Experimental Study on Combustion Modes in a Scramjet Engine

Accelerating scramjet engines for an access to space system should optimally anchor dual-mode combustion at one end of the trajectory and supersonic combustion at the other. Conventionally, dual-mode combustion engines flamehold with physical obstructions, causing prohibitive losses at high speed. Past experiments of an airframe-integrated scramjet engine at Mach 8 showed that dual-mode combustion was initiated and sustained with an unobstructed flowpath by varying the fueling rate. This flowfield is examined numerically to ascertain the fluid dynamic phenomena leading to the separated combustion modes. Two regimes of dual-mode combustion occurred above stoichiometric fueling: a separation-train-dominated flowfield downstream of the injectors and a flowfield dominated by a large separation upstream of the body-side injectors at higher fueling rates. The first mode occurs when two shocks combine and create a separation on the body-side wall that merges with the body-side injector wake separation, causing a train of separations. The second regime is from the upstream movement of combustion with increasing equivalence ratio: when combustion reaches the rear of the body-side injector, the near injector flow structure changes, allowing fuel and combustion to enter a large separation upstream. The performance of the different modes is explored.

Combustion modes of hydrogen jet combustion in a cavity-based supersonic combustor

The intensive combustion was observed in the subscale scramjet engine combustion tests, and the engine in the mode produced the large thrust. The mechanism should be made clear to apply the intensive combustion mode to the practical engine. The following mechanism was proposed in the present paper. The normally injected fuel interacted with the boundary layer behind the backward-facing step. There appeared the re-circulation region between the step and the fuel jet. The re-circulation region also appeared behind the fuel jet. This region affected the combustion condition greatly. The mixing and the combustion occurred not only in the supersonic primary flow, but also in the re-circulation region. It caused the heat release near the fuel injector, and the high combustion efficiency. Introduction The scramjet, one of the engines that will be used in the aerospace plane, is being studied at the National Aerospace Laboratory (NAL), Japan, and the sub-scale scramjet engine models have been tested in the Ramjet Engine Test Facility (RJTF). In the tests, several inflow air conditions, several engine configurations, and several fuel injection conditions were investigated. In other tests, the effects of radicals and water vapor were discussed.' In the experiments of the sub-scale scramjet engine model, there were two combustion modes, i.e., (1) the intensive combustion mode and (2) the weak combustion mode.' When the fuel flow rate was small, there was the weak combustion mode, and thrust was not produced. As the fuel flow rate was increased, the engine jumped into the intensive combustion mode. Then the thrust was large, and the net thrust was positive in several cases. The intensive combustion should be utilized in the scramjet engines. The research engine model used in the RJTF is different from the practical engine to be used in the aerospace plane. Therefore, the features and the mechanism of the intensive combustion must be made clear for its attainment in the practical engine. The mechanism of the intensive combustion was under discussion and it has not been made clear yet. In this paper, my opinion on the intensive combustion was presented. In the RJTF tests, the inflow air conditions correspond to the flight Mach numbers of 4, 6, and 8. There are many experimental results in the Mach 4 and 6 test conditions. The present discussion was based on the results of the Mach 4 and 6 test conditions. Experimental Apparatus and Methods A. Test facility The test conditions in the RJTF are shown in Table 1. They correspond to the flight conditions of Mach 4 and Mach 6 with a flight dynamic pressure of 65 kPa. Compressed air was heated by a ceramic storage heater. The supersonic nozzle had exit dimensions of 5 1 cm by 5 1 cm. In the Mach 4 test condition, the boundary layer thickness was 35 mm at the engine entrance. In the Mach 6 test condition, it

Quite different processes of ignition and combustion flowfield formation in a Mach 3 supersonic combustor depending on fuel injection and mixing schemes are observed by high speed schlieren video and wall pressure measurement. Experiments are conducted at a Mach 8 simulated flight condition using the High Enthalpy Shock Tunnel of JAXA, Japan. Parallel (12°) injection scheme of gaseous hydrogen fuel with generation of streamwise vortices by Hypermixer (HM) injector is examined and compared with the two typical normal (perpendicular) injection schemes, which are with and without backward facing step. Equivalence ratio is varied from Φ = 0.3 to 1.5. In the case of HM injector, the induction time increases as increasing Φ, to gradually push the ignition point far downstream of the injector. In larger Φ of 1.0 and 1.5, transient processes are observed, where a strong pressure wave generated by the explosive combustion of the well premixed gas at the combustor exit propagates upstream and decays around the injector to form a new quasi‐steady flowfield of a supersonic combustion mode downstream of the injector. In this new flowfield, combustion starts near downstream of the injector showing rather the features of diffusion flame combustion. The flow is choked upstream near the combustor exit and is accelerated to supersonic again in the nozzle. The properties of the observed pressure wave agreed well with those of Chapman‐Jouguet detonation wave on the basis of 1D‐flow analysis indicating that this pressure wave is a kind of detonation wave. For normal injectors, similar upstream propagation of the pressure wave is observed, but it propagates far upstream of the injector to form subsonic combustion mode. Comparing the characteristics of the pressure wave propagation for the injectors, some considerations are made on the mechanism of the transient processes, focusing on the roles of streamwise vortices to effectively operate the supersonic combustion.Quite different processes of ignition and combustion flowfield formation in a Mach 3 supersonic combustor depending on fuel injection and mixing schemes are observed by high speed schlieren video and wall pressure measurement. Experiments are conducted at a Mach 8 simulated flight condition using the High Enthalpy Shock Tunnel of JAXA, Japan. Parallel (12°) injection scheme of gaseous hydrogen fuel with generation of streamwise vortices by Hypermixer (HM) injector is examined and compared with the two typical normal (perpendicular) injection schemes, which are with and without backward facing step. Equivalence ratio is varied from Φ = 0.3 to 1.5. In the case of HM injector, the induction time increases as increasing Φ, to gradually push the ignition point far downstream of the injector. In larger Φ of 1.0 and 1.5, transient processes are observed, where a strong pressure wave generated by the explosive combustion of the well premixed gas at the combustor exit propagates upstream and decays around the inje...

In this paper, a novel two-stage hydrocarbon fuel heating and delivery system was used to more accurately simulate the state of fuel in the cooling channel of the regenerative cooling engine. Mildly cracked (<10%) experiments of n-dodecane were conducted by using the new heating system and pyrolysis products were sampled and analyzed. The ignition delay time and laminar flame speed of the cracked products were different from those of the surrogate gas fuel. When the temperature is 1100 K, the ignition delay time of the cracked products are ~50% higher than that of surrogate gas fuel. The laminar flame speed of the cracked products is about 30% lower than that of surrogate gas fuel at the equivalence ratio of 1.0. The supersonic combustion experiments of mildly cracked and different equivalence ratios n-dodecane were also employed at Mach number 3.0 of the isolator entrance, with total temperature ~1550 K and total pressure ~1.5 MPa. The static pressure along the axis of combustor was measured, while the velocity and temperature of the airflow at the exit of combustor were also measured by the tunable diode laser absorption spectroscopy (TDLAS). The distributions of airflow velocity, static temperature, and total temperature along the combustor at different conditions were analyzed through a one-dimensional analysis method. The deviations of velocity and temperature of the airflow are less than 4% and −6%, respectively, which indicates the calculation is in great accordance with the TDLAS detections. The experimental data will be helpful for validating the simulation of supersonic combustion, and studying the influence of mildly cracked fuel on supersonic combustion characteristics.

Experiments on a two-phase rotating detonation combustor operating with gasoline and high total temperature air were conducted to investigate the initiation characteristics, operation mode, and propagation characteristics of a two-phase rotating detonation wave (RDW). The outer diameter, inner diameter, and length of the annular combustor were 204 mm, 166 mm, and 155 mm, respectively. The initiation characteristics, operation mode, and propagation characteristics of the two-phase RDW were studied by varying the total air temperature. The experimental results show that the initiation time of the RDW first decreases and then increases with an increase in the total air temperature and reaches an extreme value at a total air temperature of 713 K. Four operation modes (failure, intermittent detonation, single wave, coexistence of double wave collision, and single wave) of the detonation combustor were found for different total air temperatures. The effect of the total air temperature on the peak pressure stability and propagation frequency of the RDW was studied in detail. From the results, the effect of the equivalence ratio on the working characteristics of a rotating detonation engine (RDE) was investigated at a total air temperature of 713 K. Four detonation propagation modes (sporadic detonation, intermittent detonation, single-wave mode, coexistence of double-wave collision and single wave) were obtained in the combustor. When the equivalence ratio was 0.52, the detonation initiation failed. The pressure characteristics in the combustor and propagation frequency of the RDW were studied with different equivalence ratios. In addition, a long-duration test was performed for 3 s to verify the continuous working feasibility of the two-phase RDE.

An experimental study was performed to characterize the effects of vitiation due to combustion-air preheating on dual-mode scramjet combustion.Major combustion vitiation species (H2O andCO2)were added to the freestreamof an electrical-resistance-heated, direct-connect facility simulating Mach 5 flight enthalpy. With clean, dry air, the combustor operated in the supersonic mode at fuel equivalence ratios below 0.22, and in the subsonic mode for equivalence ratios above 0.26. Hysteresis was observed in the dual-mode transition region between 0.22 and 0.26, as the mode of combustion was dependent on whether the fuel rate was increasing or decreasing. Adding increasing amounts of water vapor and carbon dioxide to the freestream decreased combustor pressures by 10 to 30% for the same fuel equivalence ratio. Vitiation also caused transition between supersonic and subsonic combustion to occur at a higher fuel equivalence ratio thanwith clean air. This work represents the first direct evaluation of the effect of testmedium vitiation on dual-mode scramjet combustion atMach 5 enthalpy simulation in the same facility. The results indicate the importance of accounting for test-medium vitiation when extrapolating from ground-testing to flight, particularly in the dual-mode transition region between subsonic and supersonic combustion regimes.

Ignition dynamics and combustion mode transitions in a rocket-based combined cycle combustor operating in the ramjet/scramjet mode

The scramjet appears to be among the most promising propulsion solutions for the future, attracting attention of many researchers and experts. The major concern is the complicated mixing of air and fuel phenomena within the combustor. The supersonic flow inside combustor causes improper air and fuel mixing in scramjet engines. The flame in stabilizes due to less air resident time, affecting output efficiency and resulting in fuel energy loss. The primary purpose of previous researches is to increase the performance of scramjet engines by improving supersonic mixing and combustion. Furthermore, beyond previous research endeavors investigating diverse fuel injection techniques within the supersonic combustor, this paper will delve into multiple fuel injection designs aimed at enhancing the combination of air and fuel.Additionally, it will demonstrate the reacting flow attributes associated with various ramp positions in a scramjet combustor featuring strut injection. Reduced fuel consumption in the combustor also exerts a considerable influence on efficient use of fuel energy in aviation field. A strut-based injector appears as an excellent design solution for a supersonic combustor. Conversely, a complex aero-thermodynamic system incorporating high-speed flow, compression waves, and thermochemical reactions represented by the scramjet combustor. The performance of a supersonic combustor is affected by numerous factors, such as the input Mach number, the type of fuel used, combustor pressure, and temperature.

High-Repetition-Rate OH Planar Laser-Induced Fluorescence of a Cavity Flameholder

This study investigates an enhanced passive control strategy for suppressing combustion oscillations in scramjet engines by introducing an auxiliary small cavity upstream of the primary flameholder. Experimental tests conducted on a direct-connect rig with Mach 3 inflow demonstrate that the staged-cavity design achieves remarkable stability improvements across equivalence ratios (Φ = 0.34–1.2). Key findings include: (1) 48.9% and 47.2% reduction in pressure oscillations at Φ = 1.2 and 0.6, respectively, with the mean normalized pressure increasing by approximately 8.57%; (2) Expansion of flammability limits (lean ignition at Φ = 0.34 vs baseline Φ = 0.48); (3) 81.2% suppression of flame front fluctuations through standing flame anchoring. Mechanistic analysis reveals five synergistic effects: (i) pre-combustion in the auxiliary cavity decentralizes heat release, (ii) the small cavity's shear layer sustains a persistent pilot flame without additional injectors, (iii) boundary layer stabilization prevents separation-induced oscillations, (iv) enhanced fuel–air mixing creates localized premixed zones within the diffusion flame, and (v) shock restructuring maintains flame anchoring despite pressure variations. High-speed imaging and schlieren visualization confirm the design's ability to maintain stable jet-like flames while eliminating chaotic combustion modes observed in conventional configurations. The auxiliary cavity approach demonstrates simultaneous improvements in combustion stability (oscillation suppression), performance (thrust increase), and operational envelope (extended ignition limits), providing a practical passive control solution for scramjet combustors. Residual low-frequency fluctuations (4%–5% amplitude) suggest future optimization opportunities in cavity-flowfield integration. These findings establish staged cavity combustion as an effective strategy for next-generation hypersonic propulsion systems.

Numerical study on effect of air throttling on combustion mode formation and transition in a dual-mode scramjet combustor

Air-breathing propulsion has been regarded in recent years as the future solution for spacecraft launchers. Hypersonic flight of vehicles designed for horizontal take-off remains nevertheless a challenging task. Especially the propulsion unit must work over a broad range of flight Mach number: efficiency can be guaranteed only with a system integrating conventional and air-breathing engines and the latter have to be suitable for dualmode operation. The Institute of Flight Propulsion of the Technische Universitat Munchen developed an air-breathing combustor concept, including a novel injection system. The combustion chamber consists of a constant cross section module hosting a strut injector followed by a diverging module which counteracts static pressure peaks consequent to combustion. Hydrogen and air are injected unlike-impinging into the strut wake, where a cylindrical pipe is inserted as flame-holder. The system allows creating a flow recirculation area. Combustion radicals are set free and promote further fuel reaction until a pilot flame stabilizes. Additional hydrogen or methane is injected through the strut sides, pre-mixes with the supersonic air flow (M=2.2) stoked to the combustor and is ignited by means of the pilot flame. The total temperature at the combustor entrance can be varied between 500K and 1200K. First results of wall static pressure measurements and Schlieren imaging demonstrated the possibility of dual-mode combustion. The transition between ramjet and scramjet modes showed to depend on the fuel total injection pressure. This paper presents the results of experiments carried out to define the fuel total injection pressure interval in which the transition occurs. Two injector versions of different thickness (i.e. 3mm and 5mm) have been used. This allowed evaluating the effects of the aerodynamic blockage due to the strut on the overall combustion process, particularly on the transition. ∗ Dipl.-Ing. Sara Rocci Denis, Research Assistant, Institute of Flight Propulsion † Professor Dr.-Ing. Hans-Peter Kau, Director, Institute of Flight Propulsion ‡ Dipl.-Ing. Armin Brandstetter, Systems Engineer, MT12 Mission Dynamics NOMENCLATURE MCC Combustor entrance Mach number MF Flight Mach number pCC Combustor entrance static pressure ptH2A Total pressure of hydrogen injected through system A pW Combustor wall static pressure TSS Flame-holder surface temperature TtCC Combustor entrance total temperature TtH2 Hydrogen injection total temperature φ Equivalence ratio

A numerical study on combustion mode characterization for locally stratified dual-fuel mixtures

A series of numerical simulations were performed to investigate the feasibility and performance of the premixed ammonia/hydrogen/air rotating detonation engines. A 19 species and 80 reactions ammonia/hydrogen/air mechanism is adopted and validated for detonation simulations. The effects of injection total temperatures (T0) and ammonia/hydrogen equivalence ratios (φNH3 and φH2) are analyzed under a fixed global equivalence ratio of 1. The propagation map of rotating detonation waves is numerically outlined. The result indicates that a higher injection total temperature and a lower ammonia equivalence ratio are beneficial to the successful propagation of rotating detonation waves. The maximum φNH3 with successful propagation of rotating detonation waves reaches 0.6, achieved at T0 = 1000 K. High total temperatures and ammonia equivalence ratios can lead to lower detonation wave speeds. The detonation height is found to account for around 20%–36% of the engine axial length. The critical accommodated detonation cell number for successful propagation of rotating detonation waves is 5.9, below which the rotating detonation wave will have difficulty maintaining propagation. Mass-flow-averaged and area-averaged methods are adopted to evaluate the pressure gain performance of NH3/H2/air RDE. The results of the two methods both indicate that the total pressure gain is significantly affected by the injection total temperature but less affected by the equivalence ratio of NH3. In addition, it is found that NOx emission is dominated by NO. The NOx emission increases with increased injection total temperatures and ammonia equivalence ratios. Negligible NOx emission is produced in pure hydrogen-fueled RDE while it reaches the maximum (0.037) at φNH3 = 0.6 and T0 = 1000 K.