阅读说明：本技术 一种用于火箭发动机的等内壁强度的冷却结构设计方法 (Method for designing cooling structure with equal inner wall strength for rocket engine ) 是由 李龙 李轩 姚卫 汪球 栗继伟 赵伟 于 2021-03-05 设计创作，主要内容包括：本发明提供一种用于火箭发动机的等内壁强度的冷却结构设计方法,包括：计算出推力室轴向上不同位置处工作时的燃气参数和绝热壁温；根据冷却剂的类型确定冷却通道围绕推力室的布置方式；由冷却通道入口至出口,计算任意一小段冷却通道中冷却剂的导热量；以整个推力室内壁强度为一个恒定值作为基础,在满足冷却通道等水力直径下调整各小段的水力直径和尺寸；通过循环迭代,以满足此条件的冷却通道形状作为设计结果,完成设计过程。本发明以推力室内壁强度为等强度的恒定值作为基础,在满足冷却通道等水利直径的条件下,将传统的冷却通道进行优化,通道内压力损失小,结构简单、可靠,轻量化作用明显可以有效的减轻发动机的质量。(The invention provides a method for designing a cooling structure with equal inner wall strength for a rocket engine, which comprises the following steps: calculating gas parameters and heat insulation wall temperature of the thrust chamber in different axial positions during working; determining the arrangement mode of the cooling channel around the thrust chamber according to the type of the coolant; calculating the heat conduction quantity of the coolant in any small section of cooling channel from the inlet to the outlet of the cooling channel; taking the strength of the inner wall of the whole thrust chamber as a constant value as a basis, and adjusting the hydraulic diameter and the size of each small section under the condition of meeting the hydraulic diameters of a cooling channel and the like; the design process is completed by loop iteration with the cooling channel shape that satisfies this condition as the design result. The invention takes the constant value of the strength of the inner wall of the thrust chamber as the equal strength as the basis, optimizes the traditional cooling channel under the condition of meeting the water conservancy diameters of the cooling channel and the like, has small pressure loss in the channel, simple and reliable structure and obvious light weight effect, and can effectively reduce the mass of the engine.)

1. A method for designing a cooling structure with equal inner wall strength for a rocket engine is characterized by comprising the following steps:

step 100, firstly, calculating gas parameters of the thrust chamber during working at different positions in the axial direction on the basis of the shape of the thrust chamber of the rocket engine to obtain the heat insulation wall temperature of the inner wall of the thrust chamber;

step 200, according to the type of the coolant, determining the arrangement mode, the shape and the inlet and outlet positions of the cooling channel around the thrust chamber in a conventional mode in the flow rate and the flow path mode of the coolant;

step 300, dividing the cooling channel into a plurality of small sections along the flow direction from the inlet to the outlet of the cooling channel, and calculating the heat conduction quantity of the coolant in the small section of the cooling channel according to the heat conduction quantity from the inner wall of the thrust chamber to the outer wall of the thrust chamber at the corresponding position of any small section of the cooling channel and the heat conduction quantity from the outer wall of the thrust chamber to the coolant;

step 400, based on the fact that the strength of the inner wall of the whole thrust chamber is a constant value, the hydraulic diameter and the size of each small section are adjusted under the condition that the hydraulic diameters of the cooling channels and the like are met by combining the flowing direction of the coolant in the cooling channels and the heat conduction quantity of each small section of cooling channel, and the small section of cooling channel meets the heat dissipation requirement of the inner wall temperature of the corresponding small section of thrust chamber;

step 500, in the adjusting process, the size of each small section of the cooling channel is adjusted by repeating steps 200 to 400 through loop iteration to change the heat exchange quantity of the small sections until the heat dissipation quantity of all the small sections enables the temperature of the inner wall of the thrust chamber to be consistent, and then the shape and the size of the cooling channel meeting the condition are taken as a design result to complete the design process.

2. The design method according to claim 1,

in the step 100, the gas parameters are as follows: mach number, temperature, pressure, and density variation data occurring at different locations along the flow of fuel axially inside the thrust chamber.

3. The design method according to claim 2,

area A at different positions_xThe calculation formula of (a) is as follows:

temperature T at different locations_xThe calculation formula of (a) is as follows:

pressure p at different locations_xThe calculation formula of (a) is as follows:

density p at different positions_xThe calculation formula of (a) is as follows:

in the formula, A_tIs throat area, T_cIs the temperature of the combustion chamber, p_cIs the pressure of the combustion chamber, p_cIs the density of the combustion chamber.

4. The design method according to claim 1,

in the step 100, the adiabatic wall temperature of the inner wall of the thrust chamber is obtained by respectively calculating and adding the results of the convective heat transfer and the radiative heat transfer after only the convective heat transfer and the radiative heat transfer are assumed to exist between the fuel gas in the thrust chamber and the inner wall of the thrust chamber;

said convective heat transfer Φ_kThe calculation process of (2) is as follows:

Φ_k＝h_gA(T_aw-T_wg) (5)

in the above formula, A is the cross-sectional area of the thrust chamber at the current position, T_wgIs the inner wall temperature, T, of the current position_awThe gas thermal insulation wall temperature for a given position x in the thrust chamber can be determined by the following equation:

h_gthe convective heat transfer coefficient of the fuel gas and the inner wall surface can be calculated according to the Batz formula:

in the formula, R is the curvature radius of the nozzle at the throat part, sigma is a correction parameter considering the performance change of the gas in the boundary layer, and can be determined according to the stagnation temperature of the nozzle, the current thrust space inner wall temperature and the current Mach number:

said radiative heat transfer Φ_rThe calculation formula of (2) is as follows:

Φ_r＝ε_w,efε_gσT_g ⁴

the temperature of the heat insulating wall phi_wgComprises the following steps:

wherein h is_gIs the overall effective heat transfer coefficient after conversion.

5. The design method according to claim 4,

in calculating the convective heat transfer Φ_kFor a particular fuel gas mixture, the data for Pr and μ, taken with the following formula, yields an approximate result:

in calculating the radiant heat transfer phi_rWhile the heat transfer by radiation inside the thrust chamber comes from water vapor and dioxygenCarbon is used as a basis for calculation.

6. The design method according to claim 1,

in the step 200, the cooling channel is a plurality of independent channels formed by fins in an isolation manner, each independent channel is uniformly distributed around the periphery of the thrust chamber, and the cross section of the cooling channel is in a fan shape; the cooling passage has an inlet at the fuel outlet end of the thrust chamber and an outlet at the fuel inlet end of the thrust chamber.

7. The design method according to claim 1,

in step 300, the heat transfer phi from the inner wall of the thrust chamber to the outer wall of the thrust chamber_tw,iThe calculation formula is as follows:

A_ithe areas of the inner wall and the outer wall of the thrust chamber corresponding to the small section of the cooling channel, T_wg,iThe temperature, T, of the inner wall of the thrust chamber at the location corresponding to the small section of the cooling passage_wfIs the outer wall temperature, delta, of the segment_iThe wall thickness of the thrust chamber;

heat transfer phi from the outer wall of the thrust chamber to the coolant_wg,iThe calculation formula is as follows:

Φ_wg,i＝h_g,iA_i(T_aw,i-T_wg,i) (21)

the calculation process of the heat exchange amount of the coolant of any small section of cooling channel is as follows:

coolant and heat transfer phi of the thrust chamber outer wall surface_cf,iThe calculation formula is as follows:

thrust at the corresponding position of the small section of cooling channelHeat insulation wall temperature phi of indoor wall_iFully conducting, the total heat transfer phi of the coolant and cooling channels in the small section of cooling channel_cf,iEqual to the heat conduction quantity phi from the gas in the thrust chamber to the inner wall of the thrust chamber_wg，iEqual to the heat transfer phi from the inner wall of the thrust chamber to the outer wall of the thrust chamber_tw,i: the heat conductivity phi of the small section of cooling channel_iComprises the following steps:

wherein A is_iThe areas of the inner wall and the outer wall of the thrust chamber, A_c,iFor cooling the cross-sectional area of the channel, T_aw,iThe temperature of the inner wall of the thrust chamber is a constant value, and then further obtained:

Φ_i＝h_g,iA_i(T_aw,i-T_wg,i)＝Const (23)

wherein h is_gTo the reduced total effective heat transfer coefficient, T_wg,iThe temperature, T, of the outer wall of the thrust chamber corresponding to the small section of the cooling passage_f,iThe temperature of the coolant in the small section of cooling passage.

8. The design method according to claim 7,

wherein, the T of the outer wall temperature of the thrust chamber corresponding to the small section of cooling channel_wf,iAnd the temperature T of the coolant in the small section of the cooling passage_f,iThe calculation process is as follows:

due to T_aw,iIs a constant value when T_wg,iTaking different values, then phaseThe required heat flux will also vary, hence T_wf,iAnd T_f,iAnd varies according to the size of the cooling passage.

9. The design method according to claim 1,

in step 400, the inner wall of the thrust chamber is subjected to the gas pressure p_gIn addition, the pressure p of the liquid in the cooling channel is borne_fTherefore, the force balance equation of the inner wall of the thrust chamber is as follows, and the force calculation mode of the inner wall of the thrust chamber is as follows:

σ₁δ₁+σ₂δ₂＝p_gR+p_fH

setting the inner wall sigma of the thrust chamber₁And cooling channel outer wall sigma₂The strain values are the same, and the tensile stress on the two is respectively as follows:

because the inner wall of the thrust chamber will bear the thermal stress caused by the temperature difference, the inner side is pressed, and can be deduced to obtain:

wherein H is the cooling channel height, E₁Is the elastic modulus of the material of the inner wall surface, a₁Is the linear expansion coefficient of the material of the inner wall surface, q is the heat flow along the vertical direction of the wall surface, upsilon is the Poisson ratio, and lambda is₁The thermal conductivity of the material of the inner wall surface.

10. The design method according to claim 9,

in step 400, the calculation process for making the strength of any small section of the inner wall of the thrust chamber constant is as follows:

to make: sigma_s,i＝σ_s,i+1When Const, there are:

the pressure drop Δ p is based on the cooling channel as per the hydraulic pipe case, etc.:

then there are:

p_f,i+1pressure of i +1 th segment in cooling channel, p_f,iFor the pressure of the i-th section in the cooling channel, d_eIs the hydraulic diameter of the cooling channel;

the following steps are provided:

then there are:

where m is the mass flow of the coolant, R₁To cool the channel bottom radius, S₁For the length of the arc at the bottom of the cooling channel, R₂To cool the channel top radius, S₂The length of the arc at the top of the cooling channel is defined, and rho is the density of the coolant;

the temperature of the coolant flowing from the inlet of the cooling passage increases for every small section of the cooling passage, when the temperature T of the i-th small section of the coolant is known_f,iThere is the temperature T of the i +1 th segment of coolant_f,i+1Comprises the following steps:

wherein c is_iM is the specific heat capacity of the coolant_iIs the mass flow rate of the coolant.

Technical Field

The invention relates to the field of rocket engines, in particular to a method for designing a cooling structure with equal inner wall strength for a rocket engine.

Background

The pressure of the thrust chamber of the liquid rocket engine is high, the heat flux density of the wall surface of the thrust chamber is high, and the temperature of fuel gas can reach thousands of degrees centigrade, which exceeds the temperature which can be born by common engine materials. However, the wall of the thrust chamber allows much smaller heat flows, and if necessary protection measures are not taken, the temperature of the wall of the thrust chamber is too high under the severe conditions, and the wall of the thrust chamber is even burnt.

In order to carry out thermal protection, a cooling channel is arranged between the inner wall surface and the outer wall surface of the engine, a cooling medium flows at high speed in the cooling channel, absorbs heat and heats up, and actively cools the engine structure to protect the engine structure from being ablated and damaged.

The cooling channel of the traditional engine is designed into a rectangular groove structure with a uniform section or a simple variable section, and large fluid pressure loss is easily caused.

Disclosure of Invention

The invention aims to provide a design method of a cooling structure with equal inner wall strength for a rocket engine.

Specifically, the invention provides a method for designing a cooling structure with equal inner wall strength for a rocket engine, which comprises the following steps:

step 100, firstly, calculating gas parameters of the thrust chamber during working at different positions in the axial direction on the basis of the shape of the thrust chamber of the rocket engine to obtain the heat insulation wall temperature of the inner wall of the thrust chamber;

step 200, according to the type of the coolant, determining the arrangement mode, the shape and the inlet and outlet positions of the cooling channel around the thrust chamber in a conventional mode in the flow rate and the flow path mode of the coolant;

step 300, dividing the cooling channel into a plurality of small sections along the flow direction from the inlet to the outlet of the cooling channel, and calculating the heat conduction quantity of the coolant in the small section of the cooling channel according to the heat conduction quantity from the inner wall of the thrust chamber to the outer wall of the thrust chamber at the corresponding position of any small section of the cooling channel and the heat conduction quantity from the outer wall of the thrust chamber to the coolant;

step 400, based on the fact that the strength of the inner wall of the whole thrust chamber is a constant value, the hydraulic diameter and the size of each small section are adjusted under the condition that the hydraulic diameters of the cooling channels and the like are met by combining the flowing direction of the coolant in the cooling channels and the heat conduction quantity of each small section of cooling channel, and the small section of cooling channel meets the heat dissipation requirement of the inner wall temperature of the corresponding small section of thrust chamber;

step 500, in the adjusting process, the size of each small section of the cooling channel is adjusted by repeating steps 200 to 400 through loop iteration to change the heat exchange quantity of the small sections until the heat dissipation quantity of all the small sections enables the temperature of the inner wall of the thrust chamber to be consistent, and then the shape and the size of the cooling channel meeting the condition are taken as a design result to complete the design process.

Based on the difference of the on-way section of the thrust chamber and the heat dissipation environment, the section of the cooling channel can be changed along with the change of the section of the thrust chamber, the strength of the inner wall of the thrust chamber is taken as a constant value with equal strength as a basis, and the traditional cooling channel is optimized under the condition of meeting the water conservancy diameters of the cooling channel and the like, so that the pressure loss in the channel is small, the structure is simple and reliable, the light weight effect is obvious, and the mass of an engine can be effectively reduced.

Drawings

FIG. 1 is a schematic illustration of the steps of a design process according to one embodiment of the present invention;

FIG. 2 is a schematic view of a thrust chamber configuration according to an embodiment of the present invention;

FIG. 3 is a schematic view of a cooling channel configuration according to an embodiment of the present invention;

FIG. 4 is a schematic heat transfer diagram of a cooling channel according to an embodiment of the present invention;

FIG. 5 is a schematic view of a cooling gallery with hoop stress according to an embodiment of the present invention;

FIG. 6 is a schematic view of an embodiment of the present invention showing the cooling passage in an axially stressed state (left) and in cross-sectional area (right);

FIG. 7 is a flow chart of a cooling channel design according to an embodiment of the present invention.

Detailed Description

The detailed structure and implementation process of the present solution are described in detail below with reference to specific embodiments and the accompanying drawings.

Generally, the larger the depth-to-width ratio of the cooling channel is, the better the cooling effect is, the prior art is a research under the condition of unchanged width, but since the convective heat transfer coefficient is directly in an inverse relation with the hydraulic diameter, the research on the influence of the change of the shape of the cooling channel or the height-to-width ratio by fixing the hydraulic diameter of the cooling channel is a more meaningful direction.

The cooling channel absorbs the heat of the wall surface of the thrust chamber to reduce the temperature of the wall surface, and under the condition that the width of the cooling channel is unchanged, along with the increase of the depth-to-width ratio of the cooling channel, the cooling effect of the wall surface of the thrust chamber is gradually improved, and the temperature of the inner wall surface on the gas side is gradually reduced; the pressure difference between the inlet and the outlet of the cooling channel gradually rises, and the pressure difference between the inlet and the outlet gradually increases. However, when the aspect ratio is increased to a certain extent, the cooling effect tends to saturate, since the negative effect on the heat transfer, which is a decrease in the area of convective heat transfer in the channels, gradually outweighs the positive effects of increasing the rib efficiency and increasing the flow rate of the cooling medium. On the basis, the influence of the height and the width of the cooling channel is researched by controlling the hydraulic diameter of the cooling channel, so that the heat on the wall surface of the thrust chamber is effectively taken away, the pressure loss of fuel in the cooling channel is reduced, and the structural weight and the thermal stress of the engine are reduced.

As shown in fig. 1, in one embodiment of the present invention, there is provided a method for designing an equi-inner-wall-strength cooling structure for a rocket engine, comprising the steps of:

step 100, firstly, calculating gas parameters of the thrust chamber during working at different positions in the axial direction on the basis of the shape of the thrust chamber of the rocket engine to obtain the heat insulation wall temperature of the inner wall of the thrust chamber;

as shown in fig. 2, when the gas flows in the thrust chamber, the relevant parameters are changed along the axial direction, assuming that the total temperature and the total pressure of the gas are constant, and the gas is an isentropic flow process, the mach numbers, the temperatures, the pressures and the densities at different positions can be deduced according to a one-dimensional adiabatic isentropic formula, and the calculation formula is as follows:

wherein A is_tIs the throat area, A_xIs the area at different positions, T_cIs the temperature of the thrust chamber, T_xIs the temperature at different locations in the thrust chamber, p_cIs the pressure of the thrust chamber, p_xIs the pressure at different locations in the thrust chamber, p_cDensity of thrust, ρ_xDensity at different locations.

The flow in the cooling passage is three-dimensional, steady-state, and turbulent, and it is necessary to consider both the change in the physical properties of the fluid and the change in the physical properties of the metal wall surface. The structure of cooling passage is shown in fig. 3, the structure of cooling passage can be regarded as a fin device, the temperature distribution of the process of gas flowing through the inner wall of the thrust chamber to transfer heat to the coolant in the cooling passage is schematically shown in fig. 4, and the specific heat transfer process is divided into three steps: the gas transfers heat to the inner wall of the thrust chamber in the thrust chamber, the inner wall of the thrust chamber transfers heat to the outer wall of the thrust chamber, and the outer wall of the thrust chamber transfers heat to the coolant in the cooling channel.

The gas transfers heat to the inner wall of the thrust chamber in a radiation and convection heat exchange mode, so that heat flow phi transferred to the inner wall of the thrust chamber_wgInvolving convective heat transfer Φ_kAnd radiative heat transfer phi_r(ii) a Convective heat transfer phi_kThe calculation formula of (2) is as follows:

Φ_k＝h_gA(T_aw-T_wg) (5)

in the above formula, A is the area of any position in the thrust chamber, h_gIs the convective heat transfer coefficient, T, of the inner wall of the gas and thrust chamber_wgIs the temperature of the inner wall of the thrust chamber, T_awThe gas thermal insulation wall temperature for a given position x in the thrust chamber can be determined by the following equation:

convective heat transfer coefficient h of inner wall of gas and thrust chamber_gCalculating according to the Batz formula:

wherein Pr is the prandtl number of the fuel gas, g is the gravitational acceleration, c^*The characteristic speed of the thrust chamber, which is the thrust chamber, is a fixed value if the engine state is determined, D_tThe diameter of a throat of the thrust chamber, R is the curvature radius of a spray pipe at the throat of the thrust chamber, and sigma is a correction parameter considering the performance change of gas in the boundary layer, and can be determined according to the stagnation temperature of the spray pipe, the inner wall temperature of the thrust chamber at the current position and the Mach number at the current position:

for a particular gas mixture, if no Pr and μ data are available, an approximate result can be obtained using the following equation:

with the radiant heat flow inside the thrust chamber coming from steam and carbon dioxide only, the radiant heat transfer Φ_rThe calculation formula of (2) is as follows:

Φ_r＝ε_w,efε_gσT_g ⁴ (11)

ε_w,efis the absorption rate of the inner wall surface of the thrust chamber, epsilon_gEmissivity of gas, T_gIs the gas temperature.

If the inner wall of the thrust chamber is arranged to face the outer wall of the thrust chamber, and the temperature of the outer wall of the thrust chamber to the coolant is completely conducted, the coolant in any small section of the cooling channel and the total heat conduction quantity phi of the cooling channel_cf,iEqual to the heat conduction quantity phi from the gas in the thrust chamber to the inner wall of the thrust chamber_wg,iEqual to the heat transfer phi from the inner wall of the thrust chamber to the outer wall of the thrust chamber_tw,i: the heat conduction phi of any small section of cooling channel_iComprises the following steps:

the total heat flow of the heat transfer of the fuel gas to the inner wall of the thrust chamber is as follows:

wherein h is_gIs the overall effective heat transfer coefficient after conversion.

In practice, the convective heat transfer of the gas is the main form of heat transfer from the gas in the thrust chamber to the inner wall of the thrust chamber, and in the thrust chamber, the convective heat flow usually accounts for more than 80% of the total heat flow, and can reach 95% near the throat and more than 98% downstream of the nozzle.

The expression of the heat conduction process from the inner wall of the thrust chamber to the outer wall of the thrust chamber is as follows:

T_wfthe temperature of the outer wall surface of the cooling channel of the thrust chamber is reduced, the thickness delta of the side wall of the whole thrust chamber is reduced, or the temperature of the inner wall of the thrust chamber can be effectively reduced by adopting a material with good heat-conducting property.

Convective heat transfer coefficient h from outer wall of thrust chamber to coolant_fThe Nu number is usually obtained using the mihajeff formula:

wherein: re ═ ρ v_fd_e/μ (15)

ρ is the coolant density, v_fFor the flow rate of the cooling liquid in the cooling channel, d_eIs the hydraulic diameter of the cooling channel, mu is the kinetic viscosity coefficient of the coolant;

Pr_w＝μc_p/λ_f (16)

c_pis the constant pressure specific heat, lambda, of the coolant_fIs the thermal conductivity of the coolant; pr (Pr) of_wIs the prandtl number of the coolant at the wall.

Convective heat transfer coefficient h of coolant_fComprises the following steps:

heat transfer heat flow phi of outer wall of thrust chamber to coolant_cfComprises the following steps:

Φ_cf＝h_fA(T_wf-T_f) (18)

T_fis the coolant temperature.

Step 200, according to the type of the coolant, determining the arrangement mode, the shape and the inlet and outlet positions of the cooling channel around the thrust chamber in a conventional mode in the flow rate and the flow path mode of the coolant;

the cooling channel is a plurality of independent channels formed by ribs in an isolation mode, the independent channels are uniformly distributed around the periphery of the thrust chamber, and the section of the cooling channel in the embodiment is in a fan shape; the inlet of the cooling passage is located at the fuel outlet end of the thrust chamber, and the outlet is located at the fuel inlet end of the thrust chamber.

Step 300, dividing the cooling channel into a plurality of small sections along the flow direction from the inlet to the outlet of the cooling channel, and calculating the heat conduction quantity of the coolant in the small section of the cooling channel according to the heat conduction quantity from the inner wall of the thrust chamber to the outer wall of the thrust chamber at the corresponding position of any small section of the cooling channel and the heat conduction quantity from the outer wall of the thrust chamber to the coolant;

in the present embodiment, the rule for dividing the cooling passages is as follows: the cooling channel is uniformly divided into a plurality of small segments along the way, the smaller the size of each small segment is, the larger the calculation amount is, therefore, the length of each small segment can be generally taken as 1mm, and the repeated calculation can be taken as 0.5mmm after the initial result is obtained. Wherein if the computer performance is strong, each small segment can be 0.1mm or even smaller.

After the cooling channel is divided into a plurality of small sections along the way according to the calculation capacity, any one of the small sections is taken for research, and the total heat convection from the gas to the inner wall of the thrust chamber, the heat conduction from the inner wall of the thrust chamber to the outer wall of the thrust chamber and the heat convection from the outer wall of the thrust chamber to the coolant are comprehensively considered. The temperature of the inner wall of the thrust chamber is reduced by absorbing the heat of the wall surface through the cooling channel, and the cooling effect on the inner wall of the thrust chamber is gradually improved along with the increase of the depth-to-width ratio of the cooling channel, so that the temperature of the inner wall of the thrust chamber is gradually reduced; the pressure difference between the inlet and the outlet of the cooling channel gradually rises (because the temperature of the coolant at the inlet is low, the temperature rises after the coolant absorbs heat at the outlet), and the pressure difference between the inlet and the outlet gradually increases. However, when the aspect ratio is increased to a certain extent, the cooling effect tends to saturate, since the negative effect on heat transfer is gradually outweighed by the positive effect on increasing the rib efficiency and increasing the flow rate of the cooling medium, due to the reduced area of convective heat transfer in the thrust chamber channels.

For any small segment of cooling channel, the thermal conductivity of the inner wall of the thrust chamber in the small segment is assumed to be constant lambda_iThe heat transfer coefficient of the outer wall of the thrust chamber is constant h_iHeat dissipation of coolant and thrust chamber outer wall surface phi_cf,iComprises the following steps:

in the formula, T_wf,iIs of T_f.iIs as a_c,iFor cooling the channel cross-sectional area, A_iThe areas of the inner wall of the thrust chamber and the outer wall of the thrust chamber;

heat conduction from inner wall of thrust chamber to outer wall of thrust chamber_tw,iComprises the following steps:

in the formula, T_wg,iThe temperature of the inner wall of the ith section;

heat transfer rate phi from gas to inner wall of thrust chamber_wg,iComprises the following steps:

Φ_wg,i＝h_g,iA_i(T_aw,i-T_wg,i) (21)

there may be:

wherein A is_iThe areas of the inner wall and the outer wall of the thrust chamber, A_c,iFor cooling the cross-sectional area of the channel, T_aw,iThe temperature of the inner wall of the thrust chamber is constant; the expression is as follows:

A_c,i＝t_iL_i (23)

P_iexpressed as the perimeter of the rib in the cooling channel:

P_i＝2(t_i+L_i) (24)

when given a phi_iThen further obtain:

Φ_i＝h_g,iA_i(T_aw,i-T_wg,i)＝Const (25)

corresponding T_wf,iAnd T_f,iThe calculation formula is as follows:

T_aw,iis a constant value when T_wg,iWhen different values are taken, the corresponding heat flows will also be different, T_wf,iAnd T_f,iAnd varies according to the size of the cooling structure.

Step 400, taking the wall temperature of the inner wall of the whole thrust chamber as a constant value as a basis, and combining the flowing direction of a coolant in a cooling channel and the heat conduction quantity of each small section of cooling channel, adjusting the hydraulic diameter and size of each small section of cooling channel under the condition of meeting the hydraulic diameter of the cooling channel and the like, so that each small section of cooling channel meets the heat dissipation requirement of the wall temperature of the inner wall of the corresponding small section of thrust chamber;

the heat dissipation requirement here means: the obtained temperature of the outer wall of each small section of the thrust chamber, the temperature of the inlet and the outlet of the coolant, and the heat flow and the heat exchange coefficient need to be taken away by the cooling liquid completely by transferring the gas in the thrust chamber to the inner wall of the thrust chamber, and the heat exchange between the outer wall of the cooling channel and the environment is not considered.

As shown in FIG. 5, the inner wall of the thrust chamber is subjected to a gas pressure p_gIn addition, the pressure p of the liquid in the cooling channel is borne_fTherefore, the force balance equation of the inner wall of the thrust chamber is as follows, and the force calculation mode of the inner wall of the thrust chamber is as follows:

σ₁δ₁+σ₂δ₂＝p_gR+p_fH (28)

setting the inner wall sigma of the thrust chamber₁And cooling channel outer wall sigma₂The strain values are the same, and the tensile stress on the two is respectively as follows:

because the thermal stress that the inner wall of thrust chamber will bear the difference in temperature and cause, inboard pressurized can derive the thermal stress that obtains the inner wall face of thrust chamber to be:

in the formula, σ_θIs the thermal stress of the inner wall surface of the thrust chamber, H is the height of the cooling passage, E₁Is the elastic modulus of the material of the inner wall surface, a₁Is the linear expansion coefficient of the material of the inner wall surface, q is the heat flow along the vertical direction of the wall surface, upsilon is the Poisson ratio, and lambda is₁The thermal conductivity of the material of the inner wall surface.

As shown in the left diagram of fig. 6, the thrust chamber is balanced in axial force by:

πR²p_c＝Aσ_z (31)

where a is the cross-sectional area of the cooling structure, as shown in the right diagram of fig. 6, is:

A＝πRδ₁+π(R+δ₁+H)δ₂+ntH (32)

then there are:

because the inner wall should bear the thermal stress caused by the temperature difference, so there are:

according to the fourth theory of strength, there are:

wherein:

in order to make the intensity of the inner wall of the thrust chamber in any small section constant, namely:

σ_s,i＝σ_s,i+1＝Const (37)

then there are:

the pressure drop Δ p is based on the cooling channel as per the hydraulic pipe case, etc.:

then there are:

p_f,i+1for the pressure in the i +1 th section of the cooling channel, p_f,iFor the pressure in the i-th section of the cooling channel, d_eIs the hydraulic diameter of the cooling channel;

the following steps are provided:

then there are:

where m is the mass flow of the coolant, R₁To cool the channel bottom radius, S₁For the length of the arc at the bottom of the cooling channel, R₂To cool the channel top radius, S₂The length of the arc at the top of the cooling channel is defined, and rho is the density of the coolant;

the coolant flowing from the inlet of the cooling channel cools each time it flows through a small section of the cooling channelThe temperature of the coolant will rise when the temperature T of the i-th section of coolant is known_f,iThere is the temperature T of the i +1 th segment of coolant_f,i+1Comprises the following steps:

wherein c is_iM is the specific heat capacity of the coolant_iIs the mass flow rate of the coolant.

When the coolant flows in from the inlet of the cooling channel at the downstream of the thrust chamber, flows along the cooling channel and flows out from the outlet of the cooling channel at the front end of the thrust chamber, along with the flowing process, the temperature of the fuel gas in the thrust chamber is gradually increased, the diameter of the thrust chamber is firstly reduced and then increased, the flowing condition inside the thrust chamber is changed, so that the heat flow at each position where the coolant flows is different, the temperature of the inner wall surface of the thrust chamber is consistent based on the design, and therefore the sizes of the cooling channel, the outer wall of the thrust chamber and the fin structure are required to be adjusted and adjusted simultaneously to meet the requirement.

Step 500, in the adjusting process, the steps 200 to 400 are repeated through loop iteration to adjust the size of each small section of the cooling channel so as to change the heat exchange quantity of the small section until the heat dissipation quantity of all the small sections enables the temperature of the inner wall of the whole thrust chamber to be consistent, and then the size of the cooling channel meeting the condition is taken as a design result to complete the design process.

In this step, the precondition that the temperature of the inner wall of the thrust chamber is kept uniform is that the conditions defined in step 400 are satisfied: the strength of the inner wall of the whole thrust chamber is a constant value.

Based on the difference of the on-way section of the thrust chamber and the heat dissipation environment, the section of the cooling channel can be changed along with the change basis, the strength of the inner wall of the thrust chamber is taken as a constant value of equal strength as a basis, the traditional cooling channel is optimized under the condition of meeting the water conservancy diameters of the cooling channel and the like, the pressure loss in the channel is small, the structure is simple and reliable, and the weight reduction effect is obvious, so that the quality of the engine can be effectively reduced.

The flow of the cooling channel design is only briefly described in the following text.

As shown in fig. 7, firstly, calculating the gas parameters at each position of the thrust chamber axis, setting the coolant flow and the coolant flow according to the gas parameters, adjusting the size and the shape of the cooling channel under the condition of meeting the equal water conservancy diameter with the determined strength of the inner wall of the thrust chamber as a constant value, iteratively calculating the total heat flow, the coolant temperature and the gas side wall temperature (inner wall of the thrust chamber) and the coolant side wall temperature (outer wall of the thrust chamber) at different positions (any small section) of the cooling channel, checking the iteration result of each time, if the heat dissipation capacity at a certain position fails to keep the strength of the inner wall of the thrust chamber consistent, keeping the inner wall temperature of the thrust chamber at the position uniform with the inner wall temperature of the whole thrust chamber, returning to adjust the flow path or the flow of the coolant, or adjusting the size of the cooling channel at the position until the inner wall temperatures of the thrust chambers at all positions keep consistent, and then designing a cooling channel according to the size of the cooling structure at the moment to obtain the cooling channel structure of the isothermal rocket engine with the equal inner wall required by the invention.

Thus, it should be appreciated by those skilled in the art that while a number of exemplary embodiments of the invention have been illustrated and described in detail herein, many other variations or modifications consistent with the principles of the invention may be directly determined or derived from the disclosure of the present invention without departing from the spirit and scope of the invention. Accordingly, the scope of the invention should be understood and interpreted to cover all such other variations or modifications.