Effects of hydrogen–air non–equilibrium chemistry on the performance of a model scramjet thrust nozzle

Abstract

Two aspects of hydrogen-air non-equilibrium chemistry related to scramjets are nozzle freezing and a process called 'kinetic afterburning' which involves continuation of combustion after expansion in the nozzle. These effects were investigated numerically and experimentally with a model scramjet combustion chamber and thrust nozzle combination. The overall model length was 0⋅5m, while precombustion Mach numbers of 3⋅1±0⋅3 and precombustion temperatures ranging from 740K to 1,400K were involved. Nozzle freezing was investigated at precombustion pressures of 190kPa and higher, and it was found that the nozzle thrusts were within 6% of values obtained from finite rate numerical calculations, which were within 7% of equilibrium calculations. When precombustion pressures of 70kPa or less were used, kinetic afterburning was found to be partly responsible for thrust production, in both the numerical calculations and the experiments. Kinetic afterburning offers a means of extending the operating Mach number range of a fixed geometry scramjet.

Scramjets and shock tunnels—The Queensland experience

“…This non-linear growth in wake displacement thickness may partly explain why the experimental results of Fig. 8 indicate that duct lengths of 7 duct heights and 18 duct heights produce approximately the same propulsive effects, and why experiments with precombustion pressures in excess of 190 kPa and a combustion duct 25 mm high yielded nozzle pressure distributions indicating complete combustion in the duct with a duct length of 7 times the duct height (29) . Another possible explanation for the relatively short duct lengths for complete combustion is associated with the relatively low precombustion Mach numbers, of 3.4 ± 0.5, which prevailed in these experiments.…”

“…(Figure 15 here) In experiments at a stagnation enthalpy of 3.5 MJ/kg and a Mach number of 6.4, the thrust with combustion increased with equivalence ratio, and became equal to the drag as the equivalence ratio approached unity, thus achieving the cruise condition of net zero thrust or drag. By comparison with results of pressure measurements in a combustion duct-thrust nozzle combination, it was concluded that combustion was taking place in the thrust nozzle (29,43) but, as indicated by a measured fuel specific impulse of 835 sec, a portion of the fuel did not burn. Thus, although the use of an ignition promoter had been avoided, the fuel specific impulse of hydrogen fuel had not been fully realised.…”

Chemical non-equilibrium flow analysis of H2 fueled scramjet nozzle

“…Nozzle is an important component of Scramjet engine, which is responsible for producing a major part of the thrust [1]. Chemical reaction inside a nozzle comes from two sources: 1, the unburned fuel in combustion chamber (if there is any) will continue their combustion reaction; 2, the dissociated species of combustion product, i.e., radicals, will recombine due to the rapid temperature drop inside the nozzle.…”

“…One-dimensional nozzle simulations with finite rate chemistry have been performed by Stalker et al [1], Sangiovanni et al [2], Ha [3], and Thomas et al [4], to evaluate the role of non-equilibrium chemistry on the performance of hypersonic nozzle. Different from the previous researches, this paper focuses on two new aspects: 1, the quantitative relationship among non-equilibrium flow, frozen flow, and equilibrium flow; 2, the influence of combustion incompleteness on the nozzle flow field.…”

Design of a 3D shape transitioning nozzle and experimental thrust measurements of an airframe integrated scramjet

“…A variation in Reynolds number of a factor of 2 had a negligible effect on thrust production. Stalker et al (2004) investigated the effects of chemical freezing of the exhaust gas and a continuation of combustion in the nozzle, a process they term kinetic afterburning, on nozzle performance. Numerical as well as experimental anlyses were conducted on a 2D model with a small thrust nozzle with an area ratio of 3.8.…”