CN103630363A - Simulation test method for high altitude ignition ability of turbine engine - Google Patents
Simulation test method for high altitude ignition ability of turbine engine Download PDFInfo
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Abstract
The invention discloses a simulation test method for high altitude ignition ability of a turbine engine. The method comprises high altitude performance test, high altitude distortion test, high altitude low-temperature test and highland starting test. A high altitude zero mach number test device of the turbine engine used by the method comprises a high altitude chamber, a calming pipe, a process air inlet channel, a turbine engine, a tail chamber, an air inlet pipe, and an air exhaust pipe, wherein the process air inlet channel is provided with a first air inlet and a first air outlet, the first air inlet and the outlet of the calming pipe are opposite and are separated, and the turbine engine is arranged in the high altitude chamber and is at the downstream side of the process air inlet channel. A distortion test device comprises a distortion simulator for the method, wherein the distortion simulator comprises a distortion simulation board and a turbine engine; the turbine engine is connected with the distortion simulator. The simulation test method for high altitude ignition ability of the turbine engine disclosed by the embodiment of the invention has the advantages of short cycle, low cost, and good practicability.
Description
Technical field
The present invention relates to engine art, relate in particular to a kind of simulation experiment method of turbine engine high-altitude ignition ability.
Background technology
Highly on the impact of turbine engine serviceability mainly from two aspects: along with highly increasing, atmospheric pressure reduces, Reynolds number raises, and surpasses critical Reynolds number rear engine component efficiencies and declines, and causes turbine engine performance to worsen or surge margin reduces; On the other hand, due to pressure decreased, be unfavorable for firing chamber tissue burning, may cause turbine engine flame-out.The aloft work performance of turbine engine self is verified and is examined by the altitude test of altitude simulation unit, but before turbine engine, install after Submerged Inlet, because Submerged Inlet has the advantages that the distortion of the low and stagnation pressure of total pressure recovery coefficient is large, charge flow rate is little, in starting process, air supply is undesirable, turbine engine fuel gas return-flow when aircraft repeatedly occurring in development starting under large Mach number, Low Angle Of Attack condition, causes the phenomenon of starting failure.
In correlation technique, study the turbine engine high-altitude ignition ability with Submerged Inlet, generally by propulsion wind tunnel or empty platform band, fly the mode of test, but this mode cycle is long, guarantee is complicated, costly.
Summary of the invention
The present invention is intended at least one of solve the problems of the technologies described above.For this reason, one object of the present invention is to propose a kind of simulation experiment method of turbine engine high-altitude ignition ability of Submerged Inlet, and the method cycle is short, and cost is low.
According to the simulation experiment method of the turbine engine high-altitude ignition ability of Submerged Inlet of the present invention, comprise the following steps: according to air intake duct blowing test test, according to number, obtain the turbine engine of described Submerged Inlet at different Mach number Ma, angle of attack, yaw angle, coefficient of flow
under inlet total pres sure recovery factor sigma and stagnation pressure degree of distortion w; According to the turbine engine working depth H of described Submerged Inlet and inlet air flow Mach number q (λ), obtain the operation interval of the turbine engine associated working of described Submerged Inlet and described Submerged Inlet; The turbine engine that obtains described Submerged Inlet according to described operation interval within the scope of operating envelope, the stagnation pressure Pt0 of the turbine engine import of described Submerged Inlet and stagnation pressure degree of distortion w, nozzle backpressure Pc under each state; According to the turbine engine inlet distortion degree w of described Submerged Inlet, generate distortion simulation plate, and calculate the degree of distortion w of described distortion simulation plate generation and the total pressure recovery coefficient of described distortion simulation plate; The degree of distortion w producing according to described distortion simulation plate and the total pressure recovery coefficient of described distortion simulation plate regulate the gas velocity of distortion simulation device outlet to determine the funtcional relationship between the distortion simulation plate exit stagnation pressure degree of distortion and analog board blocked area; According to environment temperature ts, determine the duty nh of the turbine engine of Submerged Inlet described in aberration test; According to the inlet distortion degree w of the turbine engine of described Submerged Inlet under the duty nh of the turbine engine of described Submerged Inlet and this state, determine the blocked area A of described distortion simulation plate, and regulate described distortion simulation plate according to the blocked area A of described distortion simulation plate; And after described distortion simulation plate has regulated, carry out respectively upper air performance, high-altitude aberration test, altitude low temperature test and plateau starting test.
Short according to the simulation experiment method cycle of the turbine engine high-altitude ignition ability of Submerged Inlet of the present invention, cost is low.
The degree of distortion w that distortion simulation plate produces meets following funtcional relationship: w=f(q (λ), A).
The total pressure recovery coefficient of analog board meets following funtcional relationship: σ m=g(q (λ), A); By flow field, demarcate and can determine: w=f(q (λ), A), σ m=g(q (λ), A).
Between distortion simulation plate exit stagnation pressure degree of distortion w and analog board blocked area A, meet funtcional relationship: A=f(q (λ), w).
The zero Mach number test unit for turbine engine for the simulation experiment method of the turbine engine high-altitude ignition ability of described Submerged Inlet, the described high-altitude for turbine engine zero Mach number test unit comprises: stratochamber, stratochamber comprises cabin body and hatch door, hatch door is located at pivotly on the body of cabin to open or close and on the body of Ti， cabin, cabin, is formed with cabin body import and the outlet of cabin body; Stable pipe, stable pipe is located in the body of cabin and is positioned at the draft tube installing port end of cabin body, stable pipe comprises the first pipe connected with each other and the second pipe, and the sectional area of the xsect of the first pipe is according to from cabin body import, the direction towards the outlet of cabin body increases gradually, and the second pipe is the constant cylindrical tube of sectional area; Technique air intake duct, technique air intake duct is located in stratochamber and is positioned at the downstream of stablizing pipe, technique air intake duct has the first air intake opening and first row gas port, and the outlet of the first air intake opening and stable pipe is relative and be spaced apart from each other, the central axis of technique air intake duct and the central axes of stable pipe; Turbine engine, turbine engine is located in stratochamber and is positioned at the downstream of technique air intake duct, and engine has the second air intake opening and second row gas port, and the second air intake opening is connected with first row gas port; ，Wei chamber, tail chamber is located in stratochamber and the ，Wei chamber, downstream that is positioned at engine has the 3rd air intake opening and the 3rd exhausr port, and the 3rd air intake opening is corresponding with second row gas port and be spaced apart from each other, and tail chamber has the cooling section for cooling-air; Draft tube, one end of draft tube extend in the body of cabin and with the first pipe and is connected from cabin body import; Gas outlet, one end of gas outlet extend in the body of cabin and with the 3rd exhausr port and is connected from the outlet of cabin body.
High-altitude zero Mach number test unit for turbine engine also comprises: air ejector, air ejector is located at outside stratochamber and is connected with gas outlet.
High-altitude zero Mach number test unit for turbine engine also comprises: the first support, stable pipe is located on the first support.
High-altitude zero Mach number test unit for turbine engine also comprises: the second support, and engine and technique air intake duct are located on the second support; The 3rd support, tail chamber is located on the 3rd support.
High-altitude zero Mach number test unit for turbine engine also comprises: variable valve, and variable valve is located in draft tube and is positioned at outside stratochamber; Stop valve, stop valve is located in draft tube and is positioned at the upstream side of variable valve; High-altitude zero Mach number test unit for turbine engine also comprises: for measuring the flow sensor of the flow of the air flowing out from stable pipe, flow sensor is located in a side of the contiguous technique air intake duct of stablizing pipe.
For the aberration test device of the simulation experiment method of the turbine engine high-altitude ignition ability of Submerged Inlet, aberration test device comprises: distortion simulation device, and distortion simulation device comprises distortion simulation plate; Turbine engine, turbine engine is connected with distortion simulation device.
Accompanying drawing explanation
Above-mentioned and/or the additional aspect of the present invention and advantage will become from the following description of the accompanying drawings of embodiments and obviously and easily understand, wherein,
Fig. 1 is the schematic diagram of the zero Mach number test unit for turbine engine according to an embodiment of the invention;
Fig. 2 is the schematic diagram of the zero Mach number test unit for turbine engine according to an embodiment of the invention;
Fig. 3 is that the zero Mach number test unit for turbine engine according to an embodiment of the invention completes the parameters variation diagram after experiment;
Fig. 4 is the schematic flow sheet of simulation experiment method of the turbine engine high-altitude ignition ability of Submerged Inlet according to an embodiment of the invention;
Fig. 5 is the principle schematic of simulation experiment method of the turbine engine high-altitude ignition ability of Submerged Inlet according to an embodiment of the invention;
Fig. 6 is the schematic diagram of aberration test device according to an embodiment of the invention.
Reference list:
Zero Mach number test unit 100 for turbine engine; Stratochamber 1; Cabin body 11; Cabin body import 111; Cabin body outlet 112; Hatch door 12; The second support 14; The 3rd bracket stable pipe 2; The first pipe 21; The second pipe 22; Technique air intake duct 3; The first air intake opening 31; Turbine engine 4; Second row gas port 41; Tail chamber 5; The 3rd air intake opening 51; Cooling section 53; Draft tube 6; Gas outlet 7; Variable valve 8; Stop valve 9; Air ejector 10; Stagnation temperature measuring section 201; Total pressure measurement section 202; Deng straight section 203; Distortion simulation device 204; Linkage section 205; Tail rake 207.
Embodiment
Describe embodiments of the invention below in detail, the example of described embodiment is shown in the drawings, and wherein same or similar label represents same or similar element or has the element of identical or similar functions from start to finish.Below by the embodiment being described with reference to the drawings, be exemplary, only for explaining the present invention, and can not be interpreted as limitation of the present invention.
In description of the invention, it will be appreciated that, term " " center ", " longitudinally ", " laterally ", " on ", D score, " front ", " afterwards ", " left side ", " right side ", " vertically ", " level ", " top ", " end ", " interior ", orientation or the position relationship of indications such as " outward " are based on orientation shown in the drawings or position relationship, only the present invention for convenience of description and simplified characterization, rather than device or the element of indication or hint indication must have specific orientation, with specific orientation structure and operation, therefore can not be interpreted as limitation of the present invention.In addition, term " first ", " second " be only for describing object, and can not be interpreted as indication or hint relative importance or the implicit quantity that indicates indicated technical characterictic.Thus, one or more these features can be expressed or impliedly be comprised to the feature that is limited with " first ", " second ".In description of the invention, except as otherwise noted, the implication of " a plurality of " is two or more.
In description of the invention, it should be noted that, unless otherwise clearly defined and limited, term " installation ", " being connected ", " connection " should be interpreted broadly, and for example, can be to be fixedly connected with, and can be also to removably connect, or connect integratedly; Can be mechanical connection, can be to be also electrically connected to; Can be to be directly connected, also can indirectly be connected by intermediary, can be the connection of two element internals.For the ordinary skill in the art, can concrete condition understand above-mentioned term concrete meaning in the present invention.
Below with reference to Figure of description, describe in detail according to the simulation experiment method of the turbine engine high-altitude ignition ability of Submerged Inlet of the present invention.
Simulation experiment method according to the turbine engine high-altitude ignition ability of Submerged Inlet of the present invention, with reference to figure 1 and Fig. 5, comprises the following steps:
S1: according to number, obtain the turbine engine of Submerged Inlet at different Mach number Ma, angle of attack, yaw angle, coefficient of flow according to air intake duct blowing test test
under inlet total pres sure recovery factor sigma and stagnation pressure degree of distortion w.
S2: according to the turbine engine working depth H of Submerged Inlet and inlet air flow Mach number q (λ), obtain the operation interval of the turbine engine associated working of Submerged Inlet and Submerged Inlet.
S3: the turbine engine that obtains Submerged Inlet according to operation interval within the scope of operating envelope, the stagnation pressure Pt0 of the turbine engine import of Submerged Inlet and stagnation pressure degree of distortion w, nozzle backpressure Pc under each state.
S4: according to the turbine engine inlet distortion degree w of Submerged Inlet, generate distortion simulation plate, and calculate the degree of distortion w of distortion simulation plate generation and the total pressure recovery coefficient σ of distortion simulation plate.
S5: the degree of distortion w producing according to distortion simulation plate and the total pressure recovery coefficient of distortion simulation plate regulate the gas velocity of distortion simulation device outlet to determine the funtcional relationship between the distortion simulation plate exit stagnation pressure degree of distortion and analog board blocked area.
S6: according to environment temperature ts, determine the duty nh of the turbine engine of Submerged Inlet in aberration test.
S7: according to the inlet distortion degree w of the turbine engine of Submerged Inlet under the duty nh of the turbine engine of Submerged Inlet and this state, determine the blocked area A of distortion simulation plate, and regulate distortion simulation plate according to the blocked area A of distortion simulation plate.
S8: after distortion simulation plate has regulated, carry out respectively upper air performance, high-altitude aberration test, altitude low temperature test and plateau starting test.
It should be noted that, illustrating of above-mentioned steps S1-S8 is not sequentially the enforcement order of the necessary adhere rigidly to of step S1-S8, and step S1-S8 can carry out not according to above-mentioned shown order.
Wherein, in step S4, the degree of distortion w that distortion simulation plate produces meets following funtcional relationship:
w=f（q(λ)，A）--------------------（1）
Wherein, w is the degree of distortion that distortion simulation plate produces;
Q (λ) is the inlet air flow Mach number of the turbine engine of Submerged Inlet;
A is the blocked area A of distortion simulation plate.
In step S5, distortion simulation plate also can produce certain pressure loss, and the total pressure recovery coefficient σ of distortion simulation plate meets following funtcional relationship:
σm=g（q(λ)，A）----------------------（2）
Wherein, σ m is the total pressure recovery coefficient of distortion simulation plate;
Q (λ) is the inlet air flow Mach number of the turbine engine of Submerged Inlet;
A is the blocked area A of distortion simulation plate;
Above-mentioned funtcional relationship (1) (2) can be demarcated and be obtained by flow field.That is to say, can determine funtcional relationship (1) (2) by distortion simulation device nominal data.
In step S7, between distortion simulation plate exit stagnation pressure degree of distortion w and analog board blocked area A, meet funtcional relationship:
A=f（q(λ)，w）--------------------------（3）
Wherein, q (λ) is the inlet air flow Mach number of the turbine engine of Submerged Inlet;
A is the blocked area A of distortion simulation plate;
W is the inlet distortion degree of turbine engine.
By upper air performance, study the ability to work in engine high-altitude, this test can complete at altitude simulation unit; According to the operating envelope of aircraft, in test, simulate working depth envelope curve and the flight Mach number envelope curve of turbine engine, i.e. definite engine intake stagnation pressure Pt0 and nozzle backpressure Pc.
By the high-altitude inlet engine compatibility performance after Submerged Inlet being installed before the anti-aberration test research in high-altitude turbine engine.In test, according to flying height H, Mach number Ma, environment temperature ts, determine the duty nh of aberration test engine; According to engine inlet distortion degree w under engine behavior and this state, determine the blocked area A of distortion simulation plate; According to the total pressure recovery coefficient of analog board and flying height H, Mach number Ma, determine the air-flow stagnation pressure Pt0/ σ m before analog board.
By high-altitude cold-start test, the starting capability of research engine under the cryogenic conditions of high-altitude.In test, engine is placed in cryogenic box, after engine turbine, temperature reaches after minimum temperature requires and is incubated 2h until equalized temperature, then simulated altitude Pc, incoming-flow pressure Pt0, incoming flow low temperature ts carry out test, and in test, height, temperature, Mach number should be chosen the harshest envelope curve.
Plateau starting test is that engine intake is verified without the starting under speed punching press condition, can study the rationality of engine startup control law under the low total pressure recovery coefficient condition of Submerged Inlet by this test.Test height should be the harshest height envelope curve of engine operation, and incoming flow Mach number should guarantee to be not more than 0.1; In test should not there is not the phenomenons such as suspension, delay in engine.
In a specific embodiment of invention, according to air intake duct blowing data determined in engine operation envelope curve flying height H be 6km and flight Mach number Ma be the air intake port stagnation pressure degree of distortion w of 0.7 o'clock be 8.5% and inlet total pres sure recovery factor sigma be 0.93.Submerged Inlet and Engine Matching work are calculated and obtained engine intake air-flow stagnation pressure Pt0 is that 61.4kPa and stagnation pressure degree of distortion w are 8.5%, nozzle backpressure is 47.2kPa.
According to the stagnation pressure degree of distortion w of engine intake, be 8.5%, developed the pressure distortion analog board of certain blocked area.By demarcation, the blocked area A of analog board and the funtcional relationship between stagnation pressure distortion w and total pressure recovery coefficient σ m have been obtained.
According to environment temperature, be-23.9 ℃, determine engine behavior rotation speed n h=nh* * SQRT(249.2/288.15), the stagnation pressure degree of distortion that generation value is 8.5% under this duty needs the blocked area A of analog board.
Then can carry out engine is that 6km, flight Mach number Ma are 0.5～0.8 upper air performance at flying height H, and in test, nozzle backpressure is modeled as 47.2kPa, and during Mach 2 ship 0.7, engine intake stagnation pressure is 61.4kPa.
Carrying out flying height H is that 6km, flight Mach number Ma are 0.5 the anti-aberration test in high-altitude, and distortion simulation plate is installed before engine in test, and the area that gets lodged in of analog board is A, and nozzle backpressure is 47.2kPa, and before analog board, pressure is (56.4/0.899) kPa.
The altitude low temperature test that carry out flying height H and be 6km, flight Mach number Ma and be 0.5, the temperature of turbine engine body is-45 ℃, intake air temperature-50 ℃.After the turbine of turbine engine, temperature reaches after-45 ℃, and 2h is until equalized temperature in insulation, at simulation nozzle backpressure Pc, is then that 47.2kPa, incoming-flow pressure are that 56.4kPa, incoming flow low temperature are tested during for-50 ℃.
Carrying out flying height H is the test of 6km plateau starting.This experiment can be carried out at the zero Mach number test unit for turbine engine.
The turbine engine of Submerged Inlet, after above-mentioned development test step, shows that through flight test examination these effect tests are true, effective.
The simulation experiment method that can be used for the turbine engine high-altitude ignition ability of Submerged Inlet for the zero Mach number test unit of turbine engine, especially can be used in plateau starting test.
Below with reference to Figure of description, describe according to the high-altitude for turbine engine of the embodiment of the present invention zero Mach number test unit 100.
According to the zero Mach number test unit 100 for turbine engine of the embodiment of the present invention, comprise stratochamber 1, stable pipe 2, technique air intake duct 3, turbine engine 4, tail chamber 5, draft tube 6 and gas outlet 7.
Wherein, stratochamber 1 comprises cabin body 11 and hatch door 12, and hatch door 12 is located on cabin body 11 pivotly to open or close cabin body 11, is formed with cabin body import 111 and cabin body outlet 112 on cabin body 11.One end of draft tube 6 extend in cabin body 11 from cabin body import 111.
One end of draft tube 6 extend in cabin body 11 and is connected with the first pipe 21 from cabin body import 111.First the air entering in cabin body 11 from draft tube 6 enters stable pipe 2.Be understandable that, in test, air can penetrate with lower speed from stable pipe 2, and for example the Mach number of air inlet is not more than 0.05, i.e. Ma ≮ 0.05, and wherein the implication of symbol " ≮ " is " being not more than ".
Technique air intake duct 3 is located in stratochamber 1 and is positioned at the downstream of stable pipe 2.Downstream refers in the downstream of cabin body 11 interior air-flow directions, and for example the Ji Shi downstream, right side in Fig. 1 and Fig. 2, that is to say, in Fig. 1 and Fig. 2, technique air intake duct 3 is located at the right side of stable pipe 2." downstream " hereinafter occurring again, if do not make specified otherwise, all by this understanding.
Technique air intake duct 3 has the first air intake opening 31 and first row gas port, and the outlet of the first air intake opening 31 and stable pipe 2 is relative and be spaced apart from each other.Parts for the air of stable pipe 2 stable outputs are entered technique air intake duct 3 and are entered turbine engine 4 by the first air intake opening 31, and all the other most of air-flows are from technique air intake duct 3 outer flow mistakes, and are gone out stratochamber 1 by 5 suctions of tail chamber.First row gas port is connected with turbine engine 4, and the air in technique air intake duct 3 flows in turbine engine 4 through first row gas port.Allow like this turbine engine 4 natural aspiration in starting process, and stable pipe 2 stream pressures of discharging and flow velocity stablize constantly, so just can realize the steady of intake simulation Mach number.
The central axis of technique air intake duct 3 and the central axes of stablizing pipe 2.Thereby the air that can guarantee better like this pressure and flow speed stability flows into technique air intake duct 3 and enters turbine engine 4.
Tail chamber 5 is located in stratochamber 1 and is positioned at the downstream of turbine engine 4, the namely right side in Fig. 1.
Tail chamber 5 has the 3rd air intake opening 51 and the 3rd exhausr port, and the 3rd air intake opening 51 is corresponding with second row gas port 41 and be spaced apart from each other.The air that turbine engine 4 is discharged enters tail chamber 5 by the 3rd air intake opening 51.
Tail chamber 5 has the afterbody that is positioned at cooling section 53 for cooling section 53, the three exhausr ports of cooling-air.The high-temperature gas that turbine engine 4 is discharged is cooled and discharges from the 3rd exhausr port at cooling section 53.
One end of draft tube 6 extend in cabin body 11 and is connected with the first pipe 21 from cabin body import 111, and the other end of draft tube 6 is connected with source of the gas, and draft tube 6 imports to the air in source of the gas in cabin body 11.
One end of gas outlet 7 extend in cabin body 11 and is connected with the 3rd exhausr port from cabin body outlet 112.That is to say, gas outlet 7 is connected with cooling section 53, and the cooled air of cooling section 53 is discharged outside cabin body 11 by gas outlet 7.
At the trial, first the pressure of adjusting in stratochamber 1 arrives certain numerical value, for example the pressure in stratochamber 1 is 60.2kPa, draft tube 6 has the air of certain pressure to the interior input of stratochamber 1 simultaneously, air becomes steady after stablizing pipe 2, regulate air mass flow, for example, air mass flow can be 4kg/s.After stratochamber 1 intake and exhaust balance, what the interior formation of stratochamber 1 was stable manages 2 gas channels to tail chamber 5 from stablizing, and part air-flow flows into turbine engines 4 by technique air intake duct 3, and all the other most of air-flows are from turbine engine 4 outer flow mistakes.At this moment the windmill rotating speed of turbine engine 4 is zero, and air inlet Mach number is about 0.03, amounts to wind speed and is about 36km/h.This is concerning turbine engine 4, and air-flow punching press effect does not almost have, and is equivalent to without the starting under wind friction velocity, is equivalent to zero Mach number starting.Turbine engine 4 just can complete heat run content of the test according to existing procedure like this.After turbine engine 4 stops, continue air feed until after turbine engine 4 temperature reductions, stop air feed, off-test.
According to the zero Mach number test unit 100 for turbine engine of the embodiment of the present invention, pass through technique air intake duct 3 and stable pipe 2 spaced apart, allow turbine engine 4 natural aspiration in starting process, thereby realize the steady of intake simulation Mach number, then can be by regulating pressure and the air mass flow in stratochamber 1, can realize the experiment condition for turbine engine 4 zero Mach numbers, thereby can low cost realize fast turbine engine 4 at the state of ground zero Mach number starting certification test, practicality is good.
According to the zero Mach number test unit 100 for turbine engine of the embodiment of the present invention, also comprise air ejector 10.Air ejector 10 is located at outside stratochamber 1 and is connected with gas outlet 7 with the air in suction cabin body 11, and the pressure in adjusting cabin body 11, makes the pressure in stratochamber 1 reach the pressure needing.
In addition, under the state that pressure in stratochamber 1 and air mass flow all regulate, turbine engine 4 ignition start are to slow train operating mode, in turbine engine 4 starting process, due to turbine engine 4 exhausts from ejector action, can reduce the pressure of stratochamber 1, make the pressure in stratochamber 1 produce fluctuation, regulate air ejector 10 can stablize the pressure of stratochamber 1.
According to the zero Mach number test unit 100 for turbine engine of the embodiment of the present invention, also comprise the first support.Stable pipe 2 is located on the first support.Can guarantee like this stability of stable pipe 2, improve stability and the reliability of experiment.
According to the zero Mach number test unit 100 for turbine engine of the embodiment of the present invention, also comprise the second support 14 and the 3rd support.
Wherein, turbine engine 4 and technique air intake duct 3 are located on the second support 14.Tail chamber 5 is located on the 3rd support.
The second support 14 is provided with hanger bracket, and turbine engine 4 can be suspended on hanger bracket.
On the second support 14, can be provided with pressure transducer, this pressure transducer, for measuring the pressure in stratochamber 1, to regulate as required the pressure in stratochamber 1, reaches the requirement of experiment condition.
According to the zero Mach number test unit 100 for turbine engine of the embodiment of the present invention, also comprise variable valve 8 and stop valve 9.
Wherein variable valve 8 is located in draft tube 6 and is positioned at outside stratochamber 1.Variable valve 8 can regulate the flow of the air coming from source of the gas to make the air mass flow in stratochamber 1 reach turbine engine 4 desired air mass flow under zero Mach number experiment condition.
Stop valve 9 is located in draft tube 6 and is positioned at the upstream side of variable valve 8.That is to say, stop valve 9 is located between middle pressurized air source and variable valve 8, can close draft tube 6 and open draft tube 6, can regulate better charge flow rate like this, avoids the frequent switching of variable valve 8, extends the serviceable life of variable valve 8.
According to the zero Mach number test unit 100 for turbine engine of the embodiment of the present invention, also comprise that flow sensor is located in a side of stablizing the contiguous technique air intake duct 3 of managing 2 to measure air mass flow for measuring the flow sensor of the flow of the air flowing out from stable pipe 2.
In one embodiment of the invention, according to the zero Mach number test unit 100 for turbine engine of the embodiment of the present invention, also comprise controller, controller can be connected with variable valve 8 with flow sensor, flow sensor transmits flow signal to controller, controller can regulating and controlling valve 8, regulates the flow of the air that therefrom pressurized air source comes to make air mass flow reach the desired flow of turbine engine 4 air inlet.
According to the zero Mach number test unit 100 for turbine engine of the embodiment of the present invention, also comprise inlet pressure sensor.Inlet pressure sensor can be a plurality of, and one of them inlet pressure sensor can be located near the first air intake opening 31 of technique air intake duct 3, for measuring turbine engine 4 air intake duct wall static pressures.Also have at least three inlet pressure sensors to be uniformly distributed along air inlet cross-section radial, for measuring turbine engine 4 air inlet stagnation pressures.
According to the zero Mach number test unit 100 for turbine engine of the embodiment of the present invention, also comprise for measuring the speed probe of turbine engine 4 rotating speeds.
Below the simple process of the test of describing zero Mach number test unit 100 in accordance with a preferred embodiment of the present invention.
First, open air ejector 10, the pressure in adjustment stratochamber 1, to 60.2kPa, is opened stop valve 9 simultaneously, and regulating and controlling valve 8 makes draft tube 6 in stratochamber 1, be blown into the air that air mass flow is 4kg/s by stable pipe 2.After stratochamber 1 intake and exhaust balance, what the interior formation of stratochamber 1 was stable manages 2 gas channels to tail chamber 5 from stablizing, and part air-flow flows into turbine engines 4 by technique air intake duct 3, and all the other most of air-flows are from turbine engine 4 outer flow mistakes.At this moment the windmill rotating speed of turbine engine 4 is zero, and air inlet Mach number is about 0.03, amounts to wind speed and is about 36km/h.This does not almost have turbine engine 4 air-flow punching press effects, therefore for turbine engine 4, is equivalent to without the starting under wind friction velocity.The state simulation of zero Mach number well after turbine engine 4 ignition start to slow train operating mode, turbine engine 4 starting process due to exhaust from ejector action, can reduce the pressure in stratochamber 1, need to be by the pressure surge that regulates air ejector 10 to stablize stratochamber 1.After the pressure stability of stratochamber 1, turbine engine 4 has continued other heat run contents of the test according to intended flow.After turbine engine 4 stops, continue air feed after turbine engine 4 temperature reduce, stop air feed off-test.
By above-mentioned experiment, can draw following test findings, its curve map as shown in Figure 3:
Wherein, Pm1 is engine inlets wall static pressure; PC1 is stratochamber 1 environmental pressure; N is engine speed; Ptm1, Ptm2, Ptm3 are the engine charge stagnation pressure along equally distributed three points of air inlet cross-section radial; T0 is the 4 starting moment of turbine engine, and T1 is a certain moment after turbine engine 4 startings.
Inlet pressure sensor can measure Pm1 value, namely engine inlets wall static pressure; Pressure transducer on the second support 14 can measure PC1 value, namely stratochamber 1 environmental pressure; Speed probe can measure n value, namely engine speed; At least three can measure respectively Ptm1, Ptm2, Ptm3 along the equally distributed inlet pressure sensor of air inlet cross-section radial, namely along the engine charge stagnation pressure of equally distributed three points of air inlet cross-section radial.
With reference to figure 3, turbine engine 4 prestarts, turbine engine 4 air inlet stagnation pressures and stratochamber 1 internal pressure are more steady, and force value is basic identical, illustrate that turbine engine 4 incoming flow flow velocitys are lower, and air-flow is without obvious punching press effect.Turbine engine 4 startings constantly, air inlet incoming flow is had to certain interference, but absolute figure are very little, from the parameters of test findings simulation, meet test simulation requirement.
Thus, by zero according to an embodiment of the invention Mach number test unit 100, can simulate the intake and exhaust condition of turbine engine 4 under zero Mach number of high-altitude, thereby can simulate turbine engine 4 for the aircraft starting performance under this condition, turbine engine 4 has obtained optimization for the startup control law under this condition, can realize the test objective of expection.Meanwhile, adopt zero according to an embodiment of the invention Mach number test unit 100 simulation test costs low, practicality is good.
Below with reference to Fig. 6, describe in detail according to the aberration test device of the simulation experiment method of the turbine engine high-altitude ignition ability for Submerged Inlet of the present invention, the anti-aberration test in high-altitude can complete in this aberration test device.
This aberration test device comprises: distortion simulation device 204 and turbine engine 4.
Wherein, distortion simulation device 204 comprises distortion simulation plate, and distortion simulation plate can need to regulate blocked area A according to experiment.
This aberration test device also can comprise stagnation temperature measuring section 201, total pressure measurement section 202, etc. straight section 203, linkage section 205 and tail rake 207.Wherein, environment temperature can be measured at stagnation temperature measuring section 201, and total static pressure can be measured in total pressure measurement section 202.Turbine engine 4 is located between linkage section 205 and tail rake 207, distortion simulation device 204 be located at linkage section 205 and etc. between straight section 203.
In the description of this instructions, the description of reference term " embodiment ", " some embodiment ", " illustrative examples ", " example ", " concrete example " or " some examples " etc. means to be contained at least one embodiment of the present invention or example in conjunction with specific features, structure, material or the feature of this embodiment or example description.In this manual, the schematic statement of above-mentioned term is not necessarily referred to identical embodiment or example.And the specific features of description, structure, material or feature can be with suitable mode combinations in any one or more embodiment or example.
Although illustrated and described embodiments of the invention, those having ordinary skill in the art will appreciate that: in the situation that not departing from principle of the present invention and aim, can carry out multiple variation, modification, replacement and modification to these embodiment, scope of the present invention is limited by claim and equivalent thereof.
Claims (10)
1. a simulation experiment method for the turbine engine high-altitude ignition ability of Submerged Inlet, is characterized in that, comprises the following steps:
According to air intake duct blowing test test, according to number, obtain the turbine engine of described Submerged Inlet at different Mach number Ma, angle of attack, yaw angle, coefficient of flow
under inlet total pres sure recovery factor sigma and stagnation pressure degree of distortion w;
According to the turbine engine working depth H of described Submerged Inlet and inlet air flow Mach number q (λ), obtain the operation interval of the turbine engine associated working of described Submerged Inlet and described Submerged Inlet;
The turbine engine that obtains described Submerged Inlet according to described operation interval within the scope of operating envelope, the stagnation pressure Pt0 of the turbine engine import of described Submerged Inlet and stagnation pressure degree of distortion w, nozzle backpressure Pc under each state;
According to the turbine engine inlet distortion degree w of described Submerged Inlet, generate distortion simulation plate, and calculate the degree of distortion w of described distortion simulation plate generation and the total pressure recovery coefficient of described distortion simulation plate;
The degree of distortion w producing according to described distortion simulation plate and the total pressure recovery coefficient of described distortion simulation plate regulate the gas velocity of distortion simulation device outlet to determine the funtcional relationship between the distortion simulation plate exit stagnation pressure degree of distortion and analog board blocked area;
According to environment temperature ts, determine the duty nh of the turbine engine of Submerged Inlet described in aberration test;
According to the inlet distortion degree w of the turbine engine of described Submerged Inlet under the duty nh of the turbine engine of described Submerged Inlet and this state, determine the blocked area A of described distortion simulation plate, and regulate described distortion simulation plate according to the blocked area A of described distortion simulation plate;
After described distortion simulation plate has regulated, carry out respectively upper air performance, high-altitude aberration test, altitude low temperature test and plateau starting test.
2. the simulation experiment method of the turbine engine high-altitude ignition ability of Submerged Inlet according to claim 1, is characterized in that, the degree of distortion w that distortion simulation plate produces meets following funtcional relationship:
w=f（q(λ)，A）。
3. the simulation experiment method of the turbine engine high-altitude ignition ability of Submerged Inlet according to claim 1 and 2, is characterized in that, the total pressure recovery coefficient σ of analog board meets following funtcional relationship:
σm=g（q(λ)，A）；
By flow field, demarcate and determine w=f(q (λ), A), σ m=g(q (λ), A).
4. according to the simulation experiment method of the turbine engine high-altitude ignition ability of the Submerged Inlet described in any one in claim 1-3, it is characterized in that, between distortion simulation plate exit stagnation pressure degree of distortion w and analog board blocked area A, meet funtcional relationship:
A=f（q(λ)，w）。
5. according to the simulation experiment method of the turbine engine high-altitude ignition ability of the Submerged Inlet described in claim 1-4 any one, it is characterized in that, for the zero Mach number test unit for turbine engine of the simulation experiment method of the turbine engine high-altitude ignition ability of described Submerged Inlet, the zero Mach number test unit in the described high-altitude for turbine engine comprises:
Stratochamber, described stratochamber comprises cabin body and hatch door, described hatch door is located on the body of described cabin pivotly to open or close described cabin body, is formed with cabin body import and the outlet of cabin body on the body of described cabin;
Stable pipe, described stable pipe is located in the body of described cabin and is positioned at the cabin body entrance point of described cabin body, described stable pipe comprises the first pipe connected with each other and the second pipe, the sectional area of the xsect of described the first pipe is according to from the body import of described cabin, the direction towards the outlet of described cabin body increases gradually, and described the second pipe is the constant cylindrical tube of sectional area;
Technique air intake duct, described technique air intake duct is located in described stratochamber and is positioned at the downstream of described stable pipe, described technique air intake duct has the first air intake opening and first row gas port, the outlet of described the first air intake opening and described stable pipe is relative and be spaced apart from each other, the central axis of described technique air intake duct and the central axes of described stable pipe;
Turbine engine, described turbine engine is located in described stratochamber and is positioned at the downstream of described technique air intake duct, and described engine has the second air intake opening and second row gas port, and described the second air intake opening is connected with described first row gas port;
Tail chamber, described tail chamber is located in described stratochamber and is positioned at the downstream of described engine, described tail chamber has the 3rd air intake opening and the 3rd exhausr port, and described the 3rd air intake opening is corresponding with described second row gas port and be spaced apart from each other, and described tail chamber has the cooling section for cooling-air;
Draft tube, one end of described draft tube extend in the body of described cabin and with described the first pipe and is connected from the body import of described cabin;
Gas outlet, one end of described gas outlet extend in the body of described cabin and with described the 3rd exhausr port and is connected from the outlet of described cabin body.
6. the simulation experiment method of the turbine engine high-altitude ignition ability of Submerged Inlet according to claim 5, it is characterized in that, the described high-altitude for turbine engine zero Mach number test unit also comprises: air ejector, described air ejector is located at outside described stratochamber and is connected with described gas outlet.
7. according to the simulation experiment method of the turbine engine high-altitude ignition ability of the Submerged Inlet described in any one in claim 5-6, it is characterized in that, the described high-altitude for turbine engine zero Mach number test unit also comprises: the first support, described stable pipe is located on described the first support.
8. according to the simulation experiment method of the turbine engine high-altitude ignition ability of the Submerged Inlet described in any one in claim 5-7, it is characterized in that, the described high-altitude for turbine engine zero Mach number test unit also comprises:
The second support, described engine and described technique air intake duct are located on described the second support;
The 3rd support, described tail chamber is located on described the 3rd support.
9. according to the simulation experiment method of the turbine engine high-altitude ignition ability of the Submerged Inlet described in any one in claim 5-8, it is characterized in that, the described high-altitude for turbine engine zero Mach number test unit also comprises:
Variable valve, described variable valve is located in described draft tube and is positioned at outside described stratochamber;
Stop valve, described stop valve is located in described draft tube and is positioned at the upstream side of described variable valve;
For measuring the flow sensor of the flow of the air flowing out from described stable pipe, described flow sensor is located in the side of the described technique air intake duct of vicinity of described stable pipe.
10. according to the simulation experiment method of the turbine engine high-altitude ignition ability of the Submerged Inlet described in any one in claim 1-4, it is characterized in that, for the aberration test device of the simulation experiment method of the turbine engine high-altitude ignition ability of described Submerged Inlet, described aberration test device comprises:
Distortion simulation device, described distortion simulation device comprises distortion simulation plate;
Turbine engine, described turbine engine is connected with described distortion simulation device.
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