CN112832928B - Method for designing cooling structure with equal inner wall strength for rocket engine - Google Patents

Method for designing cooling structure with equal inner wall strength for rocket engine Download PDF

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CN112832928B
CN112832928B CN202110245208.XA CN202110245208A CN112832928B CN 112832928 B CN112832928 B CN 112832928B CN 202110245208 A CN202110245208 A CN 202110245208A CN 112832928 B CN112832928 B CN 112832928B
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wall
thrust chamber
cooling channel
temperature
coolant
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CN112832928A (en
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李龙
李轩
姚卫
汪球
栗继伟
赵伟
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Institute of Mechanics of CAS
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/60—Constructional parts; Details not otherwise provided for
    • F02K9/62—Combustion or thrust chambers
    • F02K9/64—Combustion or thrust chambers having cooling arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/60—Constructional parts; Details not otherwise provided for

Abstract

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 dissipation capacity 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.

Description

Method for designing cooling structure with equal inner wall strength for rocket engine
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 dissipation capacity of the coolant in the small section of cooling channel according to the heat dissipation capacity from the inner side of the inner wall of the thrust chamber to the outer side of the inner wall of the thrust chamber at the corresponding position of any small section of cooling channel and the heat dissipation capacity from the outer side of the inner 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 a coolant in the cooling channels and the heat dissipation capacity of each small section of cooling channel, and therefore each 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:
Figure GDA0003530283810000041
Figure GDA0003530283810000042
Figure GDA0003530283810000043
Figure GDA0003530283810000044
wherein A istIs the throat area, AxIs the area at different positions, TcIs the temperature of the thrust chamber, TxIs the temperature at different locations in the thrust chamber, pcIs the pressure of the thrust chamber, pxIs the pressure at different locations in the thrust chamber, pcDensity of thrust, ρxDensity 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 chamberwgInvolving convective heat transfer ΦkAnd radiative heat transfer phir(ii) a Convective heat transfer phikThe calculation formula of (2) is as follows:
Φk＝hgA(Taw-Twg) (5)
in the above formula, A is the area of any position in the thrust chamber, hgIs the convective heat transfer coefficient, T, of the inner wall of the gas and thrust chamberwgIs the temperature of the inner wall of the thrust chamber, TawThe gas thermal insulation wall temperature for a given position x in the thrust chamber can be determined by the following equation:
Figure GDA0003530283810000051
convective heat transfer coefficient h of inner wall of gas and thrust chambergCalculating according to the Batz formula:
Figure GDA0003530283810000052
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, DtThe 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:
Figure GDA0003530283810000053
for a particular gas mixture, if no Pr and μ data are available, an approximate result can be obtained using the following equation:
Figure GDA0003530283810000054
Figure GDA0003530283810000055
with the radiant heat flow inside the thrust chamber coming from steam and carbon dioxide only, the radiant heat transfer ΦrThe calculation formula of (2) is as follows:
Φr＝εw,efεgσTg 4 (11)
εw,efis the absorption rate of the inner wall surface of the thrust chamber, epsilongEmissivity of gas, TgIs 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, any small section is coldCoolant in cooling channels and total heat dissipation of cooling channels phicf,iEqual to the heat dissipation phi from the gas in the thrust chamber to the inner wall of the thrust chamberwg,iEqual to the heat dissipation phi from the inner wall of the thrust chamber to the outer wall of the thrust chambertw,i: the heat dissipation phi of any small section of cooling channeliComprises 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:
Figure GDA0003530283810000061
wherein h isgIs 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:
Figure GDA0003530283810000062
Twfthe 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 coolantfThe Nu number is usually obtained using the mihajeff formula:
Figure GDA0003530283810000063
wherein: re ═ ρ vfde/μ(15)
ρ is the coolant density, vfFor the flow rate of the cooling liquid in the cooling channel, deIs the hydraulic diameter of the cooling channel, mu is the kinetic viscosity coefficient of the coolant;
Prw＝μcp/λf (16)
cpis the constant pressure specific heat, lambda, of the coolantfIs the thermal conductivity of the coolant; pr (Pr) ofwIs the prandtl number of the coolant at the wall.
Convective heat transfer coefficient h of coolantfComprises the following steps:
Figure GDA0003530283810000071
heat transfer heat flow phi of outer wall of thrust chamber to coolantcfComprises the following steps:
Φcf＝hfA(Twf-Tf) (18)
Tfis 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 dissipation capacity of the coolant in the small section of cooling channel according to the heat dissipation capacity from the inner side of the inner wall of the thrust chamber to the outer side of the inner wall of the thrust chamber at the corresponding position of any small section of cooling channel and the heat dissipation capacity from the outer side of the inner 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 computing capacity, any one of the small sections is taken for research, and the total heat convection from the gas to the inner side of the inner wall of the thrust chamber, the heat dissipation from the inner side of the inner wall of the thrust chamber to the outer side of the inner wall of the thrust chamber and the heat convection from the outer side of the inner 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 the cooling channel, the heat transfer coefficient of the inner wall of the thrust chamber in the small segment is assumed to be constant lambdaiThe heat transfer coefficient of the outer wall of the thrust chamber is constant hiHeat dissipation of coolant and thrust chamber outer wall surface phicf,iComprises the following steps:
Figure GDA0003530283810000081
in the formula, Twf,iIs of Tf.iIs as ac,iFor cooling the channel cross-sectional area, AiThe areas of the inner wall of the thrust chamber and the outer wall of the thrust chamber;
heat radiation phi from inner wall of thrust chamber to outer wall of thrust chambertw,iComprises the following steps:
Figure GDA0003530283810000082
in the formula, Twg,iThe temperature of the inner wall of the ith section;
heat transfer rate phi from gas to inner wall of thrust chamberwg,iComprises the following steps:
Φwg,i＝hg,iAi(Taw,i-Twg,i) (21)
there may be:
Figure GDA0003530283810000083
wherein A isiThe areas of the inner wall and the outer wall of the thrust chamber, Ac,iFor cooling the cross-sectional area of the channel, Taw,iThe temperature of the inner wall of the thrust chamber is constant; the expression is as follows:
Ac,i＝tiLi (23)
Piexpressed as the perimeter of the rib in the cooling channel:
Pi＝2(ti+Li) (24)
when given a phiiThen further obtain:
Φi＝hg,iAi(Taw,i-Twg,i)＝Const (25)
corresponding Twf,iAnd Tf,iThe calculation formula is as follows:
Figure GDA0003530283810000091
Figure GDA0003530283810000092
Taw,iis a constant value when Twg,iWhen different values are taken, the corresponding heat flows will also be different, Twf,iAnd Tf,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 adjusting the hydraulic diameter and the size of each small section under the condition of meeting the hydraulic diameters of the cooling channel and the like by combining the flowing direction of the coolant in the cooling channel and the heat dissipation capacity of each small section of cooling channel 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 pgIn addition, the pressure p of the liquid in the cooling channel is bornefTherefore, 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＝pgR+pfH (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:
Figure GDA0003530283810000101
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:
Figure GDA0003530283810000102
in the formula, σθIs 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:
πR2pc＝Aσz (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:
Figure GDA0003530283810000103
because the inner wall should bear the thermal stress caused by the temperature difference, so there are:
Figure GDA0003530283810000104
according to the fourth theory of strength, there are:
Figure GDA0003530283810000111
wherein:
Figure GDA0003530283810000112
in order to make the intensity of the inner wall of the thrust chamber in any small section constant, namely:
σs,i＝σs,i+1＝Const (37)
then there are:
Figure GDA0003530283810000113
the pressure drop Δ p is based on the cooling channel as per the hydraulic pipe case, etc.:
Figure GDA0003530283810000114
then there are:
Figure GDA0003530283810000115
pf,i+1for the pressure in the i +1 th section of the cooling channel, pf,iFor the pressure in the i-th section of the cooling channel, deIs the hydraulic diameter of the cooling channel;
the following steps are provided:
Figure GDA0003530283810000121
then there are:
Figure GDA0003530283810000122
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 knownf,iThere is the temperature T of the i +1 th segment of coolantf,i+1Comprises the following steps:
Figure GDA0003530283810000123
wherein c isiFor the specific heat capacity of the coolant, MiIs 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.

Claims (10)

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 dissipation capacity of the coolant in the small section of cooling channel according to the heat dissipation capacity from the inner side of the inner wall of the thrust chamber to the outer side of the inner wall of the thrust chamber at the corresponding position of any small section of cooling channel and the heat dissipation capacity from the outer side of the inner 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 a coolant in the cooling channels and the heat dissipation capacity of each small section of cooling channel, and therefore each 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 positionsxThe calculation formula of (a) is as follows:
Figure FDA0003530283800000021
temperature T at different locationsxThe calculation formula of (a) is as follows:
Figure FDA0003530283800000022
pressure p at different locationsxThe calculation formula of (a) is as follows:
Figure FDA0003530283800000023
density p at different positionsxThe calculation formula of (a) is as follows:
Figure FDA0003530283800000024
in the formula, AtIs throat area, TcIs the temperature of the combustion chamber, pcIs the pressure of the combustion chamber, pcIs 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 ΦkThe calculation process of (2) is as follows:
Φk＝hgA(Taw-Twg) (5)
in the above formula, A is the cross-sectional area of the thrust chamber at the current position, TwgIs the inner wall temperature, T, of the current positionawThe gas thermal insulation wall temperature for a given position x in the thrust chamber can be determined by the following equation:
Figure FDA0003530283800000025
hgthe convective heat transfer coefficient of the fuel gas and the inner wall surface can be calculated according to the Batz formula:
Figure FDA0003530283800000031
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:
Figure FDA0003530283800000032
said radiative heat transfer ΦrThe calculation formula of (2) is as follows:
Φr＝εw,efεgσTg 4
the total heat flow of the gas to the wall surface is phiwg：
Figure FDA0003530283800000033
Wherein h isgIs the overall effective heat transfer coefficient after conversion.
5. The design method according to claim 4,
in calculating the convective heat transfer ΦkFor a particular fuel gas mixture, the data for Pr and μ, taken with the following formula, yields an approximate result:
Figure FDA0003530283800000034
Figure FDA0003530283800000035
in calculating the radiant heat transfer phirThe calculation is based on the radiative heat transfer inside the thrust chamber from water vapor and carbon dioxide.
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 dissipation Φ from the inner side of the inner wall of the thrust chamber to the outer side of the inner wall of the thrust chambertw,iThe calculation formula is as follows:
Figure FDA0003530283800000041
Aithe areas of the inner wall and the outer wall of the thrust chamber corresponding to the small section of the cooling channel, Twg,iThe inner side temperature T of the inner wall of the thrust chamber at the corresponding position of the small section of cooling channelwfIs the temperature outside the inner wall of the segment, deltaiThe wall thickness of the thrust chamber;
heat radiation amount phi of heat radiation amount from gas to outer side of inner wall of thrust chamberwg,iThe calculation formula is as follows:
Φwg,i＝hg,iAi(Taw,i-Twg,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 dissipation phi of the thrust chamber outer wallcf,iThe calculation formula is as follows:
Figure FDA0003530283800000042
setting the temperature phi of the thermal insulation wall at the inner wall of the thrust chamber corresponding to the small section of cooling channeliFully conducting, the coolant in the small section of the cooling channel and the total heat dissipation phi of the cooling channelcf,iEqual to the heat dissipation phi from the gas in the thrust chamber to the inner wall of the thrust chamberwg,iEqual to the heat dissipation phi from the inner wall of the thrust chamber to the outer wall of the thrust chambertw,i: the heat dissipation phi of the small section of cooling channeliComprises the following steps:
Figure FDA0003530283800000043
wherein A isiThe areas of the inner wall and the outer wall of the thrust chamber, Ac,iFor cooling the cross-sectional area of the channel, Taw,iIf the temperature of the fuel gas is a constant value, further obtaining:
Φi＝hg,iAi(Taw,i-Twg,i)＝Const (23)
wherein h isgTo the reduced total effective heat transfer coefficient, Twg,iThe temperature, T, of the inner wall of the thrust chamber corresponding to the small section of the cooling passagef,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 channelwf,iAnd the temperature T of the coolant in the small section of the cooling passagef,iThe calculation process is as follows:
Figure FDA0003530283800000051
Figure FDA0003530283800000052
due to Taw,iIs a constant value when Twg,iWhen different values are taken, the corresponding heat flows will also be different, so Twf,iAnd Tf,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 pgIn addition, the pressure p of the liquid in the cooling channel is bornefTherefore, 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＝pgR+pfH
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:
Figure FDA0003530283800000061
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:
Figure FDA0003530283800000062
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: sigmas,i＝σs,i+1When Const, there are:
Figure FDA0003530283800000063
the pressure drop Δ p is based on the cooling channel as per the hydraulic pipe case, etc.:
Figure FDA0003530283800000064
then there are:
Figure FDA0003530283800000071
pf,i+1pressure of i +1 th segment in cooling channel, pf,iFor the pressure of the i-th section in the cooling channel, deHydraulic power for cooling channelsA diameter;
the following steps are provided:
Figure FDA0003530283800000072
then there are:
Figure FDA0003530283800000073
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 knownf,iThere is the temperature T of the i +1 th segment of coolantf,i+1Comprises the following steps:
Figure FDA0003530283800000074
wherein c isiFor the specific heat capacity of the coolant, MiIs the mass flow rate of the coolant.
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