Performance of a Dual-Mode Combustor with Multistaged Fuel Injection

Kobayashi, Kan; Tomioka, Sadatake; Kato, Kanenori; Murakami, Atsuo; Kudo, Kenji

doi:10.2514/1.19294

Cited by 49 publications

(20 citation statements)

References 13 publications

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“…The global ignition can only be achieved when strong luminescence occurred around Cavity 3 and maintained at a stable state (as shown in Fig. 5(c) and (d)), accompanying with much more heat release and pressure increase in downstream [10]. Then shock wave will be pushed upstream and flame front also propagates upstream to fuel injection, as shown in Fig.…”

Section: Ignition and Flame Propagationmentioning

confidence: 99%

Plasma-assisted ignition for a kerosene fueled scramjet at Mach 1.8

Tong

et al. 2013

Aerospace Science and Technology

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By using a plasma jet (PJ) torch with 1.5 kW input power as an igniter, successful ignition for liquidkerosene fueled combustion experiment was conducted in a direct-connected supersonic test facility. The incoming flow has total temperature of 950 K and local Mach number of 1.8, corresponding to Mach 4 flight condition. In this study, several optical techniques, including high speed photography, high speed schlieren photography, and planar laser scattering (PLS) technique, were combined to study the ignition process, flame propagation, and mixing features of liquid kerosene fuel with air around the cavity. The effect of fuel injection position, injection pressure, and feedstock gas on ignition performance has been analyzed. The results indicate that local mixing is a critical factor for ignition. It is also shown that the PJ torch with N 2 + H 2 feedstock is superior to the PJ torch with N 2 feedstock for the ignition of liquidkerosene fuel. These results are valuable for the future optimization of kerosene-fueled scramjet engine when using a PJ torch as an igniter.

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Section: Ignition and Flame Propagationmentioning

confidence: 99%

Plasma-assisted ignition for a kerosene fueled scramjet at Mach 1.8

Tong

et al. 2013

Aerospace Science and Technology

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“…This combustion mode is typical of engines designed for flight Mach numbers less than 6 (Billig et al, 1990;Heiser and Pratt, 1994;Kobayashi et al, 2006;Waltrup and Billig, 1973). Turner and Smart (2013) recently documented the occurance of dual-mode combustion in a rest scramjet designed for flight at Mach 8.…”

Section: Fuel-off Datamentioning

confidence: 99%

“…The first includes those studies that examine fundamental flow or aerothermochemical phenomena at a component level. Examples include boundary layer transition on cones and flat plates (Mee, 2002;Simeonides, 2003), inlet design studies (Gollan and Jacobs, 2013;Tan et al, 2011), fuel injector studies examining mixing performance (Ben-Yakar et al, 2006;Kawai and Lele, 2010), direct connect studies of combustion efficiency (Kobayashi et al, 2006;Tomioka et al, 2006), studies of shock-boundary layer or shock-shock interactions (Burtschell and Zeitoun, 2003;Dann and Morgan, 2011) and studies of single-expansion ramp nozzles (Hirschen et al, 2009;Tanimizu et al, 2011). The second broad category includes numerical studies that examine the design, optimisation and ascent trajectories of scramjet powered vehicles at a global system level (see for e.g.…”

Section: Research Contextmentioning

confidence: 99%

An experimental investigation of an airframe integrated three-dimensional scramjet engine at a Mach 10 flight condition

Doherty¹

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The University of Queensland in 2014 school of mechanical and mining engineering centre for hypersonics abstract Realisation of the dream of airbreathing access-to-space requires the development of a scramjet engine that produces sufficient net thrust to enable acceleration over a wide Mach number range. With engines that are highly integrated with the airframe, the net performance of a scramjet powered vehicle is closely coupled with the vehicle attitude and is difficult to determine only from component level studies. This work investigates the influence of airframe integration on the performance of an airframe integrated scramjet through the measurement of internal pressure distribution and the direct measurement of the net lift, thrust and pitching moment using a three-component stress wave force balance. The engine chosen as the basis for this study was the Mach 12 rectangularto-elliptical shape-transition (m12rest) scramjet that was developed by Suraweera and Smart (2009) as a research engine for access-to-space applications. The inlet and combustor flowpath were integrated with a slender 6°wedge forebody, streamlined external geometry and three dimensional thrust nozzle. The scale of the engine was chosen so that the entire engine would fit within the core-flow diamond (bi-conic) produced by a Mach 10 facility nozzle. The Mach 10b facility nozzle was chosen because it is the largest nozzle current in use with the t4 Stalker Tube and because the offdesign performance of a scramjet engine is of interest for access-to-space vehicles that must accelerate over a range of Mach numbers.Freejet experiments were conducted within the t4 Stalker Tube. Two trueflight Mach 10 test conditions were used: a high pressure test condition that replicated flight at a dynamic pressure of 48 kPa and a low pressure test condition that replicated flight at a dynamic pressure of 28 kPa. Scaling of the test conditions according to the established binary scaling law was not completed due to facility operational limits.The engine featured two fuel injection stations from which gaseous hydrogen was injected. The first injection station was partway along the length of the inlet while the second injection station was at the start of the combustor behind a rearward facing circumferential step. In addition to investigating inlet-only and step-only injection, a combined scheme where 68 % of the fuel was injected from the step station and 32 % from the inlet station was also investigated.To support the analysis of the experimental results, numerical simulations of the engine with no fuel injection were conducted using the nasa code vulcan. Analysis of the simulations show that the mass capture ratio with respect to the projected inlet area is approximately 60 % at each test condition. The simulations also show that spillage of flow from the slender forebody accounts for just 12 % of the flow through the projected iii inlet area, a small but non-negligible fraction. By integrating the engine surface forces, the drag coefficient with respect ...

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“…[7][8][9] When a scramjet engine is operated in ramjet mode, supersonic flow is decelerated to subsonic speed through a pseudo shock wave within the isolator, and the subsonic flow is re-accelerated to supersonic speed due to heat release by combustion. Although this conventional ramjet-mode operation assumes fuel injection within the constant-area combustor, Kanda et al 7) and Kato et al 8) attempted fuel injection at the exit of the diverging combustor to form the choking point there in their scramjet component tests.…”

Section: Ramjet-mode Operation With Downstream Combustion In a Scramjmentioning

confidence: 99%

Suppression of Combustor-Inlet Interaction in a Scramjet Engine under Mach 4 Flight Conditions

Kobayashi

Kanda

Tomioka

et al. 2007

Trans. Japan Soc. Aero. S Sci.

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A sidewall-compression-type scramjet engine was tested under Mach 4 flight conditions. The tested engine had an inlet, a constant cross-sectional area isolator, a constant cross-sectional area combustor, a diverging combustor, and an internal nozzle. In a previous study under the same flight conditions, the maximum thrust increment using fuel injection within the constant-area combustor was 1,380 N at an equivalence ratio of 0.31, and further fuel injection resulted in combustor-inlet interaction (designated as CII). To suppress the CII in the present study, we attempted (1) two-stage fuel injection within the constant-area combustor and the diverging combustor and (2) a boundary layer bleed on the top wall. The former was to suppress heat release around the first-stage fuel injectors in the constant-area combustor, and the latter was to decrease interaction length by decreasing the boundary layer thickness on the top wall. In the case of two-stage fuel injection, the maximum thrust increment was 2,230 N at an equivalence ratio of 0.63. In the case of the boundary layer bleed, on the other hand, the maximum thrust increment was 2,300 N at an equivalence ratio of 0.66. Thus, twostage fuel injection and boundary layer bleed led to 62% and 67% higher maximum thrust increments than that obtained in the previous study, respectively. Finally, both the two-stage fuel injection and boundary layer bleed were applied simultaneously to obtain the best thrust performance, and the maximum thrust increment was 2,560 N at an equivalence ratio of 0.95. As a result, we obtained an 86% higher maximum thrust increment than that in the previous study. 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%.

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Performance of a Dual-Mode Combustor with Multistaged Fuel Injection

Cited by 49 publications

References 13 publications

Plasma-assisted ignition for a kerosene fueled scramjet at Mach 1.8

Plasma-assisted ignition for a kerosene fueled scramjet at Mach 1.8

An experimental investigation of an airframe integrated three-dimensional scramjet engine at a Mach 10 flight condition

Suppression of Combustor-Inlet Interaction in a Scramjet Engine under Mach 4 Flight Conditions

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