CN115169056A

CN115169056A - Unsteady state performance estimation method for sub-combustion ramjet engine

Info

Publication number: CN115169056A
Application number: CN202210962184.4A
Authority: CN
Inventors: 陈玉春; 张至斌; 宋可染; 黄新春
Original assignee: Northwestern Polytechnical University; Beijing Power Machinery Institute
Current assignee: Northwestern Polytechnical University; Beijing Power Machinery Institute
Priority date: 2022-08-11
Filing date: 2022-08-11
Publication date: 2022-10-11

Abstract

The invention relates to a method for estimating unsteady state performance of a scramjet engine, which comprises the steps of calculating and determining design point parameters of an air inlet channel, a transition section, a combustion chamber and a tail nozzle of the ramjet engine; calculating and determining non-design points of an air inlet channel, a transition section, a combustion chamber and a tail nozzle of the ramjet according to the design point parameters and the flow balance relation of the air inlet channel and the tail nozzle; establishing an actual working point beta residual error equation of the air inlet channel; iterating the actual working point beta of the air inlet channel point by point in the whole unsteady state process to obtain the actual thrust Fn of the engine _{Fruit of Chinese wolfberry} And finally obtaining an unsteady thrust curve of the engine considering the volume effect. The method can accurately and quickly calculate the thrust under the unsteady state condition of the sub-combustion ramjet when designing the unsteady state performance of the sub-combustion ramjet, can shorten the design iteration cycle of a general designer when designing the general sub-combustion ramjet, and guides a part designer to carry out detailed design.

Description

Unsteady state performance estimation method for sub-combustion ramjet engine

Technical Field

The invention relates to the field of aero-engines, in particular to a method for estimating unsteady state performance of a sub-combustion ramjet engine.

Background

The ramjet engine uses oxygen in the air as an oxidant, and the carried fuel contains little or no oxidant, so that the specific impulse is obviously improved. Therefore, the missile powered by the ramjet has a longer range, lighter weight and better maneuverability, and can realize the whole-course supersonic cruise flight, thereby greatly improving the penetration resistance of the missile. Since the eighties of the last century, the research on the ram propulsion technology has been carried out in almost all countries with the missile development ability, and various ram and combined engines thereof become the preferred power devices for tactical missiles, interception missiles and cruise missiles in the present century.

The performance parameters of a ramjet engine are critical to the overall design and determine how advanced the engine is. The unstable-state process such as large maneuvering and the like can be carried out during the work of the ramjet, the accurate design of the unstable-state performance of the scramjet can help general designers to quickly judge whether the unstable-state performance parameters of the ramjet meet the requirements or not, the iteration cycle of the general design is shortened, and the design efficiency is improved.

Disclosure of Invention

The invention aims to avoid the defects of the prior art and provides the unsteady state performance estimation method of the sub-combustion ramjet engine, which can quickly and accurately estimate the unsteady state performance of the engine and is convenient for optimizing and iterating the overall performance of the engine when the overall scheme design of the ramjet engine is carried out.

In order to achieve the purpose, the invention adopts the technical scheme that: a method for estimating unsteady state performance of a sub-combustion ramjet engine comprises the following steps:

step one, according to a given flight Mach number Ma, a flight altitude H, inlet flow Wa of an air inlet channel and an interpolated flow coefficient of the flight Mach number Ma on an interpolation characteristic diagram of the air inlet channel

And a total pressure recovery coefficient sigma, and sequentially calculating and determining design points of an air inlet channel, a transition section, a combustion chamber and a tail nozzle of the ramjet;

the determined inlet design point parameters include: frontal area A of air inlet _c Total temperature T of outlet of air inlet passage _out And the total pressure P of the outlet of the air inlet _out (ii) a The determined design point parameters for the transition section include: total pressure P at transition section outlet ₂₂ Total outlet temperature T of transition section ₂₂ And the converted flow Wac of the inlet of the transition section _2des (ii) a The determined combustor design point parameters include: combustion chamber exit area A ₃₂ Total pressure P at the outlet of the combustion chamber ₃₂ Outlet flow Wg of combustion chamber ₃₂ (ii) a The determined jet nozzle design point parameters include: throat area A of the exhaust nozzle ₄₃ And jet nozzle exit area A ₄₂ ；

Secondly, combining parameters of the ramjet air inlet channel, the transition section, the combustion chamber and the tail nozzle design point and the flow balance relation of the air inlet channel and the tail nozzle, and capturing the flow Wa _ c and the flight Mach number Ma of the air inlet channel according to a disclosed standard atmosphere table and a non-design point _off Flying height H _off Calculating the actual inlet total temperature T _{in excess} And inlet total pressure P of air inlet _{in Shi} According to the actual air inlet channel working point beta and the flight Mach number Ma of the non-design point _off Interpolation flow coefficient on air inlet channel interpolation characteristic diagram

And total pressure recovery coefficient sigma ₁ Calculating and determining non-design points of an air inlet channel, a transition section, a combustion chamber and a tail nozzle of the ramjet;

the determined air inlet channel non-design point parameters comprise: intake passage flow Wa at non-design point _{Fruit of Chinese wolfberry} Total temperature T of outlet of air inlet channel _{out of fact} And the total pressure P of the outlet of the air inlet _{out of fact} (ii) a The determined non-design point parameters of the transition section comprise: actual total pressure P at outlet of transition section _{22 fruit} Actual total outlet temperature T of transition section _{22 fruit} Actual transition section outlet flow Wa _{22 fruit} (ii) a The determined combustion chamber non-design point parameters include: actual total outlet temperature T of combustion chamber _{32 fruit} Actual total pressure P at outlet of combustion chamber _{32 fruit} Actual outlet flow Wg of the combustor _{32 fruit} (ii) a What is neededThe determined non-design point parameters of the jet nozzle comprise: actual thrust Fn of engine _{Fruit of Chinese wolfberry} And the total pressure P required by the actual tail nozzle ₈ ；

Step three, obtaining the total pressure P required by the actual tail nozzle according to the step two ₈ And the known actual total pressure P of the tail nozzle inlet _{41 fruit} Forming a residual equation as shown in the following formula:

ERR＝(P _{41 essence} -P ₈ )/P _{41 fruit}

Changing the value of a residual error equation by changing the actual working point beta of the air inlet channel, iterating the actual working point beta of the air inlet channel by using a Newton method, and when the absolute value of the residual error is less than 0.0001, determining that iteration is converged;

step four, according to the residual air coefficient alpha and the flight Mach number Ma of the non-design point _off And a flying height H _off Iterating the actual working point beta of the air inlet channel point by point according to the change rule of the Time along with the step three, thereby obtaining the actual thrust Fn of the engine calculated in the step two _{Fruit of Chinese wolfberry} Until Time =0 to a given Time limit Time _MAX And finally obtaining the unsteady thrust curve of the engine considering the volume effect after the calculation of the whole unsteady process is finished.

Further, the step one specifically includes the following steps:

step 11: calculating the design point of the air inlet channel: given flight Mach number Ma, flight altitude H, inlet flow Wa of air inlet channel, and interpolated flow coefficient of flight Mach number Ma on inlet channel interpolation characteristic diagram

And a total pressure recovery coefficient σ;

firstly, calculating the total inlet temperature T of the air inlet according to a standard atmosphere meter, the flight Mach number Ma and the flight altitude H _in And inlet total pressure P of air inlet _in Total inlet duct outlet temperature T _out Equal to the total inlet temperature T of the air inlet _in ；

And according to the formula:

P _out ＝σP _in

is calculated to obtainTotal pressure P at outlet of air inlet _out ；

Secondly, the given parameters are substituted into a flow continuous equation to obtain:

Wa＝ρV ₀ A

calculating the windward area A of the air inlet _c In the formula, K is 0.0404 and is a fixed value; q (λ) is a flow function, ρ is the air density, V ₀ The incoming flow velocity is, and A is the flow tube area;

that is, the obtained intake passage design point parameters include: frontal area A of air inlet _c Total temperature T of outlet of air inlet channel _out And the total pressure P of the outlet of the air inlet _out ；

Step 12: calculating a design point of the total pressure loss of the transition section:

the total pressure P of the inlet of the known transition section ₂₁ Equal to the total pressure P of the outlet of the air inlet passage _out Total temperature T at the inlet of transition section ₂₁ Equal to the total temperature T of the outlet of the air inlet _out Total outlet temperature T of transition section ₂₂ Equal to the total temperature T of the inlet of the transition section ₂₁ (ii) a Transition section inlet flow Wa ₂ Equal to the inlet duct flow Wa and the transition section outlet flow Wa ₂₂ Equal to the inlet flow Wa of the transition section ₂₁ (ii) a And giving a total pressure recovery coefficient sigma of the transition section _2des ，

Thereby calculating the total pressure P of the outlet of the transition section ₂₂ From said transition section inlet total pressure P ₂₁ And the total pressure recovery coefficient sigma of the transition section _2des Calculating total pressure P of outlet of transition section ₂₂ The process of (2) represents total pressure loss;

continuously adopting known parameters, and calculating the converted flow Wac of the inlet of the transition section according to a formula _2des ，

That is, the obtained transition design point parameters include: total pressure P at outlet of transition section ₂₂ Total outlet temperature T of transition section ₂₂ And converted flow Wac at the inlet of the transition section _2des ；

Step 13: calculating the design point of the combustion chamber:

the known outlet section of the transition section is the inlet section of the combustion chamber, and the total inlet temperature T of the combustion chamber ₃₁ Total pressure P at the inlet of the combustion chamber ₃₁ And the cold total pressure recovery coefficient sigma of the combustion chamber _3des And inlet flow Wa of combustion chamber ₃₁ Mach number Ma of combustion chamber inlet ₃₁ And the combustion chamber outlet temperature T ₃₂ ；

Substituting the known parameters into the flow continuous equation to obtain:

Wa ₃₁ ＝ρ ₃ v ₃ A ₃₁

thereby calculating the inlet area A of the combustion chamber ₃₁ In the formula, ρ ₃ Is the combustion chamber air density, v ₃ Is the combustion chamber incoming flow velocity;

then, the combustion efficiency η is given _b In combination with said combustion chamber outlet temperature T ₃₂ To determine the combustion chamber fuel flow W _fb According to the fuel flow W of the combustion chamber _fb Calculating the outlet flow Wg of the combustion chamber ₃₂ ；

Continuously adopting known parameters and calculating the converted flow Wac of the inlet of the combustion chamber according to a formula _3des ：

Since the combustion chamber is an equal straight path, the outlet area A of the combustion chamber ₃₂ Equal to combustionArea of chamber entrance A ₃₁ (ii) a According to the total pressure P of the inlet of the combustion chamber ₃₁ And the cold total pressure recovery coefficient sigma of the combustion chamber _3des ，

Further according to the flow conservation equation, the inlet area A of the combustion chamber ₃₁ Inlet flow Wa of combustion chamber ₃₁ Total pressure at the inlet of the combustion chamber P ₃₁ Total inlet temperature T of combustion chamber ₃₁ Determining the velocity V of the combustion chamber inlet air flow ₃₁ (ii) a From the combustion chamber outlet area A ₃₂ Outlet flow Wg of combustion chamber ₃₂ Total pressure P at the outlet of the combustion chamber ₃₂ Total outlet temperature T of combustion chamber ₃₂ Determining the velocity V of the gas flow at the outlet of the combustion chamber ₃₂ (ii) a From the combustion chamber inlet area A ₃₁ Inlet flow Wa of combustion chamber ₃₁ Inlet total pressure P of combustion chamber ₃₁ Total inlet temperature T of combustion chamber ₃₁ Calculating the static pressure Ps at the inlet of the combustion chamber ₃₁ ；

Then, calculating the thermal state loss according to momentum conservation, substituting the known parameters into a momentum conservation formula to obtain:

Wa ₃ V ₃₁ +Ps ₃₁ A ₃₁ ＝Wg ₃₂ V ₃₂ +Ps ₃₂ A ₃₂

in the formula, the left side of the equal sign is impulse considering cold state loss, the right side is impulse after heating, and total pressure P of an outlet of the combustion chamber is obtained by iteration by taking a momentum conservation formula as residual error ₃₂ And calculating the static pressure Ps at the outlet of the combustion chamber ₃₂ And combustor exit Mach number Ma ₃₂ Namely the thermal resistance loss of the combustion chamber;

that is, the finally obtained combustion chamber design point parameters include: combustion chamber exit area A ₃₂ Total pressure P at the outlet of the combustion chamber ₃₂ Outlet flow Wg of combustion chamber ₃₂ ；

Step 14: calculating the design point of the tail nozzle:

knowing that the parameters of the inlet section of the tail nozzle are equal to the parameters of the outlet section of the combustion chamber and the total temperature T of the inlet of the tail nozzle ₄₁ Total pressure P at inlet of tail nozzle ₄₁ And the inlet flow Wg of the tail nozzle ₄₁ The section of the throat part of the tail nozzle is in a critical state, and the total temperature T of the throat part of the tail nozzle ₄₃ Equal to the total inlet temperature T of the tail nozzle ₄₁ Total pressure P in the throat of the jet nozzle ₄₃ Equal to total pressure T of the inlet of the tail nozzle ₄₁ Flow Wg of the throat of the exhaust nozzle ₄₃ Equal to the inlet flow Wg of the tail nozzle ₄₁ ；

Meanwhile, the section area A of the inlet of the tail nozzle ₄₁ Area A of the cross section of the outlet of the tail nozzle ₄₂ Sectional area A of the throat portion of the tail nozzle ₄₃ The total temperature, total pressure and flow of the three cross-sectional areas are the same, namely the total temperature T of the inlet of the tail spray pipe ₄₁ Total temperature T of tail nozzle outlet ₄₂ And total temperature T of throat part of tail nozzle ₄₃ Equal, total pressure P at inlet of tail nozzle ₄₁ Total pressure P at the outlet of the tail nozzle ₄₂ And total pressure P of throat part of tail nozzle ₄₃ Equal inlet flow Wg of tail pipe ₄₁ Outlet flow Wg of tail nozzle ₄₂ And flow Wg of the throat part of the tail nozzle ₄₃ Equal;

the known parameters are brought into the flow continuity equation:

W ₄₃ ＝ρv ₄₃ A ₄₃

thereby calculating the throat area A of the tail nozzle ₄₃ Because the back pressure of the outlet of the tail nozzle is atmospheric pressure P _s0 Then, the known parameters are continuously substituted into the flow continuity equation to obtain:

Wg ₄₂ ＝ρv ₄₂ A ₄₂

and the total static pressure relationship:

Ps ₀ ＝Ps ₄₂ ＝P ₄₂ ·π(λ)

thereby calculating the outlet area A of the tail nozzle ₄₃ According to the jet nozzle outlet velocity V ₄₃ Flow Wg of the nozzle outlet ₄₃ Area A of the outlet of the tail nozzle ₄₂ And exit static pressure P of jet nozzle _s42 And the known inlet airRoad inlet flow Wa and incoming flow velocity V ₀ Substituting the formula to obtain:

Fn＝Wg ₄₂ ·V ₄₂ -Wa·V ₀ +(P _s42 -P _s0 )*A ₄₂

thereby calculating the thrust Fn of the engine;

that is, obtaining the jet nozzle design point parameters includes: throat area A of the exhaust nozzle ₄₃ And the area A of the outlet of the exhaust nozzle ₄₂ 。

Further, the second step specifically comprises the following steps:

step 21: calculating the non-design point of the air inlet channel:

according to the disclosed standard atmosphere table and the air inlet channel capture flow Wa _ c and the flight Mach number Ma of the non-design point _off Flying height H _off Calculating the actual inlet total temperature T _{in Shi} And inlet total pressure P of air inlet _{in Shi} According to the actual air inlet channel working point beta and the flight Mach number Ma of the non-design point _off Interpolated flow coefficient on air inlet channel interpolation characteristic diagram

And total pressure recovery coefficient sigma ₁ ，

Substituting the known parameters into the flow continuous equation to obtain:

coefficient of flow

And frontal area A _c Calculating to obtain the flow Wa of the actual air inlet channel _{Fruit of Chinese wolfberry} ；

Similarly, according to the formula:

P _{out of fact} ＝σ ₁ P _{in excess}

Calculating to obtain the total outlet pressure P of a non-design point _{out of fact} (ii) a Inlet duct outlet total temperature T at non-design point _{out of fact} Equal to the total inlet temperature T of the air inlet _{in excess} ；

That is, the obtained intake passage non-design point parameters include: intake duct flow Wa at non-design point _{Fruit of Chinese wolfberry} Total temperature T of outlet of air inlet channel _{out of fact} And total pressure P at the outlet of the air inlet _{out of fact} 。

Step 22: calculating the non-design point of the transition section:

according to the total pressure P of the inlet of the actual transition section _{21 Shi} Equal to the total pressure P of the outlet of the actual air inlet channel _{out of fact} Actual total temperature T of transition section inlet _{21 Shi} Equal to the actual inlet duct outlet total temperature T _{out of fact} Actual transition section inlet flow Wa ₂₁ Equal to the flow Wa of the actual inlet duct _{Fruit of Chinese wolfberry} ；

Meanwhile, according to the fact that the total pressure loss of the transition section is in direct proportion to the square of the converted flow rate of the inlet of the transition section, the converted flow rate Wac of the inlet of the transition section is calculated according to the design point _2des And the total pressure recovery coefficient sigma of the design point _2des Calculating the actual total pressure recovery coefficient sigma ₂ The formula is as follows:

according to the total pressure P of the inlet of the actual transition section _{21 Shi} And said actual total pressure recovery coefficient sigma ₂ Calculating the actual total outlet pressure P _{22 shuai Shi} Actual total outlet temperature T of transition section _{22 shuai Shi} Equal to the total inlet temperature T of the actual transition section _{21 Shi} Actual transition section outlet flow Wa _{22 shuai Shi} Equal to the actual transition section inlet flow Wa _{21 fruit} ；

That is, the obtained non-design point parameters of the transition section include: actual total pressure P at outlet of transition section _{22 fruit} Actual total outlet temperature T of transition section _{22 shuai Shi} Actual transition section outlet flow Wa _{22 fruit} ；

Step 23: calculating a non-design point of the combustion chamber:

the known outlet section of the transition section is the inlet section of the combustion chamber, and the actual inlet total temperature T of the combustion chamber _{31 true} Actual total pressure P at the inlet of the combustion chamber _{31 fruit of Chinese wolfberry} Actual combustion chamber inlet flow W _{g31 fruit} Actual combustion chamber outlet residual gas coefficient alpha and known combustion chamber inlet area A ₃₁ And combustion chamber exit area A ₃₂ ；

Substituting the known parameters into the flow continuity equation yields:

Wa _{31 true} ＝ρv ₃ A ₃₁

In the formula, wa _{31 true} Is the actual combustion chamber inlet flow, V _{3 fact} The actual incoming flow velocity of the combustion chamber; thus, the actual combustion chamber inlet static pressure P is calculated _{s31 fruit} And inlet static temperature T _{s31 fruit} ；

The cold total pressure loss of the combustion chamber is in direct proportion to the square of the inlet converted flow, and the combustion chamber inlet converted flow Wac is calculated according to the design point _3des And a cold total pressure recovery coefficient sigma of the combustion chamber _3des Calculating actual total pressure recovery coefficient sigma of combustion chamber ₃ The calculation formula is as follows:

then, according to the actual total combustion chamber inlet pressure P _{31 fruit of Chinese wolfberry} And the actual cold total pressure recovery coefficient sigma of the combustion chamber ₃ Calculating the total outlet pressure P in the actual cold state _{32 fruit} From said actual combustor inlet Mach number Ma _{31 fruit of Chinese wolfberry} Actual inlet total pressure P _{31 fruit of Chinese wolfberry} And the actual interpolated combustion efficiency eta of the residual gas coefficient alpha on the combustion chamber characteristic diagram _{b fact} According to said actual interpolated combustion efficiency η _{b fact} And the residual gas coefficient alpha to determine the actual combustion chamber outlet temperature T _{42 fact} ；

According to the flow conservation equation, the area A of the combustion chamber inlet ₃₁ Actual combustor inlet flow Wa _{31 true} Actual total pressure P at the inlet of the combustion chamber _{31 true} Actual total combustion chamber inlet temperature T _{31 true} Determining the actual combustion chamber inlet air velocity V _{31 true} (ii) a From the combustion chamber outlet area A ₃₂ Actual combustor exit flow Wg _{32 Shi} Actual total pressure P at the outlet of the combustion chamber _{32 Shi} Actual total outlet temperature T of the combustion chamber _{32 Shi} Determining the actual combustion chamber outlet gas velocity V _{32 Shi} (ii) a From the combustion chamber inlet area A ₃₁ Actual combustor inlet flow Wa _{31 true} Actual total pressure P at the inlet of the combustion chamber _{31 true} Actual total inlet temperature T of the combustion chamber _{31 true} Calculating the actual combustion chamber inlet static pressure Ps _{31 true} ；

Calculating the thermal state loss according to the momentum conservation, and substituting the known quantity into a formula to obtain:

Wa _{31 fruit of Chinese wolfberry} V _{31 fruit of Chinese wolfberry} +Ps _{31 fruit of Chinese wolfberry} A ₃₁ ＝Wg _{32 Shi} V _{32 Shi} +Ps _{32 fruit} A ₃₂

In the formula, the left side of the equal sign is impulse considering cold state loss, the right side is impulse after heating, and the momentum conservation formula is used as the total pressure P of the actual outlet of the residual iterative combustor _{32 Shi} ；

That is, the obtained non-design point parameters of the combustion chamber include: actual total outlet temperature T of combustion chamber _{32 fruit} Actual total pressure P at the outlet of the combustion chamber _{32 Shi} Actual outlet flow Wg of the combustion chamber _{32 Shi} ；

25 Step of calculating the jet nozzle non-design point:

the inlet section parameter of the tail nozzle is equal to the outlet section parameter of the combustion chamber, and the actual inlet total temperature T of the tail nozzle is known _{41 essence} Actual total pressure P of tail nozzle inlet _{41 fruit} And the actual exhaust nozzle inlet flow Wg _{41 fruit} And the throat area A of the exhaust nozzle at the design point ₄₃ Area A of the cross section of the outlet of the tail nozzle ₄₂ ；

The throat part of the tail nozzle is calculated according to a critical state, and the known parameter is the actual total temperature T of the throat part of the tail nozzle _{43 fact} Actual exhaust nozzle throat flow W _{g43 fact} And throat area A ₄₃ Substituting the flow continuous equation to obtain:

Wg _{43 fact} ＝ρv ₈ A _{43 fact}

Calculating the total pressure P required by the actual tail nozzle ₈ ，

Finally, according to the flow conservation equation, the actual outlet flow W of the tail nozzle _{g42 fact} Actual total outlet temperature T of tail nozzle _{42 fact} Outlet area A of tail nozzle ₄₂ And actual jet nozzle outlet static pressure P _s42 Determining the velocity V of the outlet of the nozzle ₉ Combined with the actual exhaust nozzle outlet flow W _{g42 fact} Outlet area A of tail nozzle ₄₂ And actual jet nozzle outlet static pressure P _s42 And inlet flow Wa and inlet speed V of the inlet ₀ Substituting the formula to obtain:

Fn _{fruit of Chinese wolfberry} ＝W _{g42 fruit} ·V _{42 fact} -Wa·V ₀ +(P _{s42 fruit} -P _s0 )*A ₄₂

The thrust of the engine at the non-design point is obtained by calculation, namely the actual thrust Fn of the engine _{Fruit of Chinese wolfberry} ；

That is, obtaining the jet nozzle non-design point parameters includes: actual thrust Fn of engine _{Fruit of Chinese wolfberry} And the total pressure P required by the actual tail nozzle ₈ 。

Further, the step 23 further includes the step of comparing the actual total outlet temperature T of the combustion chamber with the actual total outlet temperature T of the combustion chamber _{32 fruit} Actual outlet flow Wg of combustion chamber _{32 Shi} Calculation step taking into account the volume effect:

according to the actual combustion chamber volume V _{Fruit of Chinese wolfberry} Current actual combustor exit flow W _g32 Actual total enthalpy at the outlet of the combustion chamber H _{32 Shi} Actual total pressure P at outlet of combustion chamber _{32 Shi} And calculating the total enthalpy H of the outlet of the combustion chamber after considering the volume effect compared with the change delta P and delta U of the total pressure and the internal energy at the previous moment ₄ And taking into account the volumeFlow rate W of _g4 The formula is as follows:

U＝H _{32 Shi} -R·T _{32 Shi}

Total enthalpy of combustion chamber outlet H calculated from the volume considered ₄ Calculating the total temperature T of the outlet of the combustion chamber after considering the volume ₄ (ii) a R in the formula is a gas constant, and the actual total temperature T of the outlet of the combustion chamber is obtained by considering the volume effect _{32 Shi} Is equal to T ₄ Actual outlet flow Wg of the combustor _{32 fruit} Is equal to the flow W of said considered volume _g4 The calculation of the volume effect is used to reflect the effect of unsteadiness on engine performance.

Further, the total pressure P required for the actual jet nozzle in step 24 is described ₈ The calculation of (c) further comprises the steps of:

the outlet of the tail nozzle has two states, one state is that subsonic velocity is compressed to the outlet, and the corresponding static pressure is recorded as P _s1 In the first step, the ultrasonic velocity is expanded to the outlet and then a tail normal shock wave is experienced, and the corresponding static pressure is marked as P _s2 ；

Actual atmospheric pressure P _{s0 true} Greater than P _s1 When the jet pipe is in a subsonic state, the static pressure of an outlet is equal to the atmospheric pressure, and the jet pipe is in a static pressure P _s0 And area A ₉ Calculating the total pressure P required by the tail nozzle ₈ Then recalculate the Mach number Ma of the throat ₈ ；

Actual atmospheric pressure P _{s0 fact} At P _s1 And P _s2 And if a normal shock wave exists between the throat part and the outlet, the outlet of the spray pipe is subsonic, and the static pressure of the outlet is equal to the atmospheric pressure, the normal shock wave exists between the throat part and the outlet, and the static pressure is determined according to the static pressure P _s0 And jet nozzle exit area A ₄₃ Calculating the actual total pressure P of the outlet of the tail nozzle _{42 fact} At this time, the total pressure of the throat part of the spray pipe is still P ₈ ；

Actual atmospheric pressure P _{s0 true} Less than P _s2 The outlet of the spray pipe is in a supersonic speed state according to the total pressure P ₈ And the area A of the outlet of the exhaust nozzle ₄₃ Calculating the actual outlet static pressure P of the tail nozzle _s9 。

Further, when the Time =0 in the whole unsteady state process described in the fourth step, the calculation is performed without considering the volume effect of the combustion chamber, and the total temperature T of the outlet of the combustion chamber corresponding to the volume considered in the step 23 is calculated without considering the volume ₄ Value of (a) instead of T _{32 Shi} Irrespective of the volumetric flow W _g4 Value of (1) in place of Wg _{32 Shi} 。

Further, the whole unsteady state process in the fourth step is at Time>At 0, the combustion chamber volume effect is considered, corresponding to the actual combustion chamber outlet total temperature T described in step 23 _{32 Shi} Is equal to the total temperature T of the combustion chamber outlet taking into account the volume ₄ Actual outlet flow Wg of the combustion chamber _{32 Shi} Is equal to the flow W of the volume under consideration _g4 。

The design point and the non-design point refer to: when designing an engine, a specific flight condition and engine operating state corresponding to the aerodynamic thermal parameters and the geometrical dimensions of the engine and its components are determined, called the design point, i.e. the point at which the engine geometry is determined according to the performance requirements. The engine design point may be different from the design points of its components. The flight conditions and operating conditions encountered by the engine in use, which are not at the design point, are called non-design points, i.e. points where the performance is determined according to the geometry.

The beneficial effects of the invention are: when the unsteady state performance of the sub-combustion ramjet is designed, the influence of the volume effect of the combustion chamber is taken into consideration, and the thrust under the unsteady state condition of the sub-combustion ramjet can be accurately and quickly calculated. When the overall design of the sub-combustion ramjet is carried out, the design of unsteady state performance can shorten the design iteration period of an overall designer and guide a part designer to carry out detailed design.

Drawings

FIG. 1 is a flow chart of a sub-combustion ramjet design point calculation;

FIG. 2 is a flow chart of an intake port design point calculation;

FIG. 3 is a flow chart of combustion chamber design point calculations;

FIG. 4 is a flow chart of a sub-combustion ramjet non-design point calculation;

FIG. 5 is a flow chart of an intake port non-design point calculation;

FIG. 6 is a combustion chamber off-design point calculation flow chart;

FIG. 7 is a flow chart of a jet nozzle non-design point calculation;

FIG. 8 is a ramjet unsteady state performance design flow diagram;

FIG. 9 is a graph of the effect of an example of the estimation method of the unsteady state performance of the ramjet engine.

Detailed Description

The principles and features of this invention are described below in conjunction with the following drawings, which are set forth by way of illustration only and are not intended to limit the scope of the invention.

In order to achieve the above object, the present invention provides the following embodiments:

example 1: as shown in fig. 1-8, a method for estimating unsteady state performance of a scramjet engine comprises the following steps:

(1) According to the performance requirements of the ramjet, sequentially calculating and determining design points of an air inlet, a transition section, a combustion chamber and a tail nozzle, as shown in FIG. 1;

11 Step of calculating the design point of the intake port, as shown in fig. 2: given flight Mach number Ma, flight altitude H, inlet flow Wa of air inlet channel, and interpolated flow coefficient of flight Mach number Ma on inlet channel interpolation characteristic diagram

And total pressure recovery coefficient sigma, the characteristic diagram of the air inlet is flow coefficient

And a relation graph of the total pressure recovery coefficient sigma and the flight Mach number Ma;

first, according to a standard atmosphere table and said flyCalculating total inlet temperature T of air inlet by using row Mach number Ma and flight altitude H _in And inlet total pressure P of air inlet _in Total inlet duct outlet temperature T _out Equal to the total inlet temperature T of the air inlet _in ；

And according to the formula:

P _out ＝σP _in

calculating to obtain total pressure P of an outlet of the air inlet _out ；

Secondly, the given parameters are brought into a flow continuous equation to obtain:

Wa＝ρV ₀ A

that is, the obtained intake port design point parameters include: frontal area A of air inlet _c Total temperature T of outlet of air inlet channel _out And total pressure P at the outlet of the air inlet _out 。

12 Step of calculating a design point for the total pressure loss of the transition section:

the total pressure P of the inlet of the known transition section ₂₁ Equal to the total pressure P of the outlet of the air inlet passage _out Total temperature T of inlet of transition section ₂₁ Equal to the total temperature T of the outlet of the air inlet _out Total outlet temperature T of transition section ₂₂ Equal to the total temperature T of the inlet of the transition section ₂₁ (ii) a Inlet flow Wa of transition section ₂ Equal to the inlet duct flow Wa and the transition section outlet flow Wa ₂₂ Equal to the inlet flow Wa of the transition section ₂₁ (ii) a And giving a total pressure recovery coefficient sigma of the transition section _2des ，

Thereby calculating the total pressure P of the outlet of the transition section ₂₂ From saidTotal pressure P at inlet of transition section ₂₁ And the total pressure recovery coefficient sigma of the transition section _2des Calculating total pressure P of outlet of transition section ₂₂ The process of (2) represents total pressure loss;

continuously adopting known parameters and calculating the converted flow Wac of the inlet of the transition section according to a formula _2des ，

That is, the obtained transition design point parameters include: total pressure P at transition section outlet ₂₂ Total temperature T at the outlet of the transition section ₂₂ And the converted flow Wac of the inlet of the transition section _2des 。

13 Step of calculating the design point of the combustion chamber, as shown in fig. 3: the known outlet section of the transition section is the inlet section of the combustion chamber, and the total inlet temperature T of the combustion chamber ₃₁ Total pressure P at the inlet of combustion chamber ₃₁ And the cold total pressure recovery coefficient sigma of the combustion chamber _3des Inlet flow Wa of combustion chamber ₃₁ Mach number Ma of combustor inlet ₃₁ And the combustion chamber outlet temperature T ₃₂ ；

Substituting the known parameters into the flow continuous equation to obtain:

Wa ₃₁ ＝ρ ₃ v ₃ A ₃₁

thereby calculating the inlet area A of the combustion chamber ₃₁ In the formula, rho ₃ Is the combustion chamber air density, v ₃ Is the combustion chamber incoming flow velocity;

Since the combustion chamber is an equal straight path, the combustion chamber outlet area A ₃₂ Equal to the combustion chamber inlet area A ₃₁ (ii) a According to the total pressure P of the inlet of the combustion chamber ₃₁ And the cold total pressure recovery coefficient sigma of the combustion chamber _3des ，

Further according to the flow conservation equation, the inlet area A of the combustion chamber ₃₁ Inlet flow Wa of combustion chamber ₃₁ Total pressure at the inlet of the combustion chamber P ₃₁ Total inlet temperature T of combustion chamber ₃₁ Determining the velocity V of the combustion chamber inlet air ₃₁ (ii) a From the combustion chamber outlet area A ₃₂ Outlet flow Wg of combustion chamber ₃₂ Total pressure at outlet of combustion chamber P ₃₂ Total outlet temperature T of combustion chamber ₃₂ Determining the velocity V of the gas at the outlet of the combustion chamber ₃₂ (ii) a From the combustion chamber inlet area A ₃₁ Inlet flow Wa of combustion chamber ₃₁ Inlet total pressure P of combustion chamber ₃₁ Total inlet temperature T of combustion chamber ₃₁ Calculating the static pressure Ps at the inlet of the combustion chamber ₃₁ ；

Then, calculating the thermal state loss according to the momentum conservation, substituting the known parameters into a momentum conservation formula to obtain:

Wa ₃ V ₃₁ +Ps ₃₁ A ₃₁ ＝Wg ₃₂ V ₃₂ +Ps ₃₂ A ₃₂

in the formula, the left side of the equal sign is impulse considering cold state loss, the right side is impulse after heating, and the total pressure P of the outlet of the combustion chamber is obtained by using a momentum conservation formula as residual error iteration ₃₂ And calculating the combustor exit static pressure Ps ₃₂ And the combustor exit Mach number Ma ₃₂ Namely, the thermal resistance loss of the combustion chamber is obtained;

that is, the finally obtained combustion chamber design point parameters include: combustion chamber exit area A ₃₂ Total pressure P at the outlet of the combustion chamber ₃₂ Outlet flow Wg of combustion chamber ₃₂ 。

14 Step of calculating the jet nozzle design point: known jet nozzlesThe port section parameter is equal to the combustion chamber outlet section parameter and the total inlet temperature T of the tail nozzle ₄₁ Total pressure P at inlet of tail nozzle ₄₁ And the inlet flow Wg of the tail nozzle ₄₁ The section of the throat part of the tail nozzle is in a critical state, and the total temperature T of the throat part of the tail nozzle ₄₃ Equal to the total inlet temperature T of the tail nozzle ₄₁ Total pressure P of throat part of tail nozzle ₄₃ Equal to total pressure T of the inlet of the tail nozzle ₄₁ Flow Wg of the throat of the exhaust nozzle ₄₃ Equal to the inlet flow Wg of the tail nozzle ₄₁ ；

Meanwhile, the cross-sectional area A of the inlet of the tail nozzle ₄₁ Area A of the cross section of the outlet of the tail nozzle ₄₂ Sectional area A of the throat portion of the tail nozzle ₄₃ The total temperature, total pressure and flow of the three cross-sectional areas are the same, namely the total temperature T of the inlet of the tail spray pipe ₄₁ Total outlet temperature T of tail nozzle ₄₂ And total temperature T of throat part of tail nozzle ₄₃ Equal, total pressure P at inlet of tail nozzle ₄₁ Total pressure P at outlet of tail nozzle ₄₂ And total pressure P of throat part of tail nozzle ₄₃ Equal inlet flow Wg of tail pipe ₄₁ Outlet flow Wg of tail nozzle ₄₂ And flow Wg of the throat part of the tail nozzle ₄₃ Equal;

the known parameters are brought into the flow continuity equation:

W ₄₃ ＝ρv ₄₃ A ₄₃

thereby calculating the throat area A of the tail nozzle ₄₃ Because the back pressure of the outlet of the tail nozzle is atmospheric pressure P _s0 ，

Then the known parameters are continuously brought into the flow continuous equation to obtain:

Wg ₄₂ ＝ρv ₄₂ A ₄₂

and the total static pressure relationship:

Ps ₀ ＝Ps ₄₂ ＝P ₄₂ ·π(λ)

thereby calculating and obtaining the outlet area A of the tail nozzle ₄₃ According to the jet outlet velocity V of the tail pipe ₄₃ Flow Wg of the nozzle outlet ₄₃ Outlet area A of tail nozzle ₄₂ And outlet static pressure P of tail nozzle _s42 And the known inlet flow Wa and the known incoming flow velocity V of the inlet passage ₀ Substituting the formula to obtain:

Fn＝Wg ₄₂ ·V ₄₂ -Wa·V ₀ +(P _s42 -P _s0 )*A ₄₂

thereby calculating the thrust Fn of the engine;

that is, obtaining the jet nozzle design point parameters includes: area A of the throat of the exhaust nozzle ₄₃ And the area A of the outlet of the exhaust nozzle ₄₂ 。

(2) Calculating and determining the non-design point of the performance of the ramjet according to the parameters of the design points of the air inlet channel, the transition section, the combustion chamber and the tail nozzle of the ramjet obtained in the step (1) and the flow balance relation of the air inlet channel and the tail nozzle, as shown in FIG. 4;

21 Step of calculating the port non-design point, as shown in fig. 5: according to the disclosed standard atmosphere table and the air inlet channel capture flow Wa _ c and the flight Mach number Ma of the non-design point _off Flying height H _off Calculating the actual inlet total temperature T _{in Shi} And inlet total pressure P of air inlet _{in Shi} According to the actual air inlet channel working point beta and the flight Mach number Ma of the non-design point _off Interpolated flow coefficient on air inlet channel interpolation characteristic diagram

And total pressure recovery coefficient sigma ₁ ，

Substituting the known parameters into the flow continuous equation to obtain:

coefficient of flow

Similarly, according to the formula:

P _{out of fact} ＝σ ₁ P _{in Shi}

Calculating to obtain the total outlet pressure P of the non-design point _{out of fact} (ii) a Inlet duct outlet total temperature T at non-design point _{out of fact} Equal to the total inlet temperature T of the air inlet _{in Shi} 。

That is, the obtained intake passage non-design point parameters include: intake passage flow Wa at non-design point _{Fruit of Chinese wolfberry} Total temperature T of outlet of air inlet passage _{out of fact} And total pressure P at the outlet of the air inlet _{out of fact} 。

22 Step of calculating the non-design point of the transition section: according to the actual total pressure P of the transition section inlet _{21 Shi} Equal to the total pressure P of the outlet of the actual air inlet channel _{out of fact} Actual total temperature T of transition section inlet _{21 Shi} Equal to the actual inlet duct outlet total temperature T _{out of fact} Actual transition section inlet flow Wa ₂₁ Equal to the flow Wa of the actual inlet duct _{Fruit of Chinese wolfberry} ；

Meanwhile, according to the fact that the total pressure loss of the transition section is in direct proportion to the square of the converted flow rate of the inlet of the transition section, the converted flow rate Wac of the inlet of the transition section is calculated according to the design point _2des And the total pressure recovery coefficient sigma of the design point _2des Calculating the actual total pressure recovery coefficient sigma ₂ The formula is as follows.

According to the total pressure P of the inlet of the actual transition section _{21 Shi} And said actual total pressure recovery coefficient sigma ₂ Calculating the actual total outlet pressure P _{22 shuai Shi} Actual total outlet temperature T of transition section _{22 shuai Shi} Equal to the actual transition section inletTotal temperature T _{21 fruit} Actual transition section outlet flow Wa _{22 fruit} Equal to the actual transition section inlet flow Wa _{21 fruit} ；

That is, the obtained transition section non-design point parameters include: actual total pressure P at outlet of transition section _{22 fruit} Actual total outlet temperature T of transition section _{22 fruit} Actual transition section outlet flow Wa _{22 shuai Shi} 。

23 Step of calculating the non-design point of the combustion chamber, as shown in fig. 6: the known outlet section of the transition section is the inlet section of the combustion chamber, and the actual inlet total temperature T of the combustion chamber _{31 true} Actual total pressure P at the inlet of the combustion chamber _{31 true} Actual combustion chamber inlet flow W _{g31 fruit} Actual combustion chamber outlet residual gas coefficient alpha and known combustion chamber inlet area A ₃₁ And combustion chamber exit area A ₃₂ ；

Substituting the known parameters into the flow continuity equation yields:

Wa _{31 fruit of Chinese wolfberry} ＝ρv ₃ A ₃₁

In the formula, wa _{31 fruit of Chinese wolfberry} Is the actual combustion chamber inlet flow, V _{3 fact} The actual incoming flow velocity of the combustion chamber; thus, the actual combustion chamber inlet static pressure P is calculated _{s31 fruit} And inlet static temperature T _{s31 fruit} ；

The cold total pressure loss of the combustion chamber is in direct proportion to the square of the converted inlet flow rate, and the converted inlet flow rate Wac of the combustion chamber according to the design point _3des And a cold total pressure recovery coefficient sigma of the combustion chamber _3des Calculating the actual total pressure recovery coefficient sigma of the combustion chamber ₃ The calculation formula is as follows:

then, according to the actual total combustion chamber inlet pressure P _{31 fruit of Chinese wolfberry} And the actual cold total pressure recovery coefficient sigma of the combustion chamber ₃ Calculating the actual cold outlet total pressure P _{32 Shi} From said actual combustor inlet Mach number Ma _{31 true} Actual inlet total pressure P _{31 true} And the actual interpolated combustion efficiency eta of the residual gas coefficient alpha on the combustion chamber characteristic diagram _{b fruit} According to said actual interpolated combustion efficiency η _{b fact} And the residual gas coefficient alpha to determine the actual combustion chamber outlet temperature T _{42 fruit of Chinese wolfberry} ；

According to the flow conservation equation, the inlet area A of the combustion chamber ₃₁ Actual combustor inlet flow Wa _{31 true} Actual total pressure P at the inlet of the combustion chamber _{31 true} Actual total combustion chamber inlet temperature T _{31 true} Determining the actual combustion chamber inlet air velocity V _{31 true} (ii) a From the combustion chamber outlet area A ₃₂ Actual combustor exit flow Wg _{32 fruit} Actual total pressure P at the outlet of the combustion chamber _{32 Shi} Actual total outlet temperature T of the combustion chamber _{32 fruit} Determining the actual combustion chamber outlet gas velocity V _{32 Shi} (ii) a From the combustion chamber inlet area A ₃₁ Actual combustor inlet flow Wa _{31 true} Actual total pressure P at the inlet of the combustion chamber _{31 true} Actual total combustion chamber inlet temperature T _{31 true} Calculating the actual combustion chamber inlet static pressure Ps _{31 true} ；

Calculating the thermal state loss according to momentum conservation, and substituting the known quantity into a formula to obtain:

Wa _{31 true} V _{31 true} +Ps _{31 true} A ₃₁ ＝Wg _{32 Shi} V _{32 Shi} +Ps _{32 Shi} A ₃₂

According to the actual combustion chamber volume V _{Fruit of Chinese wolfberry} Current actual combustor exit flow W _g32 Actual total enthalpy at the outlet of the combustion chamber H _{32 Shi} Actual total pressure P at the outlet of the combustion chamber _{32 Shi} And calculating the total enthalpy H of the outlet of the combustion chamber after considering the volume effect compared with the change delta P and delta U of the total pressure and the internal energy at the previous moment ₄ And a flow rate W taking into account the volume _g4 The formula is as follows:

U＝H _{32 fruit} -R·T _{32 Shi}

Total enthalpy of combustion chamber outlet H calculated from the volume considered ₄ Calculating the total temperature T of the outlet of the combustion chamber after considering the volume ₄ ；

R in the formula is a gas constant, and the actual total temperature T of the outlet of the combustion chamber is obtained by considering the volume effect _{32 Shi} Is equal to T ₄ Actual outlet flow Wg of the combustion chamber _{32 Shi} Is equal to the flow W of said considered volume _g4 The calculation of the volume effect is used to reflect the influence of unsteadiness on the engine performance;

that is, the obtained combustion chamber non-design point parameters include: actual total outlet temperature T of combustion chamber _{32 Shi} Actual total pressure P at the outlet of the combustion chamber _{32 Shi} Actual outlet flow Wg of the combustion chamber _{32 fruit} 。

24 Step of calculating the jet nozzle non-design point, as shown in FIG. 7:

the sectional parameter of the inlet of the tail nozzle is equal to the sectional parameter of the outlet of the combustion chamber, and the actual total temperature T of the inlet of the tail nozzle is known _{41 fruit} Actual total pressure P of tail nozzle inlet _{41 essence} And the actual exhaust nozzle inlet flow Wg _{41 essence} And the throat area A of the exhaust nozzle at the design point ₄₃ Sectional area A of the outlet of the tail nozzle ₄₂ ；

Wg _{43 fact} ＝ρv ₈ A _{43 fact}

Calculating the total pressure P required by the tail nozzle ₈ ，

The outlet of the tail nozzle has two key states, one is subsonic compression to the outlet, and the corresponding static pressure is recorded as P _s1 In the first step, the ultrasonic velocity is expanded to the outlet and then a tail normal shock wave is experienced, and the corresponding static pressure is marked as P _s2 ；

Actual atmospheric pressure P _{s0 true} Greater than P _s1 When the pressure is in the subsonic speed state, the static pressure at the outlet is equal to the atmospheric pressure, and the pressure is in accordance with the static pressure P _s0 And area A ₉ Calculating the total pressure P required by the tail nozzle ₈ Then recalculate the Mach number Ma of the throat ₈ ；

Actual atmospheric pressure P _{s0 true} At P _s1 And P _s2 And if a normal shock wave exists between the throat part and the outlet, the outlet of the spray pipe is subsonic, and the static pressure of the outlet is equal to the atmospheric pressure, the normal shock wave exists between the throat part and the outlet, and the static pressure is determined according to the static pressure P _s0 And the area A of the outlet of the exhaust nozzle ₄₃ Calculating the actual total pressure P of the outlet of the tail nozzle _{42 fruit of Chinese wolfberry} At this time, the total pressure of the throat part of the spray pipe is still P ₈ ；

Actual atmospheric pressure P _{s0 true} Less than P _s2 The outlet of the spray pipe is in a supersonic speed state according to the total pressure P ₈ And the area A of the outlet of the exhaust nozzle ₄₃ Calculating the actual outlet static pressure P of the tail nozzle _s9 ；

Finally, according to the flow conservation equation, the actual outlet flow W of the tail nozzle _{g42 fact} Actual total outlet temperature T of tail nozzle _{42 fact} Outlet area A of tail nozzle ₄₂ And actual jet nozzle outlet static pressure P _s42 Determining the velocity V of the outlet of the nozzle ₉ And combined with the actual exhaust nozzle outlet flow W _{g42 fruit} Outlet area A of tail nozzle ₄₂ And actual jet nozzle outlet static pressure P _s42 And inlet flow Wa and inlet velocity V of the inlet ₀ Substituting the formula to obtain:

Fn _{fruit of Chinese wolfberry} ＝W _{g42 fact} ·V _{42 fact} -Wa·V ₀ +(P _{s42 fact} -P _s0 )*A ₄₂

Thus, the thrust of the engine at the non-design point is obtained by calculation, namely the actual thrust Fn of the engine _{Fruit of Chinese wolfberry} ；

(3) The total pressure P required by the tail nozzle obtained in the step 24) ₈ Total pressure P with actual tail nozzle inlet _{41 essence} Forming a residual equation as shown in the following formula:

ERR＝(P _{41 essence} -P ₈ )/P _{41 essence}

Changing the value of a residual error equation by changing the actual working point beta of the air inlet in the step 21), iterating the actual working point beta of the air inlet by using a Newton method, and when the absolute value of the residual error is less than 0.0001, determining that iteration convergence is achieved;

(4) As shown in FIG. 8, the residual air coefficient α and the flight Mach number Ma are obtained from the non-design point _off And a flying height H _off Iterating the actual working point beta of the air inlet channel point by point according to the change rule of the Time along with the step (3) so as to obtain the actual thrust Fn of the engine calculated in the step (2) _{Fruit of Chinese wolfberry} Until the whole unsteady state process is calculated, when Time =0, the volume effect of the combustion chamber is not considered, and the volumetric total temperature T of the combustion chamber outlet is considered in the corresponding step 23 ₄ Value of (a) in place of T _{32 fruit} Irrespective of the volumetric flow W _g4 Value of (2) in place of Wg _{32 Shi} (ii) a At Time>At 0, the combustion chamber volume effect is considered, corresponding to the actual combustion chamber outlet total temperature T described in step 23 _{32 fruit} Is equal to the total temperature T of the combustion chamber outlet taking into account the volume ₄ Actual outlet flow Wg of the combustor _{32 fruit} Is equal to the flow W of the volume under consideration _g4 . (ii) a Calculating the Time until the Time reaches a given Time limit Time _MAX And finally obtaining an unsteady thrust curve of the engine considering the volume effect.

The specific calculation example is as follows:

a design method for unsteady state performance of a sub-combustion ramjet engine comprises the following calculation steps:

(1) In the specific embodiment, taking the design of unsteady state performance of a certain sub-combustion ramjet engine as an example, when calculating a design point of an air inlet, flight conditions and characteristic parameters of the air inlet are required, wherein the flight conditions and the characteristic parameters of the air inlet include a flight mach number Ma =4.0, a flight altitude H =21500m, an inlet flow Wa =10.0kg/s, a total pressure recovery coefficient σ =0.524, and a flow coefficient

According to the parameters, the atmospheric conditions and the flow rate are continuously expressed

The sum formula Wa = rho vA can be calculated to obtain the frontal area A =0.12225m of the air inlet ² Total pressure at the outlet P ₂ =344339.31Pa, total outlet temperature T ₂ ＝916.23K。

(2) According to the parameters, the total pressure recovery coefficient sigma of the design point of the transition section is given _2des =0.9, obtaining total pressure P at the outlet of the transition section ₃ =300608.22Pa, total temperature T ₃ ＝916.23K。

(3) Coefficient of restitution sigma by total pressure in given combustion chamber cold state _3des =0.93, combustion efficiency eta _b =0.92, inlet mach number Ma ₃ =0.1, total outlet temperature T ₄ =2000K, and the total outlet pressure P is obtained through calculation according to momentum conservation and energy conservation ₄ =276576.41Pa, combustion chamber area A ₃ ＝0.14805m ² Amount of oil supply W _fb =0.37660kg/s, and the outlet Mach number is Ma ₄ ＝0.1733。

(4) Determining outlet back pressure P of tail nozzle according to flight altitude _s0 =4325.2Pa, throat area a calculated from throat critical conditions ₈ ＝0.04292m ² Calculating the outlet area A of the tail nozzle according to the complete expansion condition ₉ ＝0.30138m ² Mach number at the outlet of Ma ₉ =3.2495, exit velocity V ₉ =1766.9m/s. Calculating the incoming flow velocity V from the flight Mach number and altitude ₀ =1184.3m/s. Fn =614.94kgf can be obtained from the thrust force calculation formula.

(5) On the basis of the result of the calculation of the design point, the ramjet is givenOil supply W in unsteady state process _fb The change rule of the flight Mach number Ma and the flight altitude H along with the Time. The flight height H =21500m, the flight Mach number Ma =4.0, the flight height is kept unchanged, and the oil supply quantity is linearly increased from 0.2kg/s to 0.3766kg/s within 3.6 s. The calculated time step is 0.025s. Combustion chamber volume V =1m ³ . The unsteady state performance of the engine is designed by adopting the non-design point model established by the method, the change rule of the thrust is shown in figure 9, and the thrust of the ramjet is changed from 316.22kgf to 614.94kgf.

The above description is only for the purpose of illustrating the preferred embodiments of the present invention and should not be taken as limiting the scope of the present invention, which is intended to cover any modifications, equivalents, improvements, etc. within the spirit and scope of the present invention.

Claims

1. A method for estimating unsteady state performance of a sub-combustion ramjet engine is characterized by comprising the following steps:

step one, according to a given flight Mach number Ma, a flight altitude H, an inlet flow Wa of an air inlet channel and an interpolated flow coefficient of the flight Mach number Ma on an interpolation characteristic diagram of the air inlet channel

the determined inlet design point parameters include: frontal area A of air inlet _c Total temperature T of outlet of air inlet passage _out And the total pressure P of the outlet of the air inlet _out (ii) a The determined design point parameters for the transition section include: total pressure P at transition section outlet ₂₂ Total temperature T at the outlet of the transition section ₂₂ And the converted flow Wac of the inlet of the transition section _2des (ii) a The determined combustion chamber design point parameters include: combustion chamber exit area A ₃₂ Total pressure P at the outlet of the combustion chamber ₃₂ Outlet flow Wg of combustion chamber ₃₂ (ii) a The determined jet nozzle design point parameters include: area A of the throat of the exhaust nozzle ₄₃ And the area A of the outlet of the exhaust nozzle ₄₂ ；

Secondly, combining parameters of the ramjet air inlet channel, the transition section, the combustion chamber and the tail nozzle design point and the flow balance relation of the air inlet channel and the tail nozzle, and capturing the flow Wa _ c and the flight Mach number Ma of the air inlet channel according to a disclosed standard atmosphere table and a non-design point _off Flying height H _off Calculating the actual inlet total temperature T _{in Shi} And inlet total pressure P of air inlet _{in Shi} According to the actual air inlet channel working point beta and the flight Mach number Ma of the non-design point _off Interpolated flow coefficient on air inlet channel interpolation characteristic diagram

the determined air inlet channel non-design point parameters comprise: intake passage flow Wa at non-design point _{Fruit of Chinese wolfberry} Total temperature T of outlet of air inlet passage _{out of fact} And the total pressure P of the outlet of the air inlet _{out of fact} (ii) a The determined transition section non-design point parameters include: actual total pressure P at outlet of transition section _{22 shuai Shi} Actual total outlet temperature T of transition section _{22 shuai Shi} Actual transition section outlet flow Wa _{22 shuai Shi} (ii) a The determined non-design point parameters of the combustion chamber comprise: actual total outlet temperature T of combustion chamber _{32 fruit} Actual total pressure P at outlet of combustion chamber _{32 Shi} Actual outlet flow Wg of the combustor _{32 fruit} (ii) a The determined non-design point parameters of the jet nozzle comprise: actual thrust Fn of engine _{Fruit of Chinese wolfberry} And the total pressure P required by the actual tail nozzle ₈ ；

Step three, obtaining the total pressure P required by the actual tail nozzle according to the step two ₈ And the known actual total pressure P of the inlet of the tail nozzle _{41 fruit} Forming a residual equation as shown in the following formula:

ERR＝(P _{41 fruit} -P ₈ )/P _{41 fruit}

fourthly, according to the residual air coefficient alpha of the non-design point and the flight Mach number Ma _off And a flying height H _off Iterating the actual working point beta of the air inlet channel point by point according to the change rule of the Time along with the step three, thereby obtaining the actual thrust Fn of the engine calculated in the step two _{Fruit of Chinese wolfberry} Until Time =0 to a given Time limit Time _MAX And finally obtaining the unsteady thrust curve of the engine considering the volume effect after the calculation of the whole unsteady process is finished.

2. The method for estimating the unsteady state performance of the sub-combustion ramjet engine as recited in claim 1, wherein said step one comprises the steps of:

step 11: calculating the design point of the air inlet channel: given flight Mach number Ma, flight altitude H, inlet flow Wa of the air inlet and interpolated flow coefficient of the flight Mach number Ma on the interpolation characteristic diagram of the air inlet

And a total pressure recovery coefficient sigma;

firstly, calculating the total temperature T of the inlet of the air inlet according to a standard atmosphere table and the flight Mach number Ma and the flight altitude H _in And inlet total pressure P of air inlet _in Total inlet duct outlet temperature T _out Equal to the total inlet temperature T of the air inlet _in ；

And according to the formula:

P _out ＝σP _in

calculating to obtain total pressure P of an outlet of the air inlet passage _out ；

Wa＝ρV ₀ A

calculating the windward area A of the air inlet _c In the formula, K is 0.0404 and is a fixed value; q (λ) is a flow function, ρ is the air density, V ₀ The incoming flow velocity, A is the flow tube area;

the total pressure P of the inlet of the known transition section ₂₁ Equal to the total pressure P of the outlet of the air inlet passage _out Total temperature T of inlet of transition section ₂₁ Equal to the total temperature T of the inlet duct outlet _out Total outlet temperature T of transition section ₂₂ Equal to the total temperature T of the inlet of the transition section ₂₁ (ii) a Inlet flow Wa of transition section ₂ Equal to the inlet duct flow Wa and the transition section outlet flow Wa ₂₂ Equal to the inlet flow Wa of the transition section ₂₁ (ii) a And giving a total pressure recovery coefficient sigma of the transition section _2des ，

That is, the obtained transition design point parameters include: total pressure P at outlet of transition section ₂₂ Total temperature T at the outlet of the transition section ₂₂ And converted flow Wac at the inlet of the transition section _2des ；

Step 13: calculating the design point of the combustion chamber:

known transitionsThe section outlet cross section is the combustion chamber inlet cross section, and the total temperature T of the combustion chamber inlet ₃₁ Total pressure P at the inlet of the combustion chamber ₃₁ And the cold total pressure recovery coefficient sigma of the combustion chamber _3des Inlet flow Wa of combustion chamber ₃₁ Mach number Ma of combustor inlet ₃₁ And the combustion chamber outlet temperature T ₃₂ ；

Substituting the known parameters into the flow continuous equation to obtain:

Wa ₃₁ ＝ρ ₃ v ₃ A ₃₁

Continuously adopting known parameters, and calculating the converted flow Wac of the combustion chamber inlet according to a formula _3des ：

Further according to the flow conservation equation, the inlet area A of the combustion chamber ₃₁ Inlet flow Wa of combustion chamber ₃₁ Total pressure at the inlet of the combustion chamber P ₃₁ Total inlet temperature T of combustion chamber ₃₁ Determining the velocity V of the combustion chamber inlet air ₃₁ (ii) a From the combustion chamber outlet area A ₃₂ Outlet flow Wg of combustion chamber ₃₂ Total pressure P at the outlet of the combustion chamber ₃₂ Total outlet temperature T of combustion chamber ₃₂ Determining the velocity V of the gas flow at the outlet of the combustion chamber ₃₂ (ii) a From the combustion chamber inlet area A ₃₁ Inlet flow Wa of combustion chamber ₃₁ Total pressure at the inlet of the combustion chamber P ₃₁ Total inlet temperature T of combustion chamber ₃₁ Calculating the static pressure Ps at the inlet of the combustion chamber ₃₁ ；

Wa ₃ V ₃₁ +Ps ₃₁ A ₃₁ ＝Wg ₃₂ V ₃₂ +Ps ₃₂ A ₃₂

in the formula, the left side of the equal sign is impulse considering cold state loss, the right side is impulse after heating, and the total pressure P of the outlet of the combustion chamber is obtained by using a momentum conservation formula as residual error iteration ₃₂ And calculating the combustor exit static pressure Ps ₃₂ And combustor exit Mach number Ma ₃₂ Namely the thermal resistance loss of the combustion chamber;

Step 14: calculating a design point of the tail nozzle:

Meanwhile, the section area A of the inlet of the tail nozzle ₄₁ Sectional area A of the outlet of the tail nozzle ₄₂ Sectional area A of throat part of tail nozzle ₄₃ The total temperature, total pressure and flow of the three cross-sectional areas are the same, namely the total temperature T of the inlet of the tail spray pipe ₄₁ Tail spray pipeTotal outlet temperature T ₄₂ And total temperature T of throat part of tail nozzle ₄₃ Equal, total pressure P at inlet of tail nozzle ₄₁ Total pressure P at the outlet of the tail nozzle ₄₂ And total pressure P of throat part of tail nozzle ₄₃ Equal inlet flow Wg of tail pipe ₄₁ And the outlet flow Wg of the tail spray pipe ₄₂ And flow Wg of the throat part of the tail nozzle ₄₃ Equal;

bringing the known parameters into the flow continuity equation:

W ₄₃ ＝ρv ₄₃ A ₄₃

thereby calculating the throat area A of the tail nozzle ₄₃ Because the back pressure of the outlet of the tail nozzle is atmospheric pressure P _s0 Then, the known parameters are continuously introduced into the flow continuous equation to obtain:

Wg ₄₂ ＝ρv ₄₂ A ₄₂

and the total static pressure relationship:

Ps ₀ ＝Ps ₄₂ ＝P ₄₂ ·π(λ)

Fn＝Wg ₄₂ ·V ₄₂ -Wa·V ₀ +(P _s42 -P _s0 )*A ₄₂

thereby calculating the thrust Fn of the engine;

that is, obtaining the jet nozzle design point parameters includes: throat area A of the exhaust nozzle ₄₃ And jet nozzle exit area A ₄₂ 。

3. The method for estimating the unsteady state performance of the sub-combustion ramjet engine as recited in claim 1, wherein said second step comprises the steps of:

step 21: calculating the non-design point of the air inlet channel:

And total pressure recovery coefficient sigma ₁ ，

Substituting the known parameters into the flow continuous equation to obtain:

coefficient of flow

Similarly, according to the formula:

P _{out of fact} ＝σ ₁ P _{in excess}

Calculating to obtain the total outlet pressure P of the non-design point _{out of fact} (ii) a Inlet duct outlet total temperature T at non-design point _{out of fact} Equal to the total inlet temperature T of the air inlet _{in Shi} ；

That is to say that the first and second electrodes,the obtained non-design point parameters of the air inlet channel comprise: intake passage flow Wa at non-design point _{Fruit of Chinese wolfberry} Total temperature T of outlet of air inlet channel _{out of fact} And total pressure P at the outlet of the air inlet _{out of fact} 。

Step 22: calculating the non-design point of the transition section:

according to the total pressure P of the inlet of the actual transition section _{21 Shi} Equal to the total pressure P of the outlet of the actual air inlet channel _{out of fact} Actual total temperature T at the transition section inlet _{21 Shi} Equal to the actual inlet duct outlet total temperature T _{out of fact} Actual transition section inlet flow Wa ₂₁ Equal to the flow Wa of the actual inlet _{Fruit of Chinese wolfberry} ；

according to the total pressure P of the inlet of the actual transition section _{21 Shi} And said actual total pressure recovery coefficient sigma ₂ Calculating the actual total outlet pressure P _{22 shuai Shi} Actual total outlet temperature T of transition section _{22 fruit} Equal to the actual total temperature T of the transition section inlet _{21 Shi} Actual transition section outlet flow Wa _{22 shuai Shi} Equal to the actual transition section inlet flow Wa _{21 Shi} ；

That is, the obtained transition section non-design point parameters include: actual total pressure P at outlet of transition section _{22 shuai Shi} Actual total outlet temperature T of transition section _{22 shuai Shi} Actual transition section outlet flow Wa _{22 fruit} ；

Step 23: calculating a non-design point of the combustion chamber:

the known outlet section of the transition section is the inlet section of the combustion chamber, and the actual inlet total temperature T of the combustion chamber _{31 true} Actual total pressure P at the inlet of the combustion chamber _{31 fruit of Chinese wolfberry} Actual combustion chamber inlet flow W _{g31 fruit} Fruit of Chinese wolfberryThe outlet residual gas coefficient alpha of the inter-combustion chamber and the known inlet area A of the combustion chamber ₃₁ And combustion chamber exit area A ₃₂ ；

Substituting the known parameters into the flow continuity equation yields:

Wa _{31 fruit of Chinese wolfberry} ＝ρv ₃ A ₃₁

In the formula, wa _{31 fruit of Chinese wolfberry} Is the actual combustion chamber inlet flow, V _{3 fact} The actual combustion chamber inflow speed; thus, the actual combustion chamber inlet static pressure P is calculated _{s31 fruit} And inlet static temperature T _{s31 fruit} ；

The cold total pressure loss of the combustion chamber is in direct proportion to the square of the inlet converted flow, and the combustion chamber inlet converted flow Wac is calculated according to the design point _3des And the cold total pressure recovery coefficient sigma of the combustion chamber _3des Calculating actual total pressure recovery coefficient sigma of combustion chamber ₃ The calculation formula is as follows:

then, according to the actual total combustion chamber inlet pressure P _{31 fruit of Chinese wolfberry} And the actual cold total pressure recovery coefficient sigma of the combustion chamber ₃ Calculating the total outlet pressure P in the actual cold state _{32 fruit} From said actual combustor inlet Mach number Ma _{31 true} Actual inlet total pressure P _{31 fruit of Chinese wolfberry} And the actual interpolated combustion efficiency eta of the residual gas coefficient alpha on the combustion chamber characteristic diagram _{b fact} According to said actual interpolated combustion efficiency eta _{b fact} And the residual gas coefficient alpha to determine the actual combustion chamber outlet temperature T _{42 fact} ；

According to the flow conservation equation, the area A of the combustion chamber inlet ₃₁ Actual combustor inlet flow Wa _{31 true} Actual total pressure P at the inlet of the combustion chamber _{31 true} Actual total combustion chamber inlet temperature T _{31 true} Determining actual combustionVelocity V of air at chamber inlet _{31 true} (ii) a From the combustion chamber outlet area A ₃₂ Actual combustor exit flow Wg _{32 fruit} Actual total pressure P at the outlet of the combustion chamber _{32 Shi} Actual total outlet temperature T of the combustion chamber _{32 fruit} Determining the actual combustion chamber outlet gas velocity V _{32 fruit} (ii) a From the combustion chamber inlet area A ₃₁ Actual combustion chamber inlet flow Wa _{31 fruit of Chinese wolfberry} Actual total pressure P at the inlet of the combustion chamber _{31 fruit of Chinese wolfberry} Actual total combustion chamber inlet temperature T _{31 fruit of Chinese wolfberry} Calculating the actual combustion chamber inlet static pressure Ps _{31 fruit of Chinese wolfberry} ；

Wa _{31 fruit of Chinese wolfberry} V _{31 fruit of Chinese wolfberry} +Ps _{31 fruit of Chinese wolfberry} A ₃₁ ＝Wg _{32 Shi} V _{32 fruit} +Ps _{32 fruit} A ₃₂

In the formula, the left side of the equal sign is impulse considering cold state loss, the right side is impulse after heating, and a momentum conservation formula is used as the total pressure P of the actual outlet of the residual iterative combustion chamber _{32 fruit} ；

That is, the obtained non-design point parameters of the combustion chamber include: actual total outlet temperature T of combustion chamber _{32 fruit} Actual total pressure P at outlet of combustion chamber _{32 Shi} Actual outlet flow Wg of the combustion chamber _{32 Shi} ；

Step 24, calculating the non-design point of the tail nozzle:

the inlet section parameter of the tail nozzle is equal to the outlet section parameter of the combustion chamber, and the actual inlet total temperature T of the tail nozzle is known _{41 essence} Actual total pressure P of tail nozzle inlet _{41 essence} And the actual exhaust nozzle inlet flow Wg _{41 essence} And the nozzle throat area A at the design point ₄₃ Sectional area A of the outlet of the tail nozzle ₄₂ ；

The throat part of the tail nozzle is calculated according to a critical state, and the known parameter is the actual total temperature T of the throat part of the tail nozzle _{43 fact} Actual exhaust nozzle throat flow W _{g43 fruit} And throat area A ₄₃ Substituting the flow continuous equation to obtain:

Wg _{43 fact} ＝ρv ₈ A _{43 fact}

Calculating the total pressure P required by the actual tail nozzle ₈ ，

Finally, according to the flow conservation equation, the actual outlet flow W of the tail spray pipe _{g42 fact} Actual total temperature T of tail nozzle outlet _{42 fruit of Chinese wolfberry} Outlet area A of tail nozzle ₄₂ And actual jet nozzle outlet static pressure P _s42 Determining the velocity V of the outlet of the nozzle ₉ And combined with the actual exhaust nozzle outlet flow W _{g42 fruit} Outlet area A of tail nozzle ₄₂ And actual jet nozzle outlet static pressure P _s42 And inlet flow Wa and inlet speed V of the inlet ₀ Substituting the formula to obtain:

Fn _{fruit of Chinese wolfberry} ＝W _{g42 fact} ·V _{42 fruit of Chinese wolfberry} -Wa·V ₀ +(P _{s42 fact} -P _s0 )*A ₄₂

4. The method of estimating unsteady state performance of a scramjet engine as recited in claim 3 wherein said step 23 further comprises estimating said actual combustor exit total temperature T _{32 Shi} Actual outlet flow Wg of combustion chamber _{32 Shi} Calculation step taking into account the volume effect:

according to the actual combustion chamber volume V _{Fruit of Chinese wolfberry} Current actual combustor exit flow W _g32 Actual total enthalpy at the outlet of the combustion chamber H _{32 Shi} Actual total pressure P at the outlet of the combustion chamber _{32 Shi} And calculating the total enthalpy H of the outlet of the combustion chamber after considering the volume effect compared with the change delta P and delta U of the total pressure and the internal energy at the previous moment ₄ And a volumetric flow rate W _g4 The formula is as follows:

U＝H _{32 Shi} -R·T _{32 Shi}

From the total enthalpy H of the outlet of the combustion chamber calculated taking into account the volume ₄ Calculating the total temperature T of the outlet of the combustion chamber after considering the volume ₄ (ii) a R in the formula is a gas constant, and the actual total outlet temperature T of the combustion chamber is obtained by considering the volume effect _{32 Shi} Is equal to T ₄ Actual outlet flow Wg of the combustion chamber _{32 Shi} Is equal to the flow W of said considered volume _g4 The calculation of the volume effect is used to reflect the effect of unsteadiness on engine performance.

5. The method of estimating unsteady state performance of a sub-combustion ramjet engine as recited in claim 3, wherein said step 24 comprises applying a total pressure P required for an actual tailpipe ₈ The calculation of (c) further comprises the steps of:

the outlet of the tail nozzle has two states, one state is that the subsonic velocity is compressed to the outlet, and the corresponding static pressure is recorded as P _s1 In the first step, the ultrasonic expansion is carried out to the outlet and then a tail normal shock wave is experienced, and the corresponding static pressure is recorded as P _s2 ；

Actual atmospheric pressure P _{s0 true} At P _s1 And P _s2 And if a normal shock wave exists between the throat part and the outlet, the outlet of the spray pipe is subsonic, and the static pressure of the outlet is equal to the atmospheric pressure, the normal shock wave exists between the throat part and the outlet, and the static pressure is determined according to the static pressure P _s0 And the area A of the outlet of the exhaust nozzle ₄₃ Calculating the actual total pressure P of the outlet of the tail nozzle _{42 fact} Herein, thisThe total pressure of the throat part of the spray pipe is still P ₈ ；

6. The method for estimating the unsteady state performance of the scramjet engine as claimed in claim 3, wherein the entire unsteady state process in the fourth step is calculated without considering the volume effect of the combustion chamber when Time =0, and the total temperature T of the combustion chamber outlet corresponding to the volume considered in the step 23 is calculated ₄ Value of (a) in place of T _{32 Shi} Irrespective of the volumetric flow W _g4 Value of (1) in place of Wg _{32 Shi} 。

7. The method of estimating the unsteady state performance of the scramjet engine as claimed in claim 4, wherein the entire unsteady state process in step four is at Time>At 0, the combustion chamber volume effect is considered, corresponding to the actual combustion chamber outlet total temperature T described in step 23 _{32 Shi} Is equal to the total temperature T of the combustion chamber outlet taking into account the volume ₄ Actual outlet flow Wg of the combustion chamber _{32 fruit} Is equal to the flow W of the volume under consideration _g4 。