Combustion-induced Pressure Rise Research Articles

Noise Generated in a Scramjet Combustor

This paper presents the noise levels generated inside a two-dimensional scramjet combustor at Mach 7.3 freestream using the focused laser differential interferometric (FLDI) technique. FLDI measurements were taken at the front, middle, and rear of the combustor’s center plane under unfueled, combustion-suppressed, and combustion-on conditions. Spectral analysis of the FLDI signal shows density fluctuations inside the scramjet combustor are almost two times higher than the freestream and confirms the capability of this technique for measuring the disturbances in a complex hypersonic flowfield. Though combustion-induced pressure rise is low, a trend in the increase in the normalized density fluctuations/noise levels due to supersonic combustion is observed across the measurement locations for the combustion-on condition. However, the overall increment in the noise level is modest between the three conditions and across the three probing locations. This infers that, for the current scramjet geometry, fueling rates, fuel–air mixing, and supersonic combustion are not the dominant noise sources in the scramjet combustor. Instead, the base hypersonic flowfield inside the scramjet combustor is the main contributor to the noise field. Noting that there is relatively large uncertainty in the measured noise levels and that the combustion-induced pressure rise is modest in the present study, further investigation is needed to check the applicability of these results to the scramjet engines with flow paths having more complex flowfields, fueling schemes, and fueling rates.

Thermal compression effects within a fundamental, hydrogen-fuelled scramjet

The impact of thermal compression on combustion has been studied experimentally in a hydrogen-fuelled, non-uniform scramjet flowpath. The experimental model consisted of a three-dimensional inlet, a constant area rectangular combustor, and a single-ramp expansion nozzle, all having a constant width, and was tested at an equivalent Mach 10 flight condition with 58 kPa dynamic pressure. Combustion was suppressed in different regions of the combustor by injecting helium as a replacement for hydrogen fuel in either spanwise half of the engine. Thermal compression effects increased the combustion-induced pressure rise by (11.9±5.8)% across the majority of the combustor for an equivalence ratio of 0.8, however a shock-related artefact dominated similar measurements for a case with an equivalence ratio of 1.0. Over a smaller region at the end of the combustor, the combustion-induced pressure rise was increased by (9.4±9.2)% and (6.7±6.8)% for equivalence ratios of 0.8 and 1.0, respectively. Time-integrated OH emission signals increased by 19% and 31%, whilst time-resolved signals increased by 54% and 35%, for equivalence ratios of 0.8 and 1.0, respectively. This study presents the first experimental evidence of thermal compression directly increasing combustion-induced pressure rise in a scramjet engine.

Fundamental Scramjet Combustion Experiments Using Hydrocarbon Fuel

Combustion of ethylene was examined experimentally via semi-freejet testing of a fundamental, axisymmetric combustor model at conditions representative of Mach 7–8 flight at 100 kPa dynamic pressure. The model consisted of diffuser, constant area isolator, cavity, and diverging combustor sections. Pressure data were measured along the wall of the model and were complemented with OH planar laser-induced fluorescence (PLIF) visualization at the exit of the diverging section. Robust combustion of ethylene was observed for two stagnation enthalpy cases for fuel-air equivalence ratios of 0.3 or greater. A low-enthalpy, low-equivalence-ratio case showed significant, but structurally different, OH PLIF signals even though there was little combustion-induced pressure rise. Edge-detection techniques were applied to the PLIF images, and the resulting circumferential distributions analyzed. Mean radii and standard deviations are presented for the inner and outer edges of the OH-rich band for four tunnel/fueling conditions. Finally, amplitude spectra of fluctuations in the edge radii are shown, indicating that low-wavenumber oscillations are dominant. These experimental results provide useful data for comparison with numerical simulation in future studies.

Supersonic combustion of hydrocarbons in a shape-transitioning hypersonic engine

This paper presents the results of a series of shock tunnel experiments of a hydrocarbon-fuelled, Mach 8 shape-transitioning scramjet engine with a cavity combustor. The inlet of the scramjet transitions from a quasi-rectangular capture area to an elliptical throat, which results in a highly three-dimensional flowfield at the combustor entrance. The main focus of the work was to achieve ignition and combustion of hydrocarbon fuels at a high Mach number in a flight-candidate engine that has the necessary three dimensional flow path that is typical of “practical” scramjet engines. The engine was fuelled with a surrogate fuel mixture (64% ethylene and 36% methane by volume) which mimics the extinguishing characteristics of partially cracked JP-7 fuel. Experiments were performed with and without a hydrogen-pilot to demonstrate ignition and combustion. Static pressure measurements throughout the flow path were used in conjunction with experiments where combustion was suppressed, by using nitrogen instead of air as the main flow, to confirm a combustion-induced pressure rise for different injection and piloting strategies. Ignition and supersonic combustion of the surrogate fuel mixture was achieved in all cases where a hydrogen pilot was employed. The pilot hydrogen fuel was injected at the inlet of the engine. Simultaneous injection of the surrogate fuel mixture at the inlet and in the combustor (combined injection) and combustor-only injection did not achieve combustion.

Investigation of rotating detonation combustor operation with H2-Air mixtures

The operating range and wave speed performance of a Rotating Detonation Combustor (RDC) is characterized for hydrogen-air mixtures for three fuel injection schemes and two air injection schemes. The fuel injection scheme is altered by changing the total number of injection orifices and the individual orifice area, while maintaining the same fuel mass flux across the three schemes. The operability, performance and combustion-induced pressure rise due to the addition of a back-pressurizing convergent nozzle is also characterized. While the operating range is largely unaffected by changes in the length-to-diameter ratio of the fuel injector orifices, higher length-to-diameter ratios correspond to a lower number of transitional RDC operation where there is a sudden abatement of the continuously propagating detonation wave, once established inside the combustor. Increased air injection area diminishes the operability, while producing high stochasticity in the performance of the RDC. The length-to-diameter ratio of the fuel orifices has a significant impact on the number of detonation waves that can exist in the chamber. For the highest length-to-diameter ratio of the fuel orifices, and at the highest air flow rates, the RDC supports multiple detonation waves inside the chamber. Without the convergent nozzle attachment, 80% of Chapman–Jouguet (C–J) detonation speed is achieved for all three fuel injection schemes. C–J detonation wave speed is achieved in the annulus when the RDC is back-pressurized using the nozzle. The ratio of reactant fill-height to the detonation cell-width tapers at the lean and rich operating conditions, while peaking at an equivalence ratio of around 1.2. The detonation-induced static pressure rise produced in the RDC is found to be dependent on the air flow rate and the equivalence ratio of the reactants.

Combustion enhancement in a scramjet engine using oxygen enrichment and porous fuel injection

AbstractThis paper reports on the experimental testing of oxygen-enriched porous fuel injection in a scramjet engine. Fuel was injected via inlet mounted, oxide-based ceramic matrix composite (CMC) injectors on both flow path surfaces that covered a total of 9.2 % of the intake surface area. All experiments were performed at an enthalpy of $3.93{-}4.25\pm 3.2\,\%~\text{MJ}~\text{kg}^{-1}$, flight Mach number 9.2–9.6 and an equivalence ratio of $0.493\pm 3\,\%$. At this condition, the engine was shown to be on the verge of achieving appreciable combustion. Oxygen was then added to the fuel prior to injection such that two distinct enrichment levels were achieved. Combustion was found to increase, by as much as 40 % in terms of combustion-induced pressure rise, over the fuel-only case with increasing oxygen enrichment. Further, the onset of combustion was found to move upstream with increasing levels of oxygen enrichment. Thrust, both uninstalled and specific, and specific impulse were found to be improved with oxygen enrichment. Enhanced fuel–air mixing due to the pre-mixing of oxygen with the fuel together with the porous fuel injection are believed to be the main contributors to the observed enhanced performance of the tested engine.

Experimental investigation of vitiation effects on supersonic combustor performance

A significant problem for use of combustion heated facilities in ground laboratory studies of scramjet propulsion is that the resulting high enthalpy test air are seriously vitiated by several species, for example, H2O, CO2, CO, H, OH, O, and NO, which are not of representative or very few in actual atmosphere, so-called vitiation air relative to pure air. Combustion in such vitiation air stream can be influenced by chemical and physical effects due to the different species from actual atmosphere. Therefore, the ground-test results from such vitiated facilities should be properly analyzed and corrected before extrapolated to atmospheric flight condition. The primary goal of the present efforts is to assess the net effects of vitiation air on combustion process in a supersonic combustor. Based on the direct-connected test facility of Northwestern Polytechnical University, an experimental system is developed for comparative investigation of supersonic combustion in vitiation airstream and clean airstream, respectively. Specific species at well-controlled concentration are added to the clean airstream generated from resistance heater to synthesize the vitiation airstream, duplicating the test media in a combustion heated facility. The air total temperature at combustor entrance is about 850K, typically simulating the Mach 4 flight condition. Details of the experiment system are present in this paper. With the newly constructing system, hydrogen, ethylene and kerosene fueled supersonic combustions with clean air and vitiation air stream are investigated. Individual and combined influences of H2O and CO2 at various concentrations are considered over a range of experiment condition. The combustion characteristics with clean and vitiation air stream are compared, and the influences of H2O and H2O/CO2 on supersonic combustion processes are discussed. Results show that, the combustion induced pressure rise can be significantly inhibited by H2O and/or CO2 vitiation . The direct extrapolation of vitiated test results to flight condition may possibly result in over-fueling combustion, even inlet unstart. The H2O and/or CO2 vitiation also influence the shock waves in the isolator, resulting in the tendency of supersonic combustion mode.

Experimental and Numerical Studies of Vitiated Air Effects on Hydrogen-fueled Supersonic Combustor Performance

This paper deals with the vitiation effects of test air on the scramjet performance in the ground combustion heated facilities. The primary goal is to evaluate the effects of H2O and CO2, the two major vitiated species generated by combustion heater, on hydrogen-fueled supersonic combustor performance with experimental and numerical approaches. The comparative experiments in the clean air and vitiated air are conducted by using the resistance heated direct-connected facility, with the typical Mach 4 flight conditions simulated. The H2O and CO2 species with accurately controlled contents are added to the high enthalpy clean air from resistance heater, to synthesize the vitiated air of a combustion-type heater. Typically, the contents of H2O species can be varied within the range of 3.5%–30% by mole, and 3.0%–10% for CO2 species. The total temperature, total pressure, Mach number and O2 mole fraction at the combustor entrance are well-matched between the clean air and vitiated air. The combustion experiments are completed at the fuel equivalence ratios of 0.53 and 0.42 respectively. Furthermore, three-dimensional (3D) reacting flow simulations of combustor flowpath are performed to provide insight into flow field structures and combustion chemistry details that cannot resolved by experimental instruments available. Finally, the experimental data, combined with computational results, are employed to analyze the effects of H2O and CO2 vitiated air on supersonic combustion characteristics and performance. It is concluded that H2O and CO2 contaminants can significantly inhibit the combustion induced pressure rise measured from combustor wall, and the pressure profile decreases with the increasing H2O and CO2 contents in nonlinear trend; simulation results agree well with experimental data and the overall vitiation effects are captured; direct extrapolation of the results from vitiated air to predict the performance of actual flight conditions could result in over-fueling the combustor, possible inlet un-start and inappropriate combustion mode transition. The detailed analysis and discussion are presented and the research conclusions are summarized.

OH-PLIF visualisation of radical farming supersonic combustion flows

Experimental investigations employing Planar Laser-induced fluorescence visualisation of the qualitative distribution of the OH radical (OH-PLIF), coupled with surface pressure measurements, have been made of flow in a generic, nominally two-dimensional inlet-injection radical farming supersonic combustion scramjet engine model. The test flows were provided by a hypersonic shock tunnel, and covered total enthalpies corresponding to the flight Mach number range 8.7–11.8 and approximately 150 kPa dynamic pressure. The surface pressure measurements displayed radical farming behaviour, that is a series of adjacent high and low pressure regions corresponding to successive shock/expansion structures, with no significant combustion-induced pressure rise until the second structure. OH-PLIF imaging between the first two structures provides the first direct experimental evidence of significant OH radical concentrations upstream of the ignition point in this mode of scramjet operation and shows that combustion reactions were occurring in highly localised regions in a complex turbulent and poorly micromixed fuel/air mixing layer confined to the fuel injection side of the combustor.

Boundary-Layer Separation Due to Combustion-Induced Pressure Rise in a Supersonic Flow

The HyShot supersonic combustion experiments were conducted in the T4 Stalker tube at the University of Queensland, Australia, to test supersonic combustion in a constant-area duct of rectangular cross section. The intake was deigned to spill both the compression shock and the boundary layers formed on the intake surfaces and to generate new boundary layers on the four walls of the combustor. Fuel was injected perpendicularly to the oncoming flow from four 2mm porthole fuel injectors located 40mm from the leading edge of the combustor. Choking of the combustor, attributed to boundary-layer separation due to combustion-induced pressure rises, is identified to occur within the test period in 12 fuel-on shots. The 95% confidence-interval uncertainty is estimated to be ∓11%, taking into account the uncertainty in flow conditions and pressure measurements.

Radical-Farm Ignition Processes in Two-Dimensional Supersonic Combustion

A two-dimensional numerical study was performed of the ignition processes associated with the concept of radical farming for supersonic combustion. In a preliminary parametric study, a range of freestream conditions attainable in a hypersonic shock tunnel was investigated and mapped according to whether or not the behavior known as radical farming is present: a combustion-induced pressure rise in second or subsequent hot pockets rather than the first Two such cases were analyzed in detail, both having mean conditions across the combustion-chamber entrance that would result in extremely long ignition lengths. The initiation, branching cycle, and heat release reactions in the combustion process become active in the radical farm, and H and OH radicals are produced. Their rate of production slows in the regions of flow expansion, but does not approach chemical freezing until toward the end of the localized expansion zones. Simultaneously, heat release elevates the local temperature. When the mixture flows through the shock at the second hot pocket, the elevated temperature and the presence of radicals enable the branching cycle and three-body recombination heat release reactions to accelerate, and significant pressure rise due to heat release is then able to occur. The extent to which this is completed in the second hot pocket depends on the inflow conditions.

Transverse and parallel injection of hydrogen with supersonic combustion in a shock tunnel

An experimental comparison has been made of the combustion induced pressure rise in a constant area duct when hydrogen is injected transverse to the flow by using a surface orifice, and when it is injected parallel to the flow by using a central injection strut.