Study of heat transfer processes in the flowing part of hypersonic air-ramjet engine

The technique and results of the experimental-theoretical study of gas dynamics, heat transfer and the structure of gas flow in the flowing channel of a model hypersonic air-ramjet engine are presented for Mach numbers M = (5; 6).


Introduction
At hypersonic flight speeds of rockets at low altitudes, it is supposed to use solid fuel airramjet engine.Wherein, the air from the atmosphere, which enters through the air intake into the flowing channel of the engine with velocities M > 5 [1], is used as an oxidizer in the organization of combustion.The processes of combustion and burning-out (possibly, uneven) of the surface of the solid fuel element are determined by gas-dynamic and thermal parameters of flow in the flowing part of the engine.Thus, for design of engine facility, it is necessary to have of experimental studies data from that allow one to predict these parameters with a high degree of veracity for concrete designs [2].It is especially important to solve the problem of assessing the thermal state of the flowing channel in the initial period after the switching-on of an installation.The presence of predictive estimates of the thermal state of the tract at the design stage of a specific design is necessary for a wellfounded choice of the ignition method and the provision sustainable combustion of solid fuel.It is advisable to conduct preliminary studies with laboratory models, since full-scale stand experiments are costly.
The purpose of the study was to obtain objective information about the thermal state of the walls of the flowing channel of a model hypersonic air-ramjet engine at Mach numbers M = (5,6).

Technique of the experimental study
An experimental study was carried out on a pulsed aerodynamic installation with heating of gas (up to Te ≈ 700 K) [3][4][5].The aerodynamic installation provides the speed of the inflowing stream in the range of Mach numbers M = (2 ÷ 7).Steel axisymmetric profiled https://doi.org/10.1051/matecconf/201819401037HMTTSC-2018 Matec Web of Conferences nozzles [4] were used to create a hypersonic flow for Mach numbers M = (5,6).The experiments were conducted on an axisymmetric model of an air-ramjet engine (Fig. 1, a).The measurement of the gas temperature in the flowing channel was carried out by a thermocouple method [6].A thermo-probe was manufactured for measure of the temperature along the wall of the model flowing channel.The thermos-probe represents a set of fluoroplastic rings of the equal diameter.Eight copper rings with a chromel-copel thermocouple were attached to the inner surface of each ring.The view of the axisymmetric model with the installed thermo-probe is shown in Fig. 1, b.The scheme (Fig. 2, a) and photographs of the thermo-probe in assembled form (Fig. 2, b) and the temperature sensor (Fig. 2, c The correlation of oscillatory regimes by Mach number, and temperature on the symmetry axis is noted.The Mach number decreases in the zones of compression shock, and the temperature increases.The Mach number increases in the decompression zones, and the temperature decreases.

Conclusion
The thermo-probe is developed, designed and implemented.The thermo-probe is equipped by temperature sensors with chromel-copel thermocouples.The working range of the measuring temperatures is T = (200 ÷ 600) K.
The experimentally-calculating technique of studying gas dynamics, heat transfer and the structure of gas flow in the flowing channel of hypersonic air-ramjet engine are proposed.
Objective data about the heat transfer characteristics in the flowing part of the axisymmetric model of the air-ramjet engine were obtained for Mach numbers M = (5,6) on the basis of the developed technique.This research was supported by grant (No. 8.2.09.2018) from "The Tomsk State University competitiveness improvement program".

Fig. 1 .
Fig. 1.General view of model (a), and model with the installed thermo-probe (b).

Fig. 2 .Fig. 3 .Fig. 4 .Fig. 5 .Fig. 6 .Fig. 7 .Fig. 8 .
Fig. 2. Scheme of the thermo-probe (a), the thermo-probe fully assembled (b) and the temperature sensor (c).The temperature in the prechamber and in the exit section of the model was measured with the help of special temperature sensors during process of the experiments.The view of placed in the working part of the aerodynamic installation the axisymmetric model in complex with temperature sensors is shown in Fig.3.