The effect of a recessed liquid oxygen tube in shear coaxial injection has been investigated experimentally using an optically accessible subscale rocket combustor operated at pressures between 40 and 60 bar. Different single-shear coaxial injectors have been used to inject liquid oxygen and methane at relevant operating conditions covering sub-, near-, and supercritical pressures with respect to the critical point of oxygen. Liquid oxygen was injected at 120 K and the injection temperature of gaseous methane was about 275 K. Detection of spontaneous OH and CH chemiluminescene has been performed to characterize the flame-anchoring zone near the liquid oxygen post tip. In addition, the influence of the injector geometry on the combustion roughness and stability has been investigated during steady-state operating points. An increased flame expansion was observed with a recessed injector element. At low momentum flux ratio, the pressure drop accross the injector increases with a recessed liquid oxygen tube compared with a flush tube. Furthermore, a recessed liquid oxygen tube led to be a smoother combustion in general; however, this configuration also led to additional resonant frequencies in the chamber acoustics.

Effect of Recess in High-Pressure Liquid Oxygen/Methane Coaxial Injection and Combustion

The effect of recess length on the internal flow development and spray characteristics of a gas-liquid-swirl coaxial injector was investigated both theoretically and experimentally. Three different flow patterns were identified and explored as a function of the recess length. The spray-field dynamics and droplet-size distribution were measured by means of a high-speed photographic technique. An improved theoretical model was established to correlate experimental observations. An expression of the optimal recess length for atomization was obtained.

Spray Characteristics of Recessed Gas-Liquid Coaxial Swirl Injector

The LOX/Hydrocarbon Combustion Instability Investigation Program was structured to determine if the use of light hydrocarbon combustion fuels with liquid oxygen (LOX) produces combustion performance and stability behavior similar to the LOX/hydrogen propellant combination. In particular methane was investigated to determine if that fuel can be rated for combustion instability using the same techniques as previously used for LOX/hydrogen. These techniques included fuel temperature ramping and stability bomb tests. The hot fire program probed the combustion behavior of methane from ambient to subambient temperatures. Very interesting results were obtained from this program that have potential importance to future LOX/methane development programs. A very thorough and carefully reasoned documentation of the experimental data obtained is contained. The hot fire test logic and the associated tests are discussed. Subscale performance and stability rating testing was accomplished using 40,000 lb. thrust class hardware. Stability rating tests used both bombs and fuel temperature ramping techniques. The test program was successful in generating data for the evaluation of the methane stability characteristics relative to hydrogen and to anchor stability models. Data correlations, performance analysis, stability analyses, and key stability margin enhancement parameters are discussed.

LOX/Hydrocarbon Combustion Instability Investigation

Since the late 1960s, low hydrogen injection temperature is known to have a destabilising effect on rocket engines with the propellant combination hydrogen/oxygen. Self-excited combustion instabilities of the first tangential mode have been found recently in a research rocket combustor operated with the propellant combination hydrogen/oxygen with a hydrogen temperature of 95 K. A hydrogen temperature ramping experiment has been performed with this research combustor to analyse the impact of hydrogen temperature on the self-excited combustion instabilities. The temperature was varied between 40 and 135 K. Contrary to past results found in literature, the combustor was found to be stable at low hydrogen temperatures while increased oscillation amplitudes of the first tangential mode were found at higher temperatures of around 100 K and above, which is consistent with previous observations of instabilities in this combustor. Further analysis shows that hydrogen temperature has a strong impact on the combustion chamber resonance frequencies. By varying the hydrogen injection temperature, the frequency of the first tangential mode is shifted to coincide with the second longitudinal resonance frequency of the liquid oxygen injector. Excitation of combustion chamber pressure oscillations was observed during such events.

Influence of hydrogen temperature on the stability of a rocket engine combustor operated with hydrogen and oxygen

For shear coaxial injectors, which are widely used in liquid oxygen–gaseous hydrogen rocket engines, the key parameters that characterize injector geometry are the recess length, taper angle, and wall thickness of the liquid oxygen post, which is the inner tube of the coaxial injector. In the present study, the effects of the recessed length and taper angle of the liquid oxygen post on cryogenic coaxial jets at supercritical pressures were experimentally investigated. The large-scale structure and mixing enhancement of cryogenic nitrogen and gaseous helium coaxial jets were evaluated using a high-speed backlighting technique. Instantaneous backlit images revealed that the dark core, which is believed to be dense nitrogen, exhibited sinusoidal forms induced by the presence of the outer helium jets. However, the sinusoidal oscillations were less influenced by the injector geometry. The mixing enhancement of coaxial jets was evaluated with respect to the dark-core length and the jet spread angle. The recesse...

Effects of Injector Geometry on Cryogenic Shear Coaxial Jets at Supercritical Pressures

An experimental investigation was conducted to learn more about how the specific details of a concentric tube injection element affect the screech characteristics of a hydrogen-oxygen rocket engine. The four variables investigated were (1) impingement angle, (2) oxidizer tube blunt base thickness, (3) oxidizer tube recess and extension, and (4) oxidizer tube-annulus concentricity. Tests were made using a 27.34-cm (10.77-in.) diameter heat sink combustor at nominally 300-psia chamber pressure. All of the test variables were investigated using a 157 element circular pattern injector. Additional oxidizer tube recess tests were made with a 421 element hexagonal pattern injector. Tests were conducted over the oxidant-fuel ratio range of 4 to 6. Stability evaluation for each configuration was made using the hydrogen temperature rating technique. Several element configurations were also cold flow tested using nitrogen and water as simulants. The element detail changes resulted in changes in hydrogen injector pressure drop even though the physical injection area was constant for all similar tests in both hot firing and cold flow tests. These changes in injector pressure drop produced changes in combustion stability. The data were correlated with a modified version of a previously reported injection area ratio correlation. By interpreting changes in injector pressure drop as changes in injector hydrogen flow resistance, the data were compared with a hydrogen flow response stability model and were found to be in agreement.

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L. M. Russell

Papers

Some injector element detail effects on screech in hydrogen-oxygen rockets