Constant Area Combustor Research Articles

Spatiotemporal information propagation in confined supersonic reacting flows

The interplay between mass injection, heat release, and boundary layer development plays a key role in dictating the dynamics and stability of confined supersonic flows. The relative impacts of these factors and the timescales over which they influence the upstream and downstream flow can provide critical insights into how different operating modes develop. As such, this work presents a series of simulations of an experimental axisymmetric direct connect flowpath. The mass flow rates and chemical compositions of the injection stages are varied, and subsequent information propagation and mode transitions are analyzed using spatiotemporal correlations of cross-sectional-averaged quantities. Increasing the injection flow rate decreases the time lags and durations of positive correlations between pressure and heat release at various points along the flowpath. Meanwhile, in dual-mode cases with lower injection flow rates, these correlations develop after longer time delays and persist for a longer times, illustrating how information propagates more gradually in these scenarios. Over the full flowpath, positive correlations persist for comparatively long times between (1) the upstream isolator pressure and the pressure elsewhere, and (2) the pressure in the downstream diverging combustor section and the upstream pressure. As such, the influence of the pressure in the intermediate constant-area combustor section decays more rapidly. Conditional statistics suggest that flow blockage and pressurization from the injected mass reduce the local ignition delay, thereby facilitating increased pressurization via heat release in a positive feedback loop.

Research on the operating boundary of the dual mode scramjet with a constant area combustor through thermodynamic cycle analysis

The maximum pressure ratio of the isolator is important for studying the operating boundary of the dual mode scramjet. In this paper, the pressure rise mechanism of the dual mode scramjet engine and the characteristic of the isolator maximum pressure rise are analyzed. The operating range and performance of scramjet under the pressure rise limit are investigated using the constant cross-sectional area combustor. The maximum pressure ratio of the isolator confines the back pressure from the combustor so that the scramjet engine has a minimum freestream Mach number of each equivalence ratio. When the freestream Mach number is smaller than the minimum freestream Mach number, an expansion section may be arranged between the combustor and the isolator to balance the pressure of the combustor and that of the isolator. The expression of the minimum freestream Mach number is derived and it is found that the minimum freestream Mach number of the dual mode scramjet in both modes increases with the increase of equivalence ratio and compression ratio of the inlet. A variable divergent section is suggested to balance the pressure under different freestream Mach numbers and equivalence ratios.

Ethylene flame-holding in double ramp flows

In this work, ethylene flame-holding in supersonic flows was investigated using a shock tunnel. The experiments were conducted at flow stagnation temperatures ranging from 1270 to 1810 K. The two-dimensional test model consisted of a double-ramp inlet and a constant-area combustor with a recessed wall cavity. The two fuel injectors were located at the inlet and the other upstream of the cavity. Shadowgraph and flame chemiluminescence images were captured for optical visualization. The inlet injection images showed various flame-holding patterns. At 1270 K, the flame was not maintained. At 1540 K, the flame was maintained inside the cavity, and the condition provided continuous combustion during the steady flow. At 1810 K, strong flame signals were observed from the inlet to the cavity and downstream. At 1540 K, the inlet injection with a low fuel pressure showed a gradual flame quenching in the cavity during flow establishment. On the other hand, the same injection in the combustor showed flame-holding in the shear layer above the cavity. The results showed that the flame patterns are strongly influenced by the flow stagnation temperature and the location of fuel-injection.

Scramjet performance for ideal combustion processes

Models used to analyse scramjet engine cycles are typically either very simple or focused on specific combustor geometries. This paper uses a quasi-one-dimensional, inviscid chemical equilibrium solver to examine scramjet engines defined by predetermined combustion processes. This solver is capable of rapidly analysing combustion processes without requiring predefined geometries and is validated against the Hyshot II flight experiment and a constant area combustor ground test. Combustion occurs using constant area, constant pressure, constant Mach number or constant temperature processes. Constant area combustion typically produces the highest specific impulse for given combustor entrance conditions. Maximum pressure and temperature limitations were introduced, and multiple engines with combinations of combustion process were examined. It was found that engines with a combination of all four combustion processes can have lower maximum values of pressure and temperature whilst maintaining high performance compared with constant area combustor engines.

Flow Establishment of Precombustion Shock Trains in a Shock Tunnel

This work is a fundamental investigation of the behavior of precombustion shock trains in the T4 Stalker Tube at the upper end of the dual-mode combustion regime. Experiments were conducted at a condition that represents flight at Mach 8 at an altitude of 26 km. The test model was a simple axisymmetric duct comprising a short diffuser, an isolator (), and a constant-area combustor (). Gaseous hydrogen was used as the fuel and was injected via six portholes equispaced around the perimeter of the duct. Time histories of the surface pressure in the model showed a pressure rise from stable supersonic combustion up to a fuel equivalence ratio of 0.9. At higher fuel levels, a precombustion shock train formed upstream of the point of fuel injection. Between equivalence ratios of 0.9 and 1.1, the shock train established itself and remained stable over a significant portion of the approximately 3 ms test window. Above an equivalence ratio of 1.1, the pressure time histories continued to rise throughout the test window and a stable flow was not established. These results indicate that short-duration facilities like the T4 Stalker Tube can be used for dual-mode combustion studies at fuel levels below a limit that will depend on the particular flow condition and model geometry.

Drag Reduction by Boundary-Layer Combustion: Effects of Flow Disturbances from Rectangular-to-Elliptical-Shape-Transition Inlets

A key issue that can affect the implementation of boundary-layer combustion for viscous drag reduction in scramjet combustors is the potential of nonuniform flow entering the combustor from the inlet. Experiments were conducted in the T4 Stalker tube to study the effects of these flow nonuniformities on boundary-layer combustion in a circular constant-area combustor. A rectangular-to-elliptical shape-transition inlet, typical of the type proposed for self-starting scramjets, was designed, manufactured, and attached upstream of the combustor in the current experiments. Numerical simulations of the flow through the rectangular-to-elliptical shape-transition inlet showed that the inlet generated a nonuniform outflow that had vortices, shock waves, and expansion waves. When tested at design-point inflow conditions, the total skin friction drag measured in the combustor was reduced by 61% when hydrogen burned in the boundary layer. These levels of reduction are similar to those measured in experiments conducted without a realistic scramjet inlet upstream of the combustor, thus demonstrating that the drag reduction brought about by boundary-layer combustion is not affected by the flow disturbances generated from the REST inlet. Tests conducted at off-design conditions further demonstrated that the drag reduction brought about by boundary-layer combustion was not affected by changes in the inflow conditions.

An Aerothermodynamic Design Approach for Scramjet Combustors and Comparative Performance of Low-Efficiency Systems

This paper is aimed at development of an integrated approach based on analytical and computational aerothermodynamics for the special case of design of a 75% (low process-efficiency), hydrogen-fuelled, constant area combustor of a hypersonic airbreathing propulsion (HAP) system thereafter undertaking study of two types of HAP systems. The results of configurational aerothermodynamics implied that the most appropriate constant area configuration had a 30 degrees downstream wall-mounted fuel injector with a single acoustically stable cavity placed downstream of the fuel injection point. Moreover for identical flow inlet parameters and system configurations at lower levels of thermodynamic process efficiencies, the constant combustor-area (i.e. Scramjet 1) engine is superior in its performance to the constant combustor-pressure (i.e. Scramjet 2) engine for all values of fuel-air ratios.

Experimental Study on Combustion Modes in a Supersonic Combustor

obliquefuelinjectionwasemployedtomitigatefueljet/airflowinteraction,theinteractionwasfoundtohaveasizable effect on the ignition limit. A simple model to predict the minimum combustor length needed to attain supersonic or dual-mode combustion was suggested, and it agreed relatively well with the experimental results. The model was applied to experimental results of a subscale scramjet engine tested under Mach 8 flight conditions at the Japan Aerospace Exploration Agency, Kakuda Space Center, in order to predict the occurrence of combustion (large pressure rise) within a constant area combustor in the engine, and it agreed well with the experimental results.

Measurement of Drag on a Scramjet Engine in a Shock Tunnel

The net drag on a non-fuelled, internal compression, constant-area combustor scramjet engine was measured using a single-component accelerometer balance in a shock tunnel at a freestream Mach number of 8 and a flow total enthalpy of 1.35 MJ/kg. The flow fields of the model were simulated using a commercial CFD code in order to understand the aerodynamics of the engine and compare with the measured net drag co-efficient. Measured and computed values of the net drag co-efficient were found to be in a good agreement.

Mach-8 Tests of a Combined-Cycle Engine Combustor

Under the Mach-8 flight condition, combined-cycle-engine combustor tests were conducted. The combustor model had a two-rocket fuel injector, which supplied fuel-rich, precombustion gas as fuel. Gaseous hydrogen and oxygen were used as propellants. Two combustion patterns were observed for various lengths of the straight duct connected to the rocket injector section. In the longer straight-duct configurations, combustion gas was decelerated to subsonic speed and choked thermally at the exit of the straight duct. In the shorter straight-duct configurations, combustion progressed in the downstream divergent duct. Pitot pressure and gas-sampling parameters were measured at the exit plane of the combustor model. Results showed that a certain length of a constant cross-sectional area combustor section was required to sustain combustion of the rocket plume injected parallel to the airflow.

Performance of a Dual-Mode Combustor with Multistaged Fuel Injection

A scramjet combustor, which has a constant cross-sectional area combustor (designated as combustor 1), and a diverging combustor (designated as combustor 2), was examined with both single and two-stage fuel injection to determine whether it could be operated in the ramjet mode. The combustor was directly connected to a facility nozzle, the Mach number at the exit of the nozzle being 2.4. Total temperature and total pressure were 800 K and 1.0 MPa, respectively. Pitot pressure measurements and gas sampling for an equivalence (Φ) of 0.2 showed that there were both a separated region and a supersonic core region within combustor 1 with single-stage fuel injection. For Φ = 0.4, on the other hand, a uniform subsonic region without separation was formed. One-dimensional calculation showed that two-stage fuel injection, in which the second-stage fuel was injected into combustor 2, led to formation of a large subsonic region. The maximum (dF, [thrust with fuel]-[thrust without fuel]) was 37% higher than that with single-stage fuel injection, this increment being limited by occurrence of combustor-inletinteraction in both cases. Another type of two-stage fuel injection, in which both stages were within combustor 2, was conducted, and the maximum dF was found to be almost the same as that with two-stage fuel injection within both combustor 1 and combustor 2. It was also found that dF could be described as a linear function of an effective equivalence ratio ([total equivalence ratio] x [combustion efficiency at the combustor exit]) regardless of the fuel injection configuration. The effect on dF of the difference in total pressure loss due to the difference in fuel injection configuration was negligible because the scramjet combustor was operated in the ramjet mode. Ideal was estimated by one-dimensional calculation to evaluate the achievement factor for thrust ([thrust obtained by experiment]/[ideal obtained by one-dimensional calculation]), which was found to be 60% at most for the tested combustor under the tested conditions.

M4飛行条件におけるスクラムジェットエンジンの燃焼器–インレット干渉抑制

A sidewall-compression-type scramjet engine was tested under M4 flight conditions. Tested engine had an inlet, a constant cross-sectional area isolator, a constant cross-sectional area combustor (designated as combustor-1), a diverging combustor (designated as combustor-2), and an internal nozzle. In the previous study under M4 flight conditions, the maximum thrust increment by the fuel injection within the combustor-1 was 1,380N at an equivalence ratio of 0.31, and further fuel injection resulted in the combustor-inlet interaction (designated as CII). To suppress the CII, we attempted (1) a two-stage fuel injection within the combustor-1 and the combustor-2 and (2) a boundary layer bleed on the top wall, in the present study. The former was to suppress heat release around the first-stage fuel injectors in the combustor-1 at a given total fuel mass flow rate, and the latter was to decrease interaction length by decreasing a boundary layer thickness on the top wall. With the two-stage fuel injection, the maximum thrust increment was 2,230N at an equivalence ratio of 0.63. Next, the boundary layer bleed was carried out, and the maximum thrust increment was 2,300N at an equivalence ratio of 0.66. Thus, both two-stage fuel injection and boundary layer bleed led to 60% higher maximum thrust increment than that obtained in the previous study. Finally, both two-stage fuel injection and boundary layer bleed were applied simultaneously to obtain the largest thrust performance, and the maximum thrust increment was 2,560N at an equivalence ratio of 0.95. As a result, we obtained 80% higher maximum thrust increment than that in the previous study by these methods. The thrust achievement factor, which was defined as the ratio of the maximum thrusts obtained from experiment and theoretical prediction, under this condition was estimated as 70%.

Experimental Study of Confined, Swirling, Nonpremixed Gas Flame for Validation of Simulations

An atmospheric-pressure, gaseous fuel combustor was developed to obtain data for validation of large-eddy or other high-fidelity simulations of confined combustor flow fields, such as found in gas-turbine engines. The airand fuel (methane) enter a constant-area combustor test section in separate concentric swirling jets through a 3:1 expansion. The test section is optically accessible, cylindrical in cross section, and four diameters in length. Measurements of two-component mean and rms velocity were obtained using laser Doppler velocimetry for both a combusting case and a nonreacting flow with identical inlet conditions. Measurements of major species (CO 2 , CO, O 2 , and hydrocarbons), minor species (NO), and mean temperature in the reacting flow were also obtained.

Distributed Fuel Injection for Performance Improvement of Staged Supersonic Combustor

Introduction S UBSCALE model scramjet engines have been tested at Japan Aerospace Exploration Agency—Kakuda Space Propulsion Center (formerly National Aerospace Laboratory—Kakuda Space Propulsion Laboratory).1 The models consisted of a sidecompression-type inlet, a constant (minimum) cross-sectional area isolator with steps at its exit, a constant cross-sectional area combustor with fuel injectors, a diverging combustor, and an internal nozzle. In the tests, occurrence of combustor-inlet interaction2 resulted in a limit to the fuel flow rate and the engine thrust. To mitigate this interaction, a staged combustor, which had firststage fuel injectors on a strut in the minimum cross-sectional area

スクラムジェットエンジンのマッハ４燃焼・不始動遷移模擬実験 エンジン内流れ場の観察

Combustion and unstart processes of a scramjet engine in Mach 4 flight condition were investigated experimentally focusing on the changes of the engine internal flow structures under the ‘weak’ and ‘intensive’ combustion modes. Cold flow tests were carried out using a 1/5-scale model of the scramjet engine examined in the firing tests at Ramjet Engine Test Facility (RJTF). Pressure rise due to combustion was successfully simulated by a flow-plug installed in the engine nozzle section to demonstrate good agreement with the wall pressure distributions of the engine under ‘weak’ and ‘intensive combustion’ modes. Especially, upper limit of the combustor maximum pressure at which engine fell into unstart was exactly simulated. This limit value was revealed to be the separation pressure of the turbulent boundary layer in the isolator. Shadowgraph visualization demonstrated that in the intensive combustion mode large-scale boundary layer separation was formed over the isolator, backward facing step and constant area combustor.

Control of longitudinal oscillations in a constant area combustor: Numerical simulation

The phenomenon of combustion driven oscillations has long held the interest of scientists and engineers. According to the so-called Rayleigh criterion [1], driving of the pressure oscillations occurs if the periodic heat release rate is generally coincident with the pressure oscillations. The main mechanisms involved are considered to be both thermoacoustic and vortical in nature [2-5]. Useful insight into the fundamental cause of combustion instability can be gained by considering the generalized Rayleigh criterion as advanced by Chu [6]. Due to the inherent complexity of the physical phenomena involved in combustion instability, research in this field has relied heavily on experimental observations [3-5, 7, 8], simplified linear analysis [9], or, more recently, largeeddy simulation [10]. Shyy and Udaykumar [11] have developed a one-dimensional nonlinear model to simulate the flow and heat release in a constant area combustor. The heat release model used there has been extended from experimental observations of Langhorne [12] and is a modified version of the model advanced by Bloxidge et al. [9]. The model developed in Ref. [9] follows from a onedimensional linear stability analysis for acoustic waves in ducts, relating the fluctuations of heat release rate at the gutter to the velocity fluctuations through an empirical factor determined from experiments. The heat fluctuation is then assumed to travel downstream at the average gutter velocity and thus the fluctuations at any downstream location specified. In the present model, instead of specifying the unsteady heat release at every point in time and space, the heat release rate at

Combustion of Liquid Hydrocarbons in a High-speed Airstream

A theoretical formulation and experimental results of the high-speed mixing, ignition, and combustion characteristics of liquid fuels injected directly into a high-velocity airstream are presented. Experiments were performed in a direct connect constant area combustor supplied by a Mach 2 air stream at a static pressure level of 1 atmosphere and at total temperatures ranging from ambient (530°R) to approximately 3300°R. The fuel was liquid hexane and was injected laterally into the supersonic flow through piloted and nonpiloted injector configurations. It was found that autoignition and stable combustion occurs for air total temperatures above about 2600°R. Furthermore, the liquid fuel was effectively ignited and burned stably below this autoignition level by employing a small reacting pilot j et. Minimum pilot size and energy requirements were then established. It was found that the pilot mass flows needed for ignition and sustained combustion were small compared to the main fuel and air flow rates. Comparison with theory indicates combustion efficiencies up to 93% were obtained.

A study of free-jet and enclosed supersonic diffusion flames

An experimental study has been made of supersonic diffusion flames produced by the subsonic and supersonic free-jet injection of hydrogen into a high-enthalpy air-stream. The air-stream was flowing at a Mach number=1.98 and had a total temperature of approximately 1900°K and a static pressure of 14.7 psia. An axial, mid-stream mode of fuel injection was adopted. For the case of hydrogen injected subsonically, combustion was found to be complete (i.e. the concentration of unreacted hydrogen was negligible), at a distance of approximately 9 inches from the point of injection. For the supersonic injection of hydrogen this distance was increased by approximately 30%, for the range of fuel velocities used. The tests were repeated with methane as the injected fuel, but ignition did not occur even with the methane preheated to 480°K, or when a bluff-body was inserted into the flow to create shock conditions. The above flames were then enclosed in various combustors of simple geometry, either of constant-area, constant-divergence or some combination of these two. For the conditions specified above, combustion in the diffusional mode was found to be impossible in a constantarea combustor. A diffusion flame was initiated in a combustor which consisted of a short parallel section followed by a section with a divergence of approximately 1°. However no combustion took place within a completely divergent duct even though the divergence was less than 1°. An exponential relationship between pressure, area and length has been proposed, and a one-dimensional analytical treatment for the case of heat addition in a non-constant-area duct is included in this paper. This assumption is shown to be experimentally reasonable and to result in gas dynamic equations which include the effect of process length.