angle Optimization";J.D.Zhang .et al;《New Trends in Fluid Mechanics Research》;20091231;page152-455 *
"三维并联式ＴＢＣＣ发动机排气系统设计与实验";花文达等;《航空动力学报》;20180930;第33卷(第9期);第2268-2277页 *

Also Published As

Publication number Publication date
CN110020500A (en) 2019-07-16

Similar Documents

Publication Publication Date Title
CN104632411B (en) Binary is used to become the interior rider type turbine base assembly power air intake duct of geometric ways
CN106321283B (en) The pneumatic propelling integrated layout method of hypersonic aircraft based on assembly power
US3792584A (en) Increased or variable bypass ratio engines
CN106285946B (en) The channel of double-axle rotation deformation becomes geometry air intake duct without rider formula in wedge angle
CN109670269B (en) Design method of multi-channel parallel three-power combined engine
CN105156227B (en) Pre-cooling air-breathing type variable cycle engine
CN107013368B (en) Turbine base double combustion chamber's punching press combined cycle engine control method
CN103879556B (en) Wide flight envelope morphing aircraft
CN107013367B (en) Turbine base double combustion chamber's punching press combined cycle engine
CN110020500B (en) Design method of common tail nozzle of single-side expansion four-channel combined engine
US20090313968A1 (en) Hypersonic Inlet Systems and Methods
CN106837550A (en) The method for designing of hypersonic triple channel air intake duct
CN105539862B (en) A kind of combination in series power air intake duct flatly moving type mode conversion device
CN110541773B (en) Wide-speed-range ramjet engine combustion chamber and working method thereof
CN214403792U (en) Series-parallel three-power combined engine
CN109733634B (en) Design method of three-dimensional inward-turning four-channel hypersonic combined air inlet channel
CN109236496B (en) Design method of three-power combined engine with switchable sub-combustion and super-combustion channels
US10443539B2 (en) Hybrid exhaust nozzle
CN108194224B (en) It is embedded to the TBCC parallel connection jet pipe regulating mechanism design method of diaphragm internal
CN203740126U (en) Morphing aircraft with wide flight envelope
US20160090174A1 (en) Reaction drive blade tip with turning vanes
CN109538377B (en) Design method of three-power combined engine sharing sub-combustion chamber
CN109408993B (en) Design method of turbofan and ramjet combined engine with rocket inside
CN112443422B (en) Rotary vertical take-off and landing spray pipe based on quasi-axisymmetric throat offset type pneumatic vectoring spray pipe and design method thereof
CN207297170U (en) Hypersonic triple channel air intake duct

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant