CN116562194B - Thrust evaluation method and system for ramjet rotary detonation engine - Google Patents
Thrust evaluation method and system for ramjet rotary detonation engine Download PDFInfo
- Publication number
- CN116562194B CN116562194B CN202310835905.XA CN202310835905A CN116562194B CN 116562194 B CN116562194 B CN 116562194B CN 202310835905 A CN202310835905 A CN 202310835905A CN 116562194 B CN116562194 B CN 116562194B
- Authority
- CN
- China
- Prior art keywords
- combustion chamber
- outlet
- calculating
- engine
- formula
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
- 238000005474 detonation Methods 0.000 title claims abstract description 24
- 238000011156 evaluation Methods 0.000 title claims abstract description 19
- 238000002485 combustion reaction Methods 0.000 claims abstract description 158
- 238000000034 method Methods 0.000 claims abstract description 76
- 230000003068 static effect Effects 0.000 claims abstract description 40
- 238000012360 testing method Methods 0.000 claims abstract description 24
- 239000007789 gas Substances 0.000 claims description 36
- IJGRMHOSHXDMSA-UHFFFAOYSA-N Atomic nitrogen Chemical compound N#N IJGRMHOSHXDMSA-UHFFFAOYSA-N 0.000 claims description 18
- 239000000446 fuel Substances 0.000 claims description 15
- 239000007921 spray Substances 0.000 claims description 15
- CURLTUGMZLYLDI-UHFFFAOYSA-N Carbon dioxide Chemical compound O=C=O CURLTUGMZLYLDI-UHFFFAOYSA-N 0.000 claims description 12
- 238000004364 calculation method Methods 0.000 claims description 10
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 claims description 9
- 239000002737 fuel gas Substances 0.000 claims description 9
- 229910052757 nitrogen Inorganic materials 0.000 claims description 9
- 239000001301 oxygen Substances 0.000 claims description 9
- 229910052760 oxygen Inorganic materials 0.000 claims description 9
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Chemical compound O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 claims description 9
- 229910002092 carbon dioxide Inorganic materials 0.000 claims description 6
- 239000001569 carbon dioxide Substances 0.000 claims description 6
- 230000001133 acceleration Effects 0.000 claims description 3
- UBAZGMLMVVQSCD-UHFFFAOYSA-N carbon dioxide;molecular oxygen Chemical compound O=O.O=C=O UBAZGMLMVVQSCD-UHFFFAOYSA-N 0.000 claims description 3
- 239000000295 fuel oil Substances 0.000 claims description 3
- 239000003350 kerosene Substances 0.000 claims description 3
- 238000004088 simulation Methods 0.000 abstract description 8
- 238000010998 test method Methods 0.000 abstract description 6
- 238000002347 injection Methods 0.000 description 8
- 239000007924 injection Substances 0.000 description 8
- 238000009792 diffusion process Methods 0.000 description 7
- 238000004458 analytical method Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 2
- 238000009529 body temperature measurement Methods 0.000 description 2
- 230000003247 decreasing effect Effects 0.000 description 2
- 230000000670 limiting effect Effects 0.000 description 2
- 238000013178 mathematical model Methods 0.000 description 2
- 239000000203 mixture Substances 0.000 description 2
- 239000000243 solution Substances 0.000 description 2
- 238000004891 communication Methods 0.000 description 1
- 238000007796 conventional method Methods 0.000 description 1
- 230000001419 dependent effect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000004069 differentiation Effects 0.000 description 1
- 238000012854 evaluation process Methods 0.000 description 1
- 230000010354 integration Effects 0.000 description 1
- 230000002452 interceptive effect Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000000737 periodic effect Effects 0.000 description 1
- 230000002829 reductive effect Effects 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
Classifications
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/20—Design optimisation, verification or simulation
- G06F30/28—Design optimisation, verification or simulation using fluid dynamics, e.g. using Navier-Stokes equations or computational fluid dynamics [CFD]
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M15/00—Testing of engines
- G01M15/04—Testing internal-combustion engines
- G01M15/05—Testing internal-combustion engines by combined monitoring of two or more different engine parameters
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2113/00—Details relating to the application field
- G06F2113/08—Fluids
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
- G06F2119/08—Thermal analysis or thermal optimisation
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
- G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Landscapes
- Physics & Mathematics (AREA)
- Engineering & Computer Science (AREA)
- General Physics & Mathematics (AREA)
- Theoretical Computer Science (AREA)
- Fluid Mechanics (AREA)
- Mathematical Physics (AREA)
- Mathematical Analysis (AREA)
- Mathematical Optimization (AREA)
- Computing Systems (AREA)
- Pure & Applied Mathematics (AREA)
- Computer Hardware Design (AREA)
- Evolutionary Computation (AREA)
- Geometry (AREA)
- General Engineering & Computer Science (AREA)
- Algebra (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Combined Controls Of Internal Combustion Engines (AREA)
Abstract
The invention provides a method and a system for evaluating thrust of a ramjet rotary detonation engine, which relate to the technical field of rotary detonation engines, wherein a plurality of temperature test values distributed along the circumferential direction of a combustion chamber at intervals are obtained in a critical state of an exhaust nozzle, the static temperature of the combustion chamber outlet is calculated, the Mach number of the combustion chamber outlet is calculated according to the outlet area of the combustion chamber and the throat area of the nozzle, the total temperature of the combustion chamber outlet is calculated according to the total temperature and static temperature relation, the combustion efficiency of the combustion chamber is calculated according to a temperature rise method, the physical parameters of gas in the combustion chamber are solved iteratively, the pneumatic thermal parameters of an air inlet of the engine, the total pressure of the combustion chamber outlet and the pneumatic thermal parameters of an exhaust nozzle outlet are solved, the thrust of the engine is solved according to the pneumatic thermal parameters of the air inlet of the engine and the exhaust nozzle outlet, and compared with a numerical simulation evaluation method, the whole thrust is more accurate, and compared with a free jet test method, the thrust evaluation efficiency is higher.
Description
Technical Field
The invention relates to the technical field of rotary detonation engines, in particular to a method and a system for evaluating thrust of a ramjet rotary detonation engine.
Background
The conventional method for evaluating the whole thrust of the ramjet rotary detonation engine is generally divided into a numerical simulation method and a free jet wind tunnel test method. The whole thrust is estimated by adopting a numerical simulation method, and the accuracy of the numerical simulation method is lower because the calculation method is dependent, and particularly the rotary detonation wave is difficult to accurately simulate. The thrust performance of the whole engine can be accurately estimated by adopting a free jet test method, but the free jet test period is long, the test process is complex, the test cost is high, and the engine cannot be designed in a rapid iteration mode. Because the numerical simulation method and the free jet wind tunnel test method are mutually independent, the technical guidance for integrating the two thrust evaluation methods is lacking, and the high efficiency and the accuracy of the thrust evaluation are difficult to be considered.
Disclosure of Invention
The invention aims to provide a method and a system for evaluating the thrust of a ramjet engine, which are used for relieving the technical problem that the efficiency and the accuracy are difficult to be considered in evaluating the whole thrust of the ramjet engine.
In a first aspect, the present invention provides a method for evaluating thrust of a ramjet rotary knock engine, comprising the steps of: under the critical state of the exhaust nozzle, acquiring a plurality of temperature test values distributed at intervals along the circumferential direction of the combustion chamber within preset time, and calculating the static temperature T of the outlet of the combustion chamber s4 The method comprises the steps of carrying out a first treatment on the surface of the Acquisition of combustor exit area A 4 And nozzle throat area A cr According to the outlet area A of the combustion chamber 4 And the throat area A of the spray pipe cr Calculating combustor exit Mach number M a4 The method comprises the steps of carrying out a first treatment on the surface of the According to the formulaCalculating the total temperature T of the outlet of the combustion chamber t4 The method comprises the steps of carrying out a first treatment on the surface of the Calculating the combustion efficiency of the combustion chamber according to the temperature rise method>Iteratively solving physical parameters of gas in the combustion chamber; solving aerodynamic thermal parameters of an engine air inlet; calculating the total pressure p of the outlet of the combustion chamber t4 The method comprises the steps of carrying out a first treatment on the surface of the Solving aerodynamic thermal parameters of an outlet of the exhaust nozzle; solving the thrust of the engine according to the aerodynamic thermal parameters of the air inlet and the outlet of the exhaust nozzle of the engineF。
With reference to the first aspect, the present invention provides a first possible implementation manner of the first aspect, wherein the calculating the combustor outlet static temperature T s4 The method comprises the following steps: according to the formulaCalculating the static temperature T of the outlet of the combustion chamber s4 Wherein u is the number of temperature sensors distributed at intervals along the circumferential direction of the combustion chamber, +.>For acquisition time->Is->Test values of the ith temperature sensor over time.
With reference to the first aspect, the present invention provides a second possible implementation manner of the first aspect, wherein the obtaining the outlet area a of the combustion chamber 4 And nozzle throat areaAccording to the outlet area A of the combustion chamber 4 And the throat area of the spray pipeCalculating combustor exit Mach number M a4 The method comprises the following steps: according to the formula->Calculating the outlet area A of the combustion chamber 4 Wherein r is 4i For the diameter of the inner ring of the outlet of the combustion chamber, r 4e The diameter of the outer ring of the outlet of the combustion chamber; according to the formula->Calculating the throat area of the spray pipe>Wherein r is cri Is the diameter of the inner ring of the throat of the spray pipe，r cre The diameter of the outer ring of the throat of the spray pipe is; according to the formula->Solving for combustor exit Mach number M a4 Wherein->Is the initial value of the specific heat ratio of the fuel gas at the outlet of the combustion chamber.
With reference to the first aspect, the present invention provides a third possible implementation manner of the first aspect, wherein the step of solving the aerodynamic thermal parameter of the engine air intake includes: obtaining fly height H and fly Mach number M a0 Based on the acquired flight altitude H and the flight Mach number M a0 Calculating to obtain the static temperature T of the air inlet of the engine s0 Static pressure p of engine air inlet s0 And engine air intake air flow density ρ 0 The method comprises the steps of carrying out a first treatment on the surface of the According to the formulaCalculating the flying speed v 0 Wherein (1)>Is specific heat of air, R a Is an air gas constant; according to the formula->Calculating the air flow m of the engine a Wherein, the method comprises the steps of, wherein,intake port capture area for engine, +.>The flow coefficient of the engine air inlet channel; according to the formulaCalculating total temperature T of inlet airflow of air inlet channel t0 。
With reference to the third possible implementation manner of the first aspect, the present invention provides a fourth possible implementation manner of the first aspectEmbodiments wherein the calculating obtains an engine intake air flow static temperatureStatic pressure of engine air inlet>And engine air intake flow density->The method comprises the following steps: comparing and judging the range of the flying height H; when 0 is<H<At the time of 11000m of the total length of the product,，/>the method comprises the steps of carrying out a first treatment on the surface of the When 11000m is less than or equal to H<At the time of 24000m, the number of the holes is equal to the number of the holes,，/>the method comprises the steps of carrying out a first treatment on the surface of the According to the formula->Calculating the air flow density rho of the air inlet of the engine 0 Wherein g is gravitational acceleration, +.>Is the atmospheric pressure at sea level, e is the natural logarithm.
With reference to the first aspect, the present invention provides a fifth possible implementation manner of the first aspect, wherein the combustion efficiency of the combustion chamber is calculated according to a temperature rise methodThe method comprises the following steps: according to the formula->Calculating combustion efficiency +.>Wherein->Constant pressure specific heat value for combustion chamber outlet gas, < >>Is the low heat value of fuel oil, ">For the total temperature of the combustion chamber inlet->，/>Constant pressure specific heat value for inlet air, +.>Is the specific heat value of fuel.
With reference to the first aspect, the present invention provides a sixth possible implementation manner of the first aspect, wherein the calculating the total pressure of the outlet of the combustion chamberThe method comprises the following steps: according to the formula->Calculating the fuel flow involved in combustion +.>Wherein->To fuel flow into the engine;
according to the formulaCalculating oxygen flow involved in combustion +.>Wherein L is 0 Theoretical air required for complete combustion of 1 kg of fuelAn amount of; according to the formula->Calculating the flow of carbon dioxide in the combustion productsWherein->Molecular mass of carbon dioxide, < >>The fuel molecular mass; according to the formulaCalculating the water vapor flow in the combustion products, wherein +.>Is the molecular mass of water vapor; according to the formula->Calculating the nitrogen flow which does not participate in the combustion air>Wherein->Is the mass fraction of nitrogen; according to the formula->Calculating the oxygen flow rate of the non-participated combustion air>Wherein->Is the mass fraction of oxygen; according to the formula->Calculating the mass ratio of each component gas in the combustion chamberY j Wherein the j component comprises nitrogen, oxygen, carbon dioxide, water vapor and kerosene vapor, n=5; according to the formulaCalculating the average gas constant R of gas mass in the combustion chamber ave Wherein R is j The gas constant of the j component; according to the formula->Calculating constant pressure specific heat C of combustion chamber outlet fuel gas mass average p,ave Wherein C Pj Constant pressure specific heat for the j component at the outlet of the combustion chamber, +.>A, b, c, d, e and f are polynomial calculation parameters, respectively; according to the formula->Calculating specific heat C v,ave The method comprises the steps of carrying out a first treatment on the surface of the According to the formulaCalculating the specific heat ratio value +.>The method comprises the steps of carrying out a first treatment on the surface of the Comparison of specific heat ratio value->Initial value of specific heat ratio of fuel gas with outlet of combustion chamberIf the difference of (2) is smaller than or equal to the preset difference epsilon, if not, calculating the specific heat ratio value +.>Assigning an initial value of the gas specific heat ratio at the outlet of the combustion chamber +.>Iteratively calculating the specific heat ratio calculation value +.>Until the specific heat ratio is calculated as +.>Initial value of gas specific heat ratio with the outlet of the combustion chamber +.>The difference value of (2) is smaller than or equal to the preset difference value epsilon, and the final total temperature T of the outlet of the combustion chamber is obtained t4 The method comprises the steps of carrying out a first treatment on the surface of the According to the formula->Calculating the total pressure of the outlet of the combustion chamber>Wherein->，，/>Is the velocity coefficient of the exhaust nozzle throat, and is +.>。
With reference to the first aspect, the present invention provides a seventh possible implementation manner of the first aspect, wherein the step of solving the aerodynamic thermal parameter of the exhaust nozzle outlet includes: according to the formulaCalculating the Mach number of the outlet of the exhaust nozzle>The method comprises the steps of carrying out a first treatment on the surface of the According to the formulaCalculating the static pressure +.>，/>For the total pressure of the nozzle outlet, assuming an absolute isentropic flow +.>The method comprises the steps of carrying out a first treatment on the surface of the According to the formula->Calculating the static temperature of the outlet of the combustion chamber>Wherein->The total temperature of the nozzle is; according to the formula->Calculating the outlet speed of the exhaust nozzle>。
With reference to the first aspect, the present invention provides an eighth possible implementation manner of the first aspect, wherein the step of solving the engine thrust force F according to aerodynamic thermal parameters of an engine air inlet and an exhaust nozzle outlet includes: according to the formulaCalculating an engine thrust force F, wherein +.>For the impulse of the nozzle>；/>For engine intake thrust->。
In a second aspect, the present invention provides a ramjet rotary knock engine thrust estimation system comprising: a processor for executing a program corresponding to the ramjet rotary knock engine thrust estimation method according to the first aspect; the temperature sensors are respectively arranged on flame barrels of the stamping rotary detonation engine and are arranged at intervals along the circumferential direction of the combustion chamber; and the temperature sensors are respectively connected with the processor.
The embodiment of the invention has the following beneficial effects: under the critical state of the exhaust nozzle, a plurality of temperature test values which are distributed at intervals along the circumferential direction of the combustion chamber in a preset time are obtained, the static temperature of the combustion chamber outlet is calculated, the Mach number of the combustion chamber outlet is calculated according to the area of the combustion chamber outlet and the area of the throat of the nozzle, the total temperature of the combustion chamber outlet is calculated according to the total temperature and the static temperature relation, the combustion efficiency of the combustion chamber is calculated according to a temperature rise method, the physical parameters of gas in the combustion chamber are solved iteratively, the pneumatic thermal parameters of an air inlet of an engine, the total pressure of the combustion chamber outlet and the pneumatic thermal parameters of an outlet of the exhaust nozzle are solved, the thrust of the engine is solved according to the pneumatic thermal parameters of the air inlet of the engine and the outlet of the exhaust nozzle, and compared with the numerical simulation evaluation of the thrust of the whole machine is more accurate, compared with the free jet test method, the thrust evaluation efficiency is more simplified.
In order to make the above objects, features and advantages of the present invention more comprehensible, preferred embodiments accompanied with figures are described in detail below.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the related art, the drawings that are required to be used in the description of the embodiments or the related art will be briefly described, and it is apparent that the drawings in the description below are some embodiments of the present invention, and other drawings may be obtained according to the drawings without inventive effort for those skilled in the art.
FIG. 1 is a flow chart of a method for evaluating thrust of a ramjet engine according to an embodiment of the present invention;
FIG. 2 is a schematic diagram of a ramjet rotary detonation engine;
FIG. 3 is a cross-sectional view of a ramjet rotary detonation engine.
Icon: 1-an injection section; 2-diffusion sections; 3-a flame tube; 4-an exhaust nozzle; 5-nozzle; 6-a thermal jet detonation device; 7-temperature sensor.
Detailed Description
The following description of the embodiments of the present invention will be made apparent and fully in view of the accompanying drawings, in which some, but not all embodiments of the invention are shown. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
In the description of the present invention, it should be noted that the directions or positional relationships indicated by the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc. are based on the directions or positional relationships shown in the drawings, are merely for convenience of describing the present invention and simplifying the description, and do not indicate or imply that the devices or elements referred to must have a specific orientation, be configured and operated in a specific orientation, and thus should not be construed as limiting the present invention. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance. Physical quantities in the formulas, unless otherwise noted, are understood to be basic quantities of basic units of the international system of units, or derived quantities derived from the basic quantities by mathematical operations such as multiplication, division, differentiation, or integration.
In the description of the present invention, it should be noted that, unless explicitly specified and limited otherwise, the terms "mounted," "connected," and "connected" are to be construed broadly, and may be either fixedly connected, detachably connected, or integrally connected, for example; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communication between two elements. The specific meaning of the above terms in the present invention will be understood in specific cases by those of ordinary skill in the art.
As shown in fig. 1, the method for evaluating the thrust of the ramjet rotary knock engine provided by the embodiment of the invention comprises the following steps:
under the critical state of the exhaust nozzle, acquiring a plurality of temperature test values distributed at intervals along the circumferential direction of the combustion chamber within preset time, and calculating the static temperature T of the outlet of the combustion chamber s4 ；
Acquisition of combustor exit area A 4 And nozzle throat area A cr According to the outlet area A of the combustion chamber 4 And nozzle throat area A cr Calculating combustor exit Mach number M a4 ；
According to the formulaCalculating the total temperature T of the outlet of the combustion chamber t4 ；
Calculating the combustion efficiency of the combustion chamber according to the temperature rise methodIteratively solving physical parameters of gas in the combustion chamber;
solving aerodynamic thermal parameters of an engine air inlet;
calculating the total pressure p of the outlet of the combustion chamber t4 ；
Solving aerodynamic thermal parameters of an outlet of the exhaust nozzle;
and solving the engine thrust F according to aerodynamic thermal parameters of an engine air inlet and an exhaust nozzle outlet.
The thrust evaluation method of the ramjet rotary detonation engine, which is recorded in the embodiment, integrates the combustion chamber model test data and the aerodynamic thermodynamic mathematical model analysis method, can evaluate the whole thrust of the ramjet rotary detonation engine through a part-level test, and has the advantages of accurate evaluation method, simple evaluation process and high efficiency. Compared with a numerical simulation evaluation complete machine thrust, the method is more accurate, and compared with a free jet test method, the method is more simplified and has higher thrust evaluation efficiency, so that the engine thrust evaluation has both high efficiency and accuracy.
In the present inventionIn the embodiment, the combustor outlet static temperature T is calculated s4 The method comprises the following steps: according to the formulaCalculating the static temperature T of the outlet of the combustion chamber s4 Wherein u is the number of temperature sensors distributed at intervals along the circumferential direction of the combustion chamber, +.>For acquisition time->Is->Test values of the ith temperature sensor over time. The accuracy of the circumferential average static temperature of the outlet of the combustion chamber can be improved by increasing the number of the temperature sensors, the u can take values of 4, 5 and 6 or more, the temperature sensors are uniformly distributed at intervals along the circumferential direction of the combustion chamber, the static pressure average value is solved in time and space, and the problems of uneven static temperature distribution of the outlet of the combustion chamber and time unsteadiness of static temperature measurement are solved.
Further, the outlet area A of the combustion chamber is obtained 4 And nozzle throat areaAccording to the outlet area A of the combustion chamber 4 And nozzle throat area->Calculating combustor exit Mach number M a4 The method comprises the following steps:
according to the formulaCalculating the outlet area A of the combustion chamber 4 Wherein r is 4i For the diameter of the inner ring of the outlet of the combustion chamber, r 4e The diameter of the outer ring of the outlet of the combustion chamber;
according to the formulaCalculating nozzle throatArea->Wherein r is cri Is the diameter of the inner ring of the throat of the spray pipe, r cre The diameter of the outer ring of the throat of the spray pipe is;
according to the formulaSolving for combustor exit Mach number M a4 Wherein->For the initial value of the gas specific heat ratio of the outlet of the combustion chamber, the initial value of the gas specific heat ratio of the outlet of the combustion chamber is +.>Can be preset or input in advance, r 4i 、r 4e 、r cri And r cre The knock engine can be rotated according to the pressure or detected in the test.
Further, the step of solving aerodynamic thermal parameters of the engine air intake includes:
obtaining fly height H and fly Mach number M a0 Based on the acquired flight altitude H and flight Mach number M a0 Calculating to obtain the static temperature T of the air inlet of the engine s0 Static pressure p of engine air inlet s0 And engine air intake air flow density ρ 0 ；
According to the formulaCalculating the flying speed v 0 Wherein (1)>Is specific heat of air, R a Is an air gas constant;
according to the formulaCalculating the air flow m of the engine a Wherein->Intake port capture area for engine, +.>The flow coefficient of the engine air inlet channel;
according to the formulaCalculating total temperature T of inlet airflow of air inlet channel t0 。
Further, the static temperature of the air inlet of the engine is obtained through calculationStatic pressure of engine air inlet>And engine air intake flow density->The method comprises the following steps:
comparing and judging the range of the flying height H;
when 0 is<H<At the time of 11000m of the total length of the product,，/>；
when 11000m is less than or equal to H<At the time of 24000m, the number of the holes is equal to the number of the holes,，/>；
according to the formulaCalculating the air flow density rho of the air inlet of the engine 0 Wherein g is gravitational acceleration, +.>Is the atmospheric pressure at sea level, e is the natural logarithm.
Further, the combustion efficiency of the combustion chamber is calculated according to a temperature rise methodThe method comprises the following steps:
according to the formulaCalculating combustion efficiency +.>Wherein->Constant pressure specific heat value for combustion chamber outlet gas, < >>Is the low heat value of fuel oil, ">For the total temperature of the combustion chamber inlet->，/>Constant pressure specific heat value for inlet air, +.>Is the specific heat value of fuel.
Further, calculating the total pressure of the outlet of the combustion chamberThe method comprises the following steps:
according to the formulaCalculating the fuel flow involved in combustion +.>Wherein->To fuel flow into the engine;
according to the formulaCalculating oxygen flow involved in combustion +.>Wherein L is 0 The theoretical amount of air required to completely burn 1 kg of fuel;
according to the formulaCalculating the flow of carbon dioxide in the combustion products +.>Wherein->Molecular mass of carbon dioxide, < >>The fuel molecular mass;
according to the formulaCalculating the water vapor flow in the combustion products, wherein +.>Is the molecular mass of water vapor;
according to the formulaCalculating the nitrogen flow which does not participate in the combustion air>Wherein, the method comprises the steps of, wherein,is the mass fraction of nitrogen;
according to the formulaCalculating the oxygen flow rate of the non-participated combustion air>Wherein, the method comprises the steps of, wherein,is the mass fraction of oxygen;
according to the formulaCalculating the mass ratio Y of each component gas in the combustion chamber j Wherein the j component comprises nitrogen, oxygen, carbon dioxide, water vapor and kerosene vapor, n=5;
according to the formulaCalculating the average gas constant R of gas mass in the combustion chamber ave Wherein R is j The gas constant of the j component;
according to the formulaCalculating constant pressure specific heat C of combustion chamber outlet fuel gas mass average p,ave Wherein C Pj Constant pressure specific heat for the j component at the outlet of the combustion chamber, +.>A, b, c, d, e and f are polynomial calculation parameters, respectively;
according to the formulaCalculating specific heat C v,ave ；
According to the formulaCalculating the specific heat ratio value +.>；
Comparing the calculated specific heat ratioInitial value of specific heat ratio of fuel gas with outlet of combustion chamber +.>If the difference of (2) is less than or equal to the preset difference epsilon, if not, calculating the specific heat ratio value +.>Assigned to the initial value of the specific heat ratio of the combustion chamber outlet>Iterative calculation of the specific heat ratio value +.>Until the specific heat ratio is calculated>Initial value of specific heat ratio of fuel gas with outlet of combustion chamber +.>The difference value of (2) is smaller than or equal to the preset difference value epsilon, and the final total temperature T of the outlet of the combustion chamber is obtained t4 ；
According to the formulaCalculating the total pressure of the outlet of the combustion chamber>Wherein->，，/>Is the velocity coefficient at the throat of the exhaust nozzle, at the exhaustIn the critical or supercritical state of the throat of the air jet pipe>。
Further, the step of solving aerodynamic thermal parameters of the exhaust nozzle outlet includes:
according to the formulaCalculating the Mach number of the outlet of the exhaust nozzle>；
According to the formulaCalculating the static pressure +.>，/>For the total pressure of the nozzle outlet, assuming an absolute isentropic flow +.>；
According to the formulaCalculating the static temperature of the outlet of the combustion chamber>Wherein->The total temperature of the nozzle is;
according to the formulaCalculating the outlet speed of the exhaust nozzle>。
Further, the step of solving the engine thrust F according to aerodynamic thermal parameters of the engine intake and exhaust nozzle outlets comprises:
according to the formulaCalculating an engine thrust force F, wherein +.>For the impulse of the nozzle>；For engine intake thrust->. From which the calculated engine thrust can be deduced。
In the method for evaluating the thrust of the ramjet engine, the specific heat of the gas mixture in the combustion chamber and the specific heat ratio are greatly changed with respect to the temperature and the components, but in the method for evaluating the thrust of the ramjet engine, the static temperature of the outlet of the combustion chamber and the flow rate of the combustion chamber are used as constraints, and the physical parameters such as the specific heat ratio of the mixed gas and the aerodynamic parameters such as the total pressure of the mixed gas are iteratively solved, so that the thrust of the engine is obtained. On one hand, the combustion chamber model is simplified, the test process is simple, the test period is short, and the test cost is low; on the other hand, the total pressure and the total temperature of the outlet of the combustion chamber are accurately calculated, and the accuracy of thrust evaluation is improved.
As shown in fig. 2 and 3, the thrust evaluation system for a ramjet rotary knock engine according to an embodiment of the present invention includes: a processor and a plurality of temperature sensors 7. The steps of the ramjet engine thrust estimation method may be integrated, programmed according to the flowchart shown in fig. 1, and the program may be embedded in a memory, processor. The rotary detonation engine suitable for the system is provided with an injection section 1, a diffusion section 2, a flame tube 3 and an exhaust spray pipe 4 which are sequentially arranged, wherein the injection section 1, the diffusion section 2, the flame tube 3 and the exhaust spray pipe 4 are all composed of an inner pipe and an outer pipe, the inner pipe and the outer pipe of the injection section 1 are constant in diameter along the axial direction, the inner pipe diameter of the diffusion section 2 is gradually reduced along the airflow direction, the outer pipe diameter of the diffusion section 2 is gradually increased, the outer pipe and the inner pipe of the flame tube 3 are kept constant in diameter along the airflow direction, the inner pipe diameter of the exhaust spray pipe 4 is firstly increased and then decreased, and the outer pipe diameter is firstly decreased and then increased. Inlet air from the axle head of the injection section 1 deviating from the diffusion section 2, and inject fuel into the injection section 1 by a plurality of nozzles 5 arranged on the injection section 1, and the plurality of nozzles 5 are arranged at intervals along the circumferential direction of the injection section 1. The oil-gas mixture enters the flame tube 3, takes an annular chamber inside the flame tube 3 as a combustion chamber, and can form rotary detonation waves in the flame tube 3 through the ignition and detonation of a thermal jet detonation device 6 arranged at the position of the flame tube 3 close to the diffusion section 2. The exhaust gas generated by knocking the flame tube 3 flows into the exhaust nozzle 4 and is ejected out of the tail end of the exhaust nozzle 4. The temperature sensors 7 are respectively installed on the flame tube 3 of the ramjet rotary detonation engine, are arranged at intervals along the circumferential direction of the combustion chamber, and are respectively connected with the processor. An interactive device can be additionally arranged and connected with the processor.
By adopting the method and the system for evaluating the thrust of the ramjet rotary detonation engine, the method and the system have the following beneficial effects:
(1) The method integrates the combustion chamber model test data and the aerodynamic thermodynamic mathematical model analysis method, and the whole thrust of the ramjet rotary detonation engine is estimated through the test of the component level, so that the estimation method is accurate and the estimation process is simple. Compared with a method for evaluating the thrust of the whole machine by numerical simulation, the method provided by the invention is based on the test data of the rotary detonation combustion for evaluation, and has high accuracy. Compared with free jet test, the invention has the advantages of simple combustion chamber model, simple test process, short test period and low test cost.
(2) A plurality of temperature openers are uniformly distributed in the circumferential direction of the combustion chamber so as to collect the static temperature of the outlet of the combustion chamber, and the problem of space non-uniformity of the static temperature of the outlet of the combustion chamber is solved. In the thrust evaluation method, time average is carried out on a plurality of periodic static temperatures acquired by the temperature sensor to obtain the static temperature of the outlet of the combustion chamber, so that the problem of time unsteadiness of static temperature measurement is solved.
(3) And (3) taking the static temperature of the outlet of the combustion chamber and the flow of the combustion chamber as constraints, and iteratively solving physical parameters such as the specific heat ratio of the mixed gas and aerodynamic parameters such as the total pressure of the mixed gas, thereby solving the thrust of the whole engine. The influence of temperature on physical parameters such as specific heat ratio and gas constant is avoided pertinently, the total pressure and total temperature of the outlet of the combustion chamber are calculated more accurately, and the estimation accuracy of the engine thrust is higher.
Finally, it should be noted that: the above embodiments are only for illustrating the technical solution of the present invention, and not for limiting the same; although the invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical scheme described in the foregoing embodiments can be modified or some or all of the technical features thereof can be replaced by equivalents; such modifications and substitutions do not depart from the spirit of the invention.
Claims (7)
1. A method of evaluating thrust of a ramjet rotary knock engine, comprising the steps of:
under the critical state of the exhaust nozzle, acquiring a plurality of temperature test values distributed at intervals along the circumferential direction of the combustion chamber within preset time, and calculating the static temperature T of the outlet of the combustion chamber s4 ；
Acquisition of combustor exit area A 4 And nozzle throat area A cr According to the outlet area A of the combustion chamber 4 And the throat area A of the spray pipe cr Calculating combustor exit Mach number M a4 ；
According to the formulaCalculating the total temperature T of the outlet of the combustion chamber t4 ；
Calculating the combustion efficiency of the combustion chamber according to the temperature rise methodIteratively solving physical parameters of gas in the combustion chamber;
solving aerodynamic thermal parameters of an engine air inlet; the step of solving the aerodynamic thermal parameters of the engine air inlet comprises the following steps: obtaining fly height H and fly Mach number M a0 Based on the acquired flight altitude H and the flight Mach number M a0 Calculating to obtain the static temperature T of the air inlet of the engine s0 Static pressure p of engine air inlet s0 And engine air intake air flow density ρ 0 The method comprises the steps of carrying out a first treatment on the surface of the According to the formulaCalculating the flying speed v 0 Wherein (1)>Is specific heat of air, R a Is an air gas constant; according to the formulaCalculating the air flow m of the engine a Wherein->Intake port capture area for engine, +.>The flow coefficient of the engine air inlet channel; according to the formula->Calculating total temperature T of inlet airflow of air inlet channel t0 ；
Calculating the total pressure p of the outlet of the combustion chamber t4 The method comprises the steps of carrying out a first treatment on the surface of the Said calculating the total pressure p of the outlet of the combustion chamber t4 The method comprises the following steps: according to the formulaCalculating the fuel flow involved in combustion +.>Wherein->To fuel flow into the engine; according to the formulaCalculating oxygen flow involved in combustion +.>Wherein L is 0 The theoretical amount of air required to completely burn 1 kg of fuel; according to the formula->Calculating the flow of carbon dioxide in the combustion products +.>Wherein->Molecular mass of carbon dioxide, < >>The fuel molecular mass; according to the formula->Calculating the water vapor flow in the combustion products, wherein +.>Is the molecular mass of water vapor; according to the formula->Calculating the nitrogen flow which does not participate in the combustion air>Wherein->Is the mass fraction of nitrogen; according to the formulaCalculating the oxygen flow rate of the non-participated combustion air>Wherein->Is the mass fraction of oxygen; according to the formula->Calculating the mass ratio Y of each component gas in the combustion chamber j Wherein the j component comprises nitrogen, oxygen, carbon dioxide, water vapor and kerosene vapor, n=5; according to the formula->Calculating the average gas constant R of gas mass in the combustion chamber ave Wherein R is j The gas constant of the j component; according to the formula->Calculating constant pressure specific heat C of combustion chamber outlet fuel gas mass average p,ave Wherein C Pj The constant pressure specific heat of the j component at the outlet of the combustion chamber,a, b, c, d, e and f are polynomial calculation parameters, respectively; according to the formulaCalculating specific heat C v,ave The method comprises the steps of carrying out a first treatment on the surface of the According to the formula->Calculating the specific heat ratio value +.>The method comprises the steps of carrying out a first treatment on the surface of the Comparison of specific heat ratio value->Initial value of specific heat ratio of fuel gas with outlet of combustion chamber +.>If the difference of (2) is smaller than or equal to the preset difference epsilon, if not, calculating the specific heat ratio value +.>Assigning an initial value of the gas specific heat ratio at the outlet of the combustion chamber +.>Iteratively calculating the specific heat ratio calculation value +.>Until the specific heat ratio is calculated as +.>Initial value of gas specific heat ratio with the outlet of the combustion chamber +.>The difference value of (2) is smaller than or equal to the preset difference value epsilon, and the final total temperature T of the outlet of the combustion chamber is obtained t4 The method comprises the steps of carrying out a first treatment on the surface of the According to the formula->Calculating the total pressure of the outlet of the combustion chamber>Wherein->，/>，/>Is the velocity coefficient of the exhaust nozzle throat, and is +.>；
Solving aerodynamic thermal parameters of an outlet of the exhaust nozzle; the step of solving the aerodynamic thermal parameters of the outlet of the exhaust nozzle comprises the following steps: according to the formulaCalculating the Mach number of the outlet of the exhaust nozzle>The method comprises the steps of carrying out a first treatment on the surface of the According to the formula->Calculating the static pressure +.>，/>For the total pressure of the nozzle outlet, assuming an absolute isentropic flow +.>The method comprises the steps of carrying out a first treatment on the surface of the According to the formula->Calculating the static temperature of the outlet of the combustion chamber>Wherein->The total temperature of the nozzle is; according to the formula->Calculating the outlet speed of the exhaust nozzle>；
And solving the engine thrust F according to aerodynamic thermal parameters of an engine air inlet and an exhaust nozzle outlet.
2. The ramjet rotary knock engine thrust estimation method according to claim 1, wherein the calculation of the combustion chamber outlet static temperature T s4 The method comprises the following steps:
according to the formulaCalculating the static temperature T of the outlet of the combustion chamber s4 Wherein u is the number of temperature sensors distributed at intervals along the circumferential direction of the combustion chamber, +.>For acquisition time->Is->Test values of the ith temperature sensor over time.
3. The ramjet rotary knock engine thrust estimation method according to claim 1, wherein the acquisition of the combustion chamber outlet area a 4 And nozzle throat areaAccording to the outlet area A of the combustion chamber 4 And the throat area of the spray pipe>Calculating combustor exit Mach number M a4 The method comprises the following steps:
according to the formulaCalculating the outlet area A of the combustion chamber 4 Wherein r is 4i For the diameter of the inner ring of the outlet of the combustion chamber, r 4e The diameter of the outer ring of the outlet of the combustion chamber;
according to the formulaCalculating the throat area of the spray pipe>Wherein r is cri Is the diameter of the inner ring of the throat of the spray pipe, r cre The diameter of the outer ring of the throat of the spray pipe is;
according to the formulaSolving for combustor exit Mach number M a4 Wherein->Is the initial value of the specific heat ratio of the fuel gas at the outlet of the combustion chamber.
4. The ramjet rotary knock engine thrust estimation method according to claim 1, wherein the calculation obtains an engine intake air flow static temperature T s0 Static pressure p of engine air inlet s0 And engine air intake air flow density ρ 0 The method comprises the following steps:
comparing and judging the range of the flying height H;
when 0 is<H<At the time of 11000m of the total length of the product,，/>；
when 11000m is less than or equal to H<At the time of 24000m, the number of the holes is equal to the number of the holes,，/>；
according to the formulaCalculating the air flow density rho of the air inlet of the engine 0 Wherein g is gravitational acceleration, +.>Is the atmospheric pressure at sea level, e is the natural logarithm.
5. The ramjet rotary knock engine thrust force evaluation method according to claim 1, wherein the combustion efficiency of the combustion chamber is calculated by a temperature increase methodThe method comprises the following steps:
according to the formulaCalculating combustion efficiency +.>Wherein->Constant pressure specific heat value for combustion chamber outlet gas, < >>Is the low heat value of fuel oil, ">For the total temperature of the combustion chamber inlet->，/>Constant pressure specific heat value for inlet air, +.>Is the specific heat value of fuel.
6. The ramjet rotary detonation engine thrust estimation method of claim 1, wherein the step of solving for the engine thrust F based on aerodynamic thermal parameters of an engine intake and an exhaust nozzle outlet comprises:
according to the formulaCalculating an engine thrust force F, wherein +.>For the impulse of the nozzle>；/>For engine intake thrust->。
7. A ramjet rotary knock engine thrust estimation system, comprising:
a processor for executing a program corresponding to the ramjet rotary knock engine thrust force evaluation method according to any one of claims 1 to 6;
a plurality of temperature sensors (7), wherein the temperature sensors (7) are respectively arranged on flame tubes (3) of the ramjet rotary detonation engine and are arranged at intervals along the circumferential direction of the combustion chamber;
wherein a plurality of temperature sensors (7) are respectively connected with the processor.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310835905.XA CN116562194B (en) | 2023-07-10 | 2023-07-10 | Thrust evaluation method and system for ramjet rotary detonation engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310835905.XA CN116562194B (en) | 2023-07-10 | 2023-07-10 | Thrust evaluation method and system for ramjet rotary detonation engine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN116562194A CN116562194A (en) | 2023-08-08 |
| CN116562194B true CN116562194B (en) | 2023-09-19 |
Family
ID=87488374
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310835905.XA Active CN116562194B (en) | 2023-07-10 | 2023-07-10 | Thrust evaluation method and system for ramjet rotary detonation engine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN116562194B (en) |
Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2011160441A1 (en) * | 2010-06-21 | 2011-12-29 | Jin Beibiao | Thermal ramjet |
| CN106407571A (en) * | 2016-09-22 | 2017-02-15 | 北京机械设备研究所 | A hypersonic velocity air-breathing type ramjet pneumatic thrust analysis method |
| CN109460626A (en) * | 2018-12-06 | 2019-03-12 | 北京空天技术研究所 | Punching engine performance parameter calculation method |
| CN111157248A (en) * | 2020-01-06 | 2020-05-15 | 中国人民解放军国防科技大学 | Ramjet based on ground direct connection test and combustion chamber performance evaluation method thereof |
| CN114722743A (en) * | 2022-05-24 | 2022-07-08 | 中国人民解放军国防科技大学 | Combustion chamber chemical balance-based scramjet engine one-dimensional performance estimation method |
| CN115060504A (en) * | 2022-06-24 | 2022-09-16 | 中国人民解放军国防科技大学 | Method for determining combustion mode and isolation section airflow parameters of ramjet in real time |
| CN115169056A (en) * | 2022-08-11 | 2022-10-11 | 西北工业大学 | Unsteady state performance estimation method for sub-combustion ramjet engine |
| CN115982943A (en) * | 2022-11-25 | 2023-04-18 | 南京航空航天大学 | Volumetric dynamics one-dimensional modeling method for scramjet engine |
-
2023
- 2023-07-10 CN CN202310835905.XA patent/CN116562194B/en active Active
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2011160441A1 (en) * | 2010-06-21 | 2011-12-29 | Jin Beibiao | Thermal ramjet |
| CN106407571A (en) * | 2016-09-22 | 2017-02-15 | 北京机械设备研究所 | A hypersonic velocity air-breathing type ramjet pneumatic thrust analysis method |
| CN109460626A (en) * | 2018-12-06 | 2019-03-12 | 北京空天技术研究所 | Punching engine performance parameter calculation method |
| CN111157248A (en) * | 2020-01-06 | 2020-05-15 | 中国人民解放军国防科技大学 | Ramjet based on ground direct connection test and combustion chamber performance evaluation method thereof |
| CN114722743A (en) * | 2022-05-24 | 2022-07-08 | 中国人民解放军国防科技大学 | Combustion chamber chemical balance-based scramjet engine one-dimensional performance estimation method |
| CN115060504A (en) * | 2022-06-24 | 2022-09-16 | 中国人民解放军国防科技大学 | Method for determining combustion mode and isolation section airflow parameters of ramjet in real time |
| CN115169056A (en) * | 2022-08-11 | 2022-10-11 | 西北工业大学 | Unsteady state performance estimation method for sub-combustion ramjet engine |
| CN115982943A (en) * | 2022-11-25 | 2023-04-18 | 南京航空航天大学 | Volumetric dynamics one-dimensional modeling method for scramjet engine |
Non-Patent Citations (2)
| Title |
|---|
| Sub-nA Low-Current HZO Ferroelectric Tunnel Junction for High-Performance and Accurate Deep Learning Acceleration;Tzu-Yun Wu ET AL;《2019 IEEE International Electron Devices Meeting (IEDM)》;全文 * |
| 吴云 ; 王健 ; 宋慧敏 ; 贾敏 ; 崔巍.电磁复合等离子体气动激励控制激波的原理研究.《中国人民解放军空军工程大学》.2013,全文. * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN116562194A (en) | 2023-08-08 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| EP3584557B1 (en) | System and method for estimating an air mass flow of air flowing in a bypass duct of a gas turbine engine | |
| CN110362783A (en) | Rocket-based combined cycle engine thrust calculation method | |
| CN111157248B (en) | Ramjet based on ground direct connection test and combustion chamber performance evaluation method thereof | |
| Lang et al. | Statistical evaluation of ignition phenomena in turbojet engines | |
| CN109141907A (en) | A kind of supersonic speed engine testsand | |
| Fievisohn et al. | Experimental measurements of equivalent available pressure-lessons learned | |
| CN116562194B (en) | Thrust evaluation method and system for ramjet rotary detonation engine | |
| CN111173644B (en) | Method for determining performance of ramjet engine | |
| CN116127815B (en) | Modeling method of turbofan engine with injection nozzle | |
| CN109282989A (en) | A kind of supersonic speed engine testsand gas handling system | |
| Choi et al. | Infrared signal of the lobed mixer with external air mixing | |
| CN116593168B (en) | Method and system for evaluating fuel consumption rate of ramjet rotary detonation engine | |
| CN111090936B (en) | Multi-stage ignition performance matching simulation calculation method for gas generator | |
| CN115236265A (en) | Method and device for estimating internal ballistic performance of solid rocket engine | |
| CN115221638A (en) | Performance time response analysis method for unsteady state process of sub-combustion ramjet engine | |
| CN116562193B (en) | Combustion efficiency analysis method and system for rotary detonation engine | |
| CN114996978B (en) | Method and device for calculating airflow parameters at inlet of combustion chamber of turbofan engine | |
| Gallis et al. | Development and Validation of a 0-D/1-D Model to Evaluate Pulsating Conditions from a Constant Volume Combustor | |
| CN115062866A (en) | Discrete spectrum contribution-based scramjet engine flame infrared spectrum radiation prediction method | |
| Galigher et al. | Evaluation of boattail geometry and exhaust plume temperature effects on nozzle afterbody drag at transonic Mach numbers | |
| Cadiou | Experimental method for the combustion efficiency calculation in a reheat duct | |
| CN117993319A (en) | Self-adaptive airborne real-time modeling method for scramjet engine by perturbation optimization method | |
| Olson et al. | Method for estimating combustion efficiency at altitude flight conditions from combustor tests at low pressures | |
| CN109141909A (en) | A kind of application method of supersonic speed engine testsand | |
| Le et al. | Preliminary investigation of full model of two-mode scramjet |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |