CN112948967B - Series-parallel three-power combined engine design method - Google Patents

Series-parallel three-power combined engine design method Download PDF

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CN112948967B
CN112948967B CN202110173210.0A CN202110173210A CN112948967B CN 112948967 B CN112948967 B CN 112948967B CN 202110173210 A CN202110173210 A CN 202110173210A CN 112948967 B CN112948967 B CN 112948967B
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朱剑锋
韦宇卿
尤延铖
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Xiamen University
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Abstract

A series-parallel three-power combined engine design method comprises the following steps that 1) the maximum windward runner sectional area of combined power is determined according to the size index of an aircraft to the combined power; the engine air inlet passage adopts an axial symmetry structure that an air inlet blocking cone can move back and forth, and the capture area of the air inlet passage is selected as the maximum windward runner sectional area; 2) calculating the flow coefficient and the Mach number of the maximum section of the turbine engine and the inlet sectional area of a sub-combustion combustor of the ramjet engine according to the lowest Mach number of the single stamping operation; 3) calculating the throat area of the adjustable nozzle according to the outlet flow, the total pressure and the total temperature of the secondary combustion chamber of the ramjet engine under different flight conditions; 4) calculating to obtain the thrust and the oil consumption rate of the engine according to the outlet flow, the total pressure, the total temperature and the like of the spray pipe under different flight conditions; 5) completing the design of the length along the flow direction; 6) and the size and the number of the rocket engines are designed by combining the maximum thrust requirement of the rocket engines in the whole flight section.

Description

Series-parallel three-power combined engine design method
Technical Field
The invention relates to the field of combined engines, in particular to a design method of a series-parallel series-parallel three-power combined engine.
Background
Hypersonic flight is considered as a third revolution after propeller and jet propulsion, and a hypersonic aircraft is developed, wherein the most important technology is a hypersonic propulsion technology, and the performance of a propulsion system directly influences the success of the design scheme of the whole aircraft. However, in different mach number ranges, various air-breathing propellers have different economical efficiency, and therefore, to meet the power requirements of the air-breathing aircraft in the full working range, the combined power system is an ideal scheme with great engineering application prospect. At present, the common combined power system available for the choice of the hypersonic aircraft has two main forms: turbine-based combined cycle power plants (TBCC) and rocket-based combined cycle power plants (RBCC). The turbine-based combined cycle engine is a power device formed by combining a turbine engine and other types of engines, has the advantages of high specific thrust, wide flight speed range, conventional take-off and landing, reusability, good low-speed performance and the like, and is one of key power systems for realizing autonomous acceleration, landing with a power level and repeated use of the hypersonic aircraft.
The turbine base combined cycle can realize a variable cycle working process, so that the aircraft can obtain good propulsion performance under different flight conditions (subsonic speed, supersonic speed and hypersonic speed). When the aircraft flies at low speed, airflow enters the turbine engine, and the combined engine works in a turbine engine mode; during high-speed flight of the aircraft, the air flow enters the ramjet engine, and the combined engine operates in the ramjet engine mode.
The turbine-based combined power has the advantage of high propulsion efficiency, but at present, two main problems exist. Firstly, the optimal operating speed range of the current turbine engine is Ma 0-2; the beneficial operating speed range Ma3 ~ 5 of the scramjet, there is operating speed range difference between the two, shows as "thrust gap" phenomenon. Secondly, the turbine and the ramjet are used as air suction type power, the mutual combination of the two powers inevitably increases the size and the weight, the windward area is increased due to the size increase in the high-speed flight process, the unit windward thrust is reduced, the aircraft resistance is inevitably increased, the combined power is increased, and the thrust-resistance imbalance phenomenon of the aircraft/the combined power is aggravated.
Disclosure of Invention
The invention aims to solve the problems of thrust gap and low unit head-on thrust of turbine-based combined power in the prior art, and provides a design method of a three-power combined engine in which a turbine engine, a rocket engine and a sub-combustion ramjet are connected in series-parallel, so that the overall performance of a propulsion system is improved while the advantages of wide-speed-range flight and high thrust are maintained.
In order to achieve the purpose, the invention adopts the following technical scheme:
a design method of a series-parallel series-parallel three-power combined engine comprises the following steps:
1) determining the maximum windward flow passage sectional area of the combined power according to the size index of the aircraft to the combined power; for combined power with the working Mach number of Ma 0-5, an engine air inlet adopts an axisymmetric structure that an air inlet blocking cone can move back and forth, and the capture area of the air inlet is selected as the maximum windward flow passage sectional area;
2) calculating the flow coefficient and the Mach number of the maximum section of the turbine engine in the state according to the lowest Mach number of the single stamping operation, and calculating the inlet sectional area of the scramjet sub-combustion combustor;
3) calculating to obtain the throat area of the adjustable nozzle according to the outlet flow, the total pressure and the total temperature of the secondary combustion chamber of the ramjet engine under different flight conditions;
4) calculating to obtain the Mach number of the outlet of the adjustable spray pipe according to the flow, the total pressure, the total temperature and the size indexes of the outlet of the spray pipe under different flight conditions, obtaining the outlet speed and the pressure of the spray pipe according to the Mach number of the outlet, and further calculating to obtain the thrust and the oil consumption rate of an engine;
5) the sectional area of the typical section along the flow direction can be obtained through the steps 1-4), and the length design along the flow direction is further completed based on the sectional area;
6) and (3) based on the combined engine profile obtained in the step 5), combining the maximum thrust requirement of the rocket engines in the whole flight section, and completing the size and number design of the rocket engines, wherein the rocket engines are arranged on the outlet end face of the turbine engine.
In the step 1), the maximum windward flow passage sectional area A of the combined powermax＝k·AIndex (I)，AIndex (I)For the dimensional index, k is a coefficient for measuring the relative radial proportion of the casing and the accessory outside the inlet runner.
In step 2), the flow coefficient q (M) of the maximum section of the turbine engine1) And Mach number M1Q (M) is continuous according to flow0)·A0＝σ1·q(M1)(AIndex (I)-ATE) Is solved for, σ1For the total pressure recovery coefficient of the combined air inlet passage, ATEIs the maximum cross-sectional area of the turbine engine, AIndex (I)As a size index, A0For air intake trapping area, M0Minimum Mach number, q (M), for single stamping operations0) Is M0The corresponding flow coefficient; the inlet cross-sectional area A of the scramjet sub-combustion chamber3Q (M) is continuous according to flow0)·A0＝σ1·σ2·q(M3)·A3Is solved for, σ2For the total pressure recovery coefficient, M, of the outer ring flow channel of the turbine channel3Mach number, q (M) of the inlet cross-section of the combustion chamber3) The flow coefficient of the inlet section of the scramjet sub-combustion chamber is shown.
In step 3), the throat area A of the variable nozzle8Is m8＝K·P8·/T8 0.5·q(M8)·A8，m8The flow rate of the outlet of the secondary combustion chamber of the ramjet engine is K, the gas constant is P8Total pressure at outlet of secondary combustion chamber of ramjet engine8For the total temperature of the outlet of the secondary combustion chamber of the ramjet engine, M8Is Mach number, q (M) at the throat of the variable nozzle8) Is M8The corresponding flow coefficient.
In step 4), the Mach number M of the outlet of the adjustable nozzle9From the flow equation m9＝K·P9·/T9 0.5·q(M9)·AIndex (I)Is calculated to obtain m9Is the nozzle outlet flow, P9Is total pressure and T at the outlet of the spray pipe9Is the total temperature of the nozzle outlet, AIndex (I)For dimensional indications, K is the gas constant.
In the step 5), the length of the air inlet blocking cone of the combined air inlet channel is determined by the angle alpha of the blocking cone1And the maximum windward flow passage sectional area AmaxAnd determining the cone angle of the blocking cone according to the designed Mach number of the combined air inlet channel and a cone shock wave pneumatic formula.
In the step 5), the length from the throat section of the air inlet channel to the inlet section of the turbine engine is determined according to the expansion angle of the flow channel and the space requirement of a flow distribution plate of a drainage channel at the inlet of the turbine engine; the length of the section from the outlet of the turbine engine to the inlet of the scramjet engine sub-combustion chamber is determined by the space requirement of a diversion channel splitter plate at the outlet of the turbine engine; the length of the nozzle is the equivalent diameter of the maximum cross-sectional area of the nozzle outlet.
In the step 6), the maximum outlet diameter of the nozzle of the rocket engine is limited by the inner diameter and the outer diameter of the outlet end face, the thrust of a single rocket engine can be calculated according to the outlet area of the nozzle of the rocket engine, the number of the rocket engines can be calculated by combining the maximum thrust requirement of an aircraft on the rocket engine, and the rocket engines are circumferentially distributed along the outlet end face of the turbine engine.
A series-parallel three-power combined engine comprises an air inlet blocking cone, an engine air inlet channel, an inlet drainage channel splitter plate, a turbine engine, an outlet drainage channel splitter plate, a rocket engine, a ramjet sub-combustion chamber and a shared tail nozzle which are sequentially arranged;
the turbine engine is arranged in an air inlet channel of the turbine engine, and an ejector rocket-sub-combustion stamping outer ring combined channel is arranged outside the air inlet channel of the turbine engine in parallel;
the air inlet blocking cone is arranged at an inlet at the front end of an engine air inlet channel, the engine air inlet channel is provided with two outlets which are respectively connected with an inlet of an air inlet channel of the turbine engine and an inlet of an ejection rocket-sub-combustion stamping outer ring combined channel; the inlet drainage channel splitter plate is arranged between an inlet of an air inlet channel of the turbine engine and an inlet of the ejection rocket-sub-combustion stamping outer ring combined channel, and the outlet drainage channel splitter plate is arranged between an outlet of the air inlet channel of the turbine engine and an outlet of the ejection rocket-sub-combustion stamping outer ring combined channel;
the rocket engine is arranged behind the turbine engine and the ejector rocket-sub-combustion ramjet outer ring combined channel, the ramjet sub-combustion chamber is arranged behind the rocket engine, and the shared tail nozzle is arranged behind the ramjet sub-combustion chamber through the throat of the area-adjustable nozzle.
The inlet drainage channel flow distribution plate and the outlet drainage channel flow distribution plate are arranged in an openable fish scale-shaped structure in a multi-layer crossed and overlapped mode.
Compared with the prior art, the technical scheme of the invention has the following beneficial effects:
the three-power combined engine of the ejection rocket-sub-combustion ramjet outer-ring combined channel and the series-parallel hybrid combines the characteristics of the full speed domain of the rocket engine and the high performance of the turbine engine, solves the problem of thrust gap on the premise of not reducing the thrust performance, and can effectively reduce the space requirement of the engine by sharing the afterburning chamber of the turbine engine, the combustion chamber of the sub-combustion ramjet engine and the combustion chamber of the ejection rocket engine. The combined engine designed by the method solves the problems of 'thrust gap' and limited space design of turbine-based combined power, and has the advantages of compact structure, outstanding overall performance and moderate technical difficulty.
The invention realizes integrated design by sharing the afterburner of the turbine engine and the sub-combustion combustor of the ramjet engine; by introducing the rocket engine, on one hand, the turbine/ram thrust bridging is realized, and simultaneously, the promotion of the combined power unit head-on thrust is synchronously realized, so that the contradiction between the requirement of high thrust in the takeoff-acceleration stage and the requirement of lower thrust and high efficiency in cruising is effectively solved. In addition, the rocket engine is arranged at the front end of the scramjet sub-combustion chamber, and ignition and stable working range of the scramjet sub-combustion chamber can be improved by advanced work of the rocket engine under the environment that high altitude and low pressure are not beneficial to ignition of the scramjet sub-combustion chamber. In addition, the air inlet blocking cone and the fish scale type spray pipe related to the design method are already practical in engineering in conventional turbines and ramjet engines, and the realizability is high.
Drawings
Fig. 1 is a schematic cross-sectional view of the overall layout of a three-power combined engine.
FIG. 2 is a partially enlarged schematic view of a combined turbine-sub-combustion path of a three power combined engine.
Fig. 3 is a schematic diagram of the overall structure of the three-power combined engine.
FIG. 4 is a schematic cross-sectional structural view of a three-power combined engine in a closed state of an ejection rocket-sub-combustion stamping outer ring combined passage.
FIG. 5 is a schematic cross-sectional structural view of an open state of an ejection rocket-sub-combustion stamping outer ring combined passage of the three-power combined engine.
FIG. 6 is a front view of a diversion plate structure of the drainage channel.
Reference numerals: the device comprises an air inlet blocking cone 1, an engine air inlet channel 2, a turbine engine air inlet channel 4, an ejection rocket-sub-combustion ramjet outer ring combined channel 5, a rocket engine 6, a ramjet sub-combustion chamber 7, a shared tail nozzle 8, an area-adjustable nozzle throat 9, a flow guide channel splitter plate 10 and a turbine engine 11; the flow guide channel splitter plate 10 is arranged on the front end of the flow guide channel splitter plate, and the area adjustable nozzle throat 9 is arranged on the rear end of the flow guide channel splitter plate.
Detailed Description
In order to make the technical problems, technical solutions and advantageous effects of the present invention more clear and obvious, the present invention is further described in detail below with reference to the accompanying drawings and embodiments.
The method is designed by taking the working Mach number of a combined engine as Ma 0-5, the independent working Mach number of a turbine engine as 0-2.5, the independent working Mach number of a ramjet as 3-5 and the Mach number as 2.5-3 as an example of turbine/ramjet mode conversion, and comprises the following steps of:
1) dimension index A for aircraft to combined powerIndex (es)Determining the maximum windward flow passage sectional area A of the combined powermax＝k·AIndex (I)K is a coefficient for measuring the radial relative proportion of the casing and accessories outside the channel of the air inlet channel, and can be generally 0.95;
2) for combined power with the working Mach number of Ma 0-5, an engine air inlet adopts an axisymmetric structure that an air inlet blocking cone can move back and forth, and the capture area A of the air inlet is0The maximum windward flow passage sectional area A is selectedmax。
3) Considering that the turbine engine in the turbine-based combined power is generally a certain type of off-the-shelf engine, the maximum cross-sectional area A of the engine isTEIs generally a fixed value, and in this embodiment, assume ATE＝0.7AIndex (I)。
4) Calculating the flow coefficient q (M) of the largest section of the turbine engine in this state according to the lowest Mach number of the individual stamping work1) And Mach number M1Which is continuous q (M) according to the flow rate0)·A0＝σ1·q(M1)(AIndex (I)-ATE) The solution is carried out, in the embodiment, the combined power working speed range is 0-5, and the lowest working Mach of the ramjet engineNumber Ma0Is 3, σ1The total pressure recovery coefficient of the combined air inlet passage is calculated by a formula sigma1＝1-0.075(Ma0-1)1.35，σ10.809, flow coefficient q (M)1) 0.829, corresponding to Mach number M10.585; a. theTEIs the maximum cross-sectional area of the turbine engine, AIndex (I)As a size index, A0For air intake trapping area, M0Minimum Mach number, q (M), for single stamping operations0) Is M0The corresponding flow coefficient; design M considering flow loss and flow capacity1It is recommended to choose between 0.5 and 0.6, if M is present1Higher or lower, the performance index of the combined air inlet or turbine engine needs to be further improved.
5) Calculating the inlet cross-sectional area A of the sub-combustion chamber of the ramjet according to the lowest Mach number of the single stamping work3Which is continuous q (M) according to the flow rate0)·A0＝σ1·σ2·q(M3)·A3The solution is carried out, and the working speed range of the combined power is 0-5 and the lowest working Mach number Ma of the ramjet in the embodiment0Is 3, σ2Is the total pressure recovery coefficient of the outer ring flow passage (namely the ejection rocket-sub-combustion stamping outer ring combined passage) of the turbine passage, sigma2It may be generally 0.97, M3Mach number, q (M) of the inlet cross-section of the combustion chamber3) The Mach number M of the inlet section of the general combustion chamber is taken into consideration of the combustion organization problem of the scramjet sub-combustion chamber for the flow coefficient at the inlet section of the scramjet sub-combustion chamber3Taken as 0.2, q (M)3) Is 0.337, the cross-sectional area A of the scramjet sub-combustion chamber can be obtained3。
6) Calculating to obtain the throat area A of the adjustable nozzle according to the outlet flow, the total pressure and the total temperature of the secondary combustion chamber of the ramjet engine under different flight conditions8Wherein the throat area A of the adjustable nozzle8Is m8＝K·P8·/T8 0.5·q(M8)·A8，m8The flow rate of the outlet of the secondary combustion chamber of the ramjet engine is K, the gas constant is P8For sub-combustion of ramjet enginesTotal pressure at the outlet of the furnace, T8For the total temperature of the outlet of the secondary combustion chamber of the ramjet engine, M8The Mach number, M, at the throat of the adjustable nozzle8＝1，q(M8) Is M8Corresponding flow coefficient, q (M)8) 1, total pressure recovery coefficient sigma in a sub-combustion chamber of a ramjet engine3It can be 0.98, the Mach number of the throat section is 1.0, and the total temperature T of the outlet temperature of the combustion chamber8Is determined by the amount of oil supplied.
7) Calculating to obtain the Mach number of the outlet of the adjustable spray pipe according to the indexes of the outlet flow, the total pressure, the total temperature and the size of the spray pipe under different flight conditions, obtaining the outlet speed and the pressure of the spray pipe according to the Mach number of the outlet, and further calculating to obtain the thrust and the oil consumption rate of the engine; wherein the Mach number M of the outlet of the adjustable nozzle9From the flow equation m9＝K·P9·/T9 0.5·q(M9)·AIndex (I)Is calculated to obtain m9Is the nozzle outlet flow, P9Is total pressure and T at the outlet of the spray pipe9Is the total temperature of the nozzle outlet, AIndex (I)For dimensional index, K is the gas constant.
8) The above 7 steps can obtain the sectional area of the typical section along the flow direction, and based on the sectional area, the length along the flow direction can be further designed. Wherein the air inlet of the combined air inlet channel blocks the length L of the awl1(length from the front edge point of the plugging cone to the cross section of the throat) is determined by the angle alpha of the plugging cone1And the maximum windward flow passage sectional area AmaxAnd determining the cone angle of the blocking cone according to the designed Mach number of the combined air inlet channel and a cone shock wave pneumatic formula.
9) Length L from throat section of air inlet channel to inlet section of turbine engine2Two factors need to be considered. One is the expansion flow loss of the flow channel, the expansion angle of the flow channel is made to be smaller as much as possible, but the length is too long due to the excessively low expansion angle, and the expansion angle is selected to be about 15 degrees; and secondly, the space requirement of a flow distribution plate (an annular fish scale structure which can be controlled to open and close) of a drainage channel at the inlet of the turbine engine needs to be considered, and the space requirement needs to be comprehensively determined by combining the specific structural form of the fish scale.
10) Turbine engine length L3It is started by spotThe machine decision is typically a fixed value.
11) Section length L from outlet of turbine engine to inlet of scramjet sub-combustion chamber4The turbine engine outlet flow guide channel splitter plate is mainly determined by the space requirement of a turbine engine outlet flow guide channel splitter plate (an annular fish scale structure capable of being controlled to open and close), and the space requirement needs to be comprehensively determined by combining the specific structural form of the fish scale.
12) Ramjet sub-combustion chamber length L5Which is determined by the combustion texture characteristics, can be designed to be 1.2m in general.
13) Length L of nozzle5Which is determined by the minimum area of the throat of the nozzle and the maximum cross-sectional area of the outlet, can generally be designed to be the equivalent diameter of the maximum cross-sectional area of the outlet.
14) Based on the combined engine profiles obtained in the steps 1-13, the size and number of rocket engines can be designed by combining the maximum thrust requirement of the rocket engines in the whole flight profile. The rocket engine is arranged on the outlet end surface of the turbine engine, and the maximum outlet diameter of the spray pipe is limited by the inner diameter and the outer diameter of the outlet end surface. The thrust of a single rocket engine can be calculated according to the area of the outlet of the rocket engine nozzle, and the number of the rocket engines can be calculated by combining the maximum thrust requirement of the aircraft on the rocket engines, wherein the rocket engines are circumferentially distributed along the outlet end face of the turbine engine.
As shown in fig. 1 to 5, the series-parallel three-power combined engine of the invention has an axisymmetric structure as a whole, and comprises an air inlet blocking cone 1, an engine air inlet 2, a flow guide channel splitter plate 10, a turbine engine 11, a rocket engine 6, a ramjet sub-combustion chamber 7 and a common tail nozzle 8 which are sequentially arranged;
the turbine engine 11 is arranged in a turbine engine air inlet passage 4, and an ejection rocket-sub-combustion stamping outer ring combined passage 5 is arranged outside the turbine engine air inlet passage 4 in parallel;
the air inlet blocking cone 1 is arranged at an inlet at the front end of an engine air inlet channel 2, the engine air inlet channel 2 is provided with two outlets which are respectively connected with an inlet of a turbine engine air inlet channel 4 and an inlet of an ejection rocket-sub-combustion stamping outer ring combined channel 5;
the flow guide channel splitter plate 10 comprises an inlet flow guide channel splitter plate and an outlet flow guide channel splitter plate, specifically, the inlet flow guide channel splitter plate is arranged between an inlet of a turbine engine air inlet channel 4 and an inlet of an ejection rocket-sub combustion stamping outer ring combined channel 5, and the outlet flow guide channel splitter plate is arranged between an outlet of the turbine engine air inlet channel and an outlet of the ejection rocket-sub combustion stamping outer ring combined channel; wherein, the diversion channel splitter plate 10 is used for controlling the conversion of airflow between the turbine engine air inlet channel 4 and the ejection rocket-sub-combustion stamping outer ring combined channel 5;
the rocket engine 6 is arranged behind the turbine engine 11 and the ejection rocket-sub-combustion ramjet outer ring combined channel 5, the ramjet sub-combustion chamber 7 is arranged behind the rocket engine 6, and the shared tail nozzle 8 is arranged behind the ramjet sub-combustion chamber 7 through the area-adjustable nozzle throat 9.
As shown in fig. 6, the diversion channel flow distribution plate 10 is provided with an openable fish scale-shaped structure, and is arranged in a multi-layer crossed and overlapped mode, so that good sealing performance is achieved.
The working principle of the invention is as follows:
when the flying Mach number is 0-2, the diversion plates 10 of the front and rear diversion channels of the turbine engine 11 are all located at the first position, the turbine engine 11 is ignited to work, at the moment, the ejection rocket-secondary combustion stamping outer ring combined channel 5 is in a closed state, and the throat 9 of the area-adjustable spray pipe is in a state of three;
when the flying Mach number is 2-3, the turbine engine 11 stops working, the diversion plate 10 of the drainage channel moves downwards to the position II, the air inlet channel 4 of the turbine engine is closed, the rocket engine 6 and the ramjet engine sub-combustion chamber 7 are ignited to work, airflow at the outlet of the rocket engine 6 flows into the ramjet engine sub-combustion chamber 7 to be mixed and combusted with fuel oil, the throat 9 of the area-adjustable spray pipe is in a solid line state shown in figure 1, and then the area-adjustable spray pipe expands through the shared tail spray pipe 8 to do work to generate thrust;
when the flying Mach number is 3-5, the rocket engine 6 stops working, the diversion plate 10 of the drainage channel keeps the position II, the scramjet engine sub-combustion chamber 7 continues to burn, the throat 9 of the area-adjustable jet pipe is in the state IV, and airflow expands through the shared tail jet pipe 8 to do work to generate thrust.
According to the requirements of flow and thrust when the engine works, the throat area of the shared tail nozzle 8 is adjusted through the throat 9 of the area-adjustable nozzle in different states of the third stage to improve the thrust performance of the shared tail nozzle 8.
Rocket engines, as a non-aspirated power, can operate at full speed, although at a lower thrust than turbine and ramjet engines, and have significant advantages in terms of specific head-on thrust. Therefore, the small-size rocket engine is introduced into the turbine-based combined power, on one hand, the bridge connection of a thrust gap of the turbine/ramjet engine can be realized, on the other hand, the acceleration characteristic of the aircraft can be improved by using the large unit head-on thrust of the rocket engine, the comprehensive optimization of the aircraft and the combined power is realized, the problem that the low-Mach-number rocket engine cannot generate the thrust without starting is solved, and the effect of continuous thrust is achieved.
The invention not only introduces the nest mortar for carrying out thrust bridging of a turbine/ramjet by a rocket engine, but also focuses on the power design requirements of high thrust and high cruising efficiency in the takeoff-acceleration stage by introducing a rocket engine technology and a series-parallel hybrid combination mode, and realizes the comprehensive optimization of the performance of the wide-speed-range hypersonic aircraft/combined power. The invention has the advantages of simple structure, moderate technical difficulty, easy realization, repeated use and the like.

Claims (10)

1. A design method of a series-parallel series-parallel three-power combined engine is characterized by comprising the following steps:
1) determining the maximum windward flow passage sectional area of the combined power according to the size index of the aircraft to the combined power; for combined power with the working Mach number of Ma 0-5, an engine air inlet adopts an axial symmetry structure with an air inlet blocking cone moving back and forth, and the capture area of the air inlet is selected as the maximum windward flow channel sectional area;
2) calculating the flow coefficient and the Mach number of the maximum section of the turbine engine in the state according to the lowest Mach number of the single stamping operation, and calculating the inlet sectional area of the scramjet sub-combustion combustor;
3) calculating to obtain the throat area of the adjustable nozzle according to the outlet flow, the total pressure and the total temperature of the secondary combustion chamber of the ramjet engine under different flight conditions;
4) calculating to obtain the Mach number of the outlet of the adjustable spray pipe according to the indexes of the outlet flow, the total pressure, the total temperature and the size of the spray pipe under different flight conditions, obtaining the outlet speed and the pressure of the spray pipe according to the Mach number of the outlet, and further calculating to obtain the thrust and the oil consumption rate of the engine;
5) obtaining the sectional area of the typical section along the flow direction through the steps 1) to 4), and designing the length along the flow direction based on the sectional area;
6) based on the combined engine profile obtained in the step 5), the size and number design of the rocket engines is completed by combining the maximum thrust requirement of the rocket engines in the whole flight profile, and the rocket engines are arranged on the outlet end face of the turbine engine.
2. The method for designing a series-parallel combined three-power engine as claimed in claim 1, wherein: in step 1), the maximum upwind flow passage sectional area of the combined powerA max =k·A Index (I) ，A Index (I) As an index of the size,kthe radial relative proportion coefficient of the casing and accessories outside the inlet passage is measured.
3. The method for designing a series-parallel combined three-power engine as claimed in claim 1, wherein: in step 2), the flow coefficient of the largest cross section of the turbine engineq(M 1 )And Mach numberM 1 According to flow rateq(M 0 ) ·A 0 =σ 1 ·q(M 1 ) (A Index (I) -A TE ) Is solved for, σ 1 In order to recover the coefficient of the total pressure of the combined air inlet channel,A TE for a turbine hairThe maximum cross-sectional area of the motive machine,A index (I) As an index of the size,A 0 in order to capture the area of the air inlet channel,M 0 at the lowest mach number of the individual stamping operation,q(M 0 ) Is M 0 What is needed is Corresponding flow coefficient(ii) a Inlet cross-sectional area of the scramjet sub-combustion chamberA 3 According to flow rateq(M 0 ) ·A 0 =σ 1 ·σ 2 ·q(M 3 ) ·A 3 Is solved for, σ 2 The total pressure recovery coefficient of the outer ring flow passage of the turbine channel is obtained,M 3 mach of the combustion chamber inlet cross section Number, q (M) 3 )The flow coefficient of the inlet section of the scramjet sub-combustion chamber is shown.
4. The method for designing a series-parallel combined three-power engine as claimed in claim 1, wherein: in step 3), the throat area of the nozzle is adjustedA 8 Is calculated by the formulam 8 =K·P 8 / T 8 0.5 ·q(M 8 )·A 8 ，m 8 Is the outlet flow of the scramjet sub-combustion chamber, K is a gas constant,P 8 the total pressure of the outlet of the sub-combustion chamber of the ramjet engine,T 8 is the total temperature of the outlet of the scramjet sub-combustion chamber,M 8 is the Mach number at the throat of the adjustable nozzle,q(M 8 )is composed ofM 8 The corresponding flow coefficient.
5. The method for designing a series-parallel combined three-power engine as claimed in claim 1, wherein: in step 4), the Mach number of the outlet of the adjustable nozzleM 9 From the flow equationm 9 =K·P 9 /T 9 0.5 ·q(M 9 )·A Index (I) The calculation results in that,m 9 is the flow rate of the outlet of the spray pipe,P 9 is the total pressure of the nozzle outlet,T 9 Is the total temperature of the outlet of the spray pipe,A index (I) For dimensional indications, K is the gas constant.
6. The method for designing a series-parallel combined three-power engine as claimed in claim 1, wherein: in step 5), the length of the air inlet blocking cone of the combined air inlet channel is determined by the angle of the blocking coneα 1 And maximum windward flow passage cross-sectional areaA max And determining the cone angle of the blocking cone according to the designed Mach number of the combined air inlet channel and a cone shock wave pneumatic formula.
7. The method for designing a series-parallel combined three-power engine as claimed in claim 1, wherein: in the step 5), the length from the throat section of the air inlet channel to the inlet section of the turbine engine is determined according to the expansion angle of the flow channel and the space requirement of a flow distribution plate of a drainage channel at the inlet of the turbine engine; the length of the section from the outlet of the turbine engine to the inlet of the scramjet engine sub-combustion chamber is determined by the space requirement of a diversion channel splitter plate at the outlet of the turbine engine; the length of the nozzle is the equivalent diameter of the maximum cross-sectional area of the nozzle outlet.
8. The method for designing a series-parallel three-power combined engine as claimed in claim 1, wherein: in the step 6), the maximum outlet diameter of the nozzle of the rocket engine is limited by the inner diameter and the outer diameter of the outlet end surface, the thrust of a single rocket engine is obtained through calculation according to the outlet area of the nozzle of the rocket engine, the maximum thrust requirement of an aircraft on the rocket engine is combined, the number of the rocket engines is obtained through calculation, and the rocket engines are circumferentially distributed along the outlet end surface of the turbine engine.
9. A series-parallel-series three-power combined engine designed based on the method for designing the series-parallel three-power combined engine of claim 1, characterized in that: the device comprises an air inlet blocking cone, an engine air inlet channel, an inlet drainage channel splitter plate, a turbine engine, an outlet drainage channel splitter plate, a rocket engine, a ramjet sub-combustion chamber and a shared tail nozzle which are sequentially arranged;
the turbine engine is arranged in an air inlet channel of the turbine engine, and an ejector rocket-sub-combustion stamping outer ring combined channel is arranged outside the air inlet channel of the turbine engine in parallel;
the air inlet blocking cone is arranged at an inlet at the front end of an engine air inlet channel, the engine air inlet channel is provided with two outlets which are respectively connected with an inlet of an air inlet channel of the turbine engine and an inlet of an ejection rocket-sub-combustion stamping outer ring combined channel; the inlet drainage channel splitter plate is arranged between an inlet of an air inlet channel of the turbine engine and an inlet of the ejection rocket-sub-combustion stamping outer ring combined channel, and the outlet drainage channel splitter plate is arranged between an outlet of the air inlet channel of the turbine engine and an outlet of the ejection rocket-sub-combustion stamping outer ring combined channel;
the rocket engine is arranged behind the turbine engine and the ejector rocket-sub-combustion ramjet outer ring combined channel, the ramjet sub-combustion chamber is arranged behind the rocket engine, and the shared tail nozzle is arranged behind the ramjet sub-combustion chamber through the throat of the area-adjustable nozzle.
10. A series-parallel three-power combined engine according to claim 9, characterized in that: the inlet drainage channel flow distribution plate and the outlet drainage channel flow distribution plate are arranged in a fish scale-shaped structure which can be opened and closed, and are arranged in a multilayer crossing and overlapping mode.
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