RU-2861612-C1 - METHOD FOR DETERMINING HEAT LOSS COEFFICIENT IN COMBUSTION CHAMBER

RU2861612C1RU 2861612 C1RU2861612 C1RU 2861612C1RU-2861612-C1

Abstract

FIELD: engine building. SUBSTANCE: invention relates to the field of research of heat and mass transfer processes in solid-propellant, ramjet and hybrid rocket engines and can be used to determine the heat loss coefficient in the engine combustion chamber. The method includes measuring the dependence of pressure in the chamber on time p(t) during adiabatic outflow through the nozzle of solid-propellant charge combustion products. The dependence p(t) is preliminarily measured in a combustion chamber with a coating that thermally insulates the inner walls during free outflow after forced charge extinguishing by pressure relief, and the nozzle discharge coefficient is determined. Then the dependence p(t) is measured in the combustion chamber without a heat-insulating coating during free outflow of combustion products through an identical nozzle after complete combustion of the charge, and the heat loss coefficient is determined. To calculate, from the obtained experimental dependences p(t), the nozzle discharge coefficient, the heat loss coefficient and to select the relaxation time ensuring adiabatic outflow, specified algebraic relationships are used. EFFECT: enabling determination of the heat loss coefficient in combustion chambers of solid-propellant power plants in an extended temperature range without the use of thermocouples. 1 cl, 4 dwg, 2 tbl

Inventors

Arkhipov Vladimir Afanasevich
Basalaev Sergei Aleksandrovich
Perfileva Kseniia Grigorevna
Poriazov Vasilii Andreevich

Dates

Publication Date: 20260506
Application Date: 20251118

Claims (20)

A method for determining the heat loss coefficient in a combustion chamber, including measuring the dependence of the pressure in the chamber on time p ( t ) during the adiabatic outflow of combustion products of a solid propellant charge through a nozzle, characterized in that the dependence p ( t ) is first measured in a combustion chamber with a heat-insulating coating on the inner walls during the free outflow after forced quenching of the charge by releasing the pressure and the nozzle flow coefficient is determined, then the dependence p ( t ) is measured in a combustion chamber without a heat-insulating coating during the free outflow of combustion products through an identical nozzle after complete combustion of the charge and the heat loss coefficient is determined, wherein the flow coefficient and the heat loss coefficient are determined in accordance with the relationships
, ,
, ,
and the relaxation time of the free volume of the combustion chamber is selected in accordance with the inequality
,
where ϕ is the nozzle flow coefficient;
n – number of experimental points And , evenly distributed over the pressure relief area;
k – adiabatic index of combustion products;
τ = t / t k – dimensionless time;
t – time;
– relaxation time of the free volume of the combustion chamber;
V – free volume of the combustion chamber;
S cr – critical cross-sectional area of the nozzle;
– function of the adiabatic index of combustion products;
R – gas constant of combustion products;
T 0 – temperature of combustion products before pressure release;
x ( τ )= p ( t )/ p 0 – dimensionless pressure in the combustion chamber with heat-insulating coating;
p ( t ) – dependence of pressure in the combustion chamber on time during the period of free flow of combustion products;
p 0 – pressure in the combustion chamber before pressure relief;
– heat loss coefficient;

Description

The invention relates to the field of studying heat and mass transfer processes in solid-fuel, ramjet and hybrid rocket engines and can be used to determine the heat loss coefficient in the engine combustion chamber. The heat loss coefficient χ in the combustion chamber is one of the main characteristics of a solid propellant rocket engine (SPRE) and is determined by the ratio [1] , (1) Where- area of the inner surface of the combustion chamber walls, m2 ; α - heat transfer coefficient, W/( m2 K); T0 - gas temperature in the combustion chamber, K; - temperature of the inner surface of the combustion chamber walls in contact with gases, K; - heat of combustion of solid fuel, J/kg; - mass gas input per second during charge combustion, kg/s. The heat loss coefficient takes into account the combustion energy loss of the solid propellant charge in the chamber due to two factors: incomplete combustion and heat loss in the combustion chamber walls [2]. The value of the heat loss coefficient is determined by a number of parameters: the type of charge, the surface area of the chamber in contact with the gases, the presence of thermal insulation coatings, the temperature and composition of the combustion products, the presence of condensed particles in the gas, and the nature of the movement of the gaseous combustion products in the flow path of the chamber. The theoretical determination of the heat loss coefficient using relation (1) is associated with the need to carry out calculations (usually in a three-dimensional formulation) of the gas dynamics of combustion products in the engine flow path to determine the heat transfer coefficient α, and also requires the use of a large number of empirical constants for the thermophysical characteristics of the charge and chamber materials, thermodynamic parameters of combustion products, empirical dependencies for the Nusselt number and does not always provide sufficient accuracy of calculations. A method for experimentally determining the heat loss coefficient in a model solid propellant rocket motor with an insert charge, with combustion interrupted at various points in time, is known [1]. The amount of heat accumulated by the motor body during charge combustion was determined in a calorimeter in which the motor was placed after combustion was interrupted. This method is labor-intensive and can only be used for small-sized solid propellant rocket motors. To determine the heat loss coefficient in the combustion chamber, one can use the method of directly measuring the temperature of combustion products T0 in the solid propellant rocket motor chamber and calculating the value of χ using the formula [2]: , where Tad is the adiabatic combustion temperature of the fuel, obtained from thermodynamic calculations. A known method for measuring the temperature of combustion products in an experimental gas generator [3] involves a conical insert with connections for pressure and temperature sensors soldered between the housing and the nozzle block. The measured temperature of the combustion products was (800 ÷ 2100)°C. A method for measuring combustion product temperature is known, implemented in a setup for studying temperature and mass transfer processes in solid-propellant rocket motor designs [4]. The setup includes a solid-fuel gas generator, a gas duct with a cylindrical channel, and a nozzle. The outer surface of the gas duct, which has one or more faces, is fitted with fittings for temperature sensors. These methods are applicable only for engines with a short operating time and are limited by the range of fuel combustion temperatures, since modern thermoelectric sensors remain operational at temperatures no higher than 1300°C (chromel-alumel), 1760°C (platinum-rhodium) and 2500°C (tungsten-rhenium) [5]. Currently, solid propellant charges are infused with metal powders and active combustible binders, which significantly increase combustion temperatures [6]. Furthermore, condensed combustion products clog the thermocouple junction, increasing measurement error. The closest in technical essence to the claimed invention is a semi-empirical method for determining the parameters of adiabatic gas flow from a receiver through a nozzle [7]. The method is based on measuring the pressure-time dependence p ( t ) during the period of free gas flow upon sudden opening of the nozzle. The measured dependence p ( t ) is used to calculate the time to establish a quasi-steady state of flow and the nozzle discharge coefficient. The technical result of the present invention is the development of a method for determining the heat loss coefficient in a combustion chamber without using thermocouples placed in the combustion chamber and thereby ensuring an expansion of the class of studied compositions of solid fuels with a high combustion temperature. The technical result is achieved by developing a method for determining the heat loss coefficient in a combustion chamber. This method involves mea