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51st AIAA/SAE/ASEE Joint Propulsion Conference (July 27-29, 2015.), Orlando, Florida, USA, 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2017-02-27, 資料番号: PA1710043000

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: PA1710005000

資料番号: ARDS06311000, レポート番号: ISTS2006-a-13

Flame propagation experiments on n-decane spray were performed in microgravity to investigate theflame propagation mechanism of a spray for less volatile fuels. After the spray was dispersed into an acrylic propagation tube with an inner diameter of 62 mm and a length of 535 mm, about 5 s were required for a quiescent spray to form under conditions of microgravity. The spray was ignited by an electrically heated nichrome wire just after the Sauter mean diameter (SMD) and the concentration of spray were measured by a laser droplet analyzer. The variation of flame propagation speed with the SMD was determined. Results showed that the change of the flame propagation speed with the SMD had a maximum for a constant equivalence ratio. This can be reasonably explained by analogy with the flame spread from a droplet to a neighboring droplet in a droplet array, that is, the flame spread of a droplet array and the flame propagation speed of spray reach a maximum when the spacing between a droplet and a neighboring droplet is approximately equal to the flame radius of the droplet. Therefore, all experimental data could be arranged on one curve when the lateral axis was taken as the spacing between droplets divided by the flame diameter. Even when methane of 1 vol %, which is outside the flammability limit, was added to spray with a constant overall equivalence ratio, the flame propagation speeds fell on the integrated curve.

For LOX/LH2 shear coaxial injectors, it is well-known that high-frequency combustion instabilities may occur when the injection temperature of hydrogen decreases below a certain value, but the mechanism of the initiation of combustion instability with a coaxial injector is still not clear. In the present study, firing tests were conducted with five types of single shear coaxial injector elements by using LOX and LH2 as propellants to further investigate the mechanism of the initiation of combustion instability during temperature-ramping changes during hydrogen injection. Results showed that unstable combustion was initiated when the hydrogen injection temperature decreased to less than a certain cryogenic temperature. The combustion instabilities observed in the present firing tests are discussed and classified into three different types.

A liquid jet disintegration phenomenon in the recessed region of a shear coaxial injector was examined to find the relation between the difference of the recess depth and flow instability which may induce combustion instability of liquid rocket engines. An injector with a rectangular cross section and a recess for a central post, which modeled the shear coaxial injector element employed in liquid rocket engines, was constructed. The injector was made of transparent acrylic glass to allow observation of the disintegration phenomenon in the recess. Cold-flow tests with water and nitrogen gas with ambient pressures of 0.2, 0.3 and 0.4 MPa were conducted. Results showed that a condition arose in which the flow in the recess was choked by two-phase flow. The choked flow was accompanied by vibration of the central post which caused a significant change of the disintegration pattern from moderate disintegration to violent disintegration. A similar transition from a fiber-type flow to a super-pulsating disintegration flow reported by Chigier and Reitz was also observed under non-choked conditions for a coaxial injector without a recessed region. The boundary of the transition was found to depend on certain values of ReL/(WeG) 0.5 for each recess depth, including the two-phase choked flow condition. This means that the transition from a fibertype flow to super-pulsating disintegration leads to the transition from a non-choked flow to a choked flow.

To study the mechanism of the initiation of combustion instability as hydrogen injection temperature decreases, a hydrogen temperature ramping test was conducted with a single coaxial injection element with LOX/LH2 at a chamber pressure of 8.0 MPa. Two types of injectors were used in the tests. One employed a straight bore LOX post and the other employed a taper-reamed LOX post for better atomization of LOX. The combustion flame was visualized with a high-speed video camera at a rate of 6,000 frames per second. Results showed that unstable combustion was initiated when the hydrogen injection temperature decreased to less than a certain cryogenic temperature. By observing the movement of the prominent pattern of OH emission on the flame, the flame was found to propagate downstream at a constant speed with the flame angle remaining constant during stable combustion. On the other hand, injection pressure peaks appeared during unstable combustion. In this case, a block of flame with strong OH emission was occasionally observed. A block of flame caught up with an anterior block and coalesced into a large block with strong OH emission. This coalesced block of the flame is herein termed “flame burst”.

Some rocket engines have a fuel mixer upstream of the injector to mix two hydrogen flows of different temperatures. In the mixing process, this fuel mixer may generate large fluctuations of flow properties, which in turn may increase combustion pressure fluctuations. Therefore, fuel mixers must be designed carefully to prevent such large fluctuations. In addition, fuel mixers must have good mixing characteristics and be free of large flow property fluctuations even at off-design points when rocket engines require deep-throttling capability. In this study, we experimentally and numerically investigated the effects of fuel mixer configuration in a rocket engine on the downstream flow properties. In the experiments, we used three different mixer models with different cryogenic hydrogen injection hole configurations (small holes, large holes, and a mixture of both size holes), and conducted experiments using cryogenic hydrogen and gaseous hydrogen under different flow conditions corresponding to engine throttling. The mixers with large injection holes showed better mixing characteristics than the mixer with smaller holes even under conditions of throttling.

Combustion flowfields in GH2/LOX sub-scale calorimeter chambers with multi-injector elements and full-scale thrust chamber are investigated using Reynolds-Averaged NavierStokes simulation, in which the finite rate chemistry with the H2/O2 detailed reaction mechanism is taken into account. The computed wall heat flux distributions are compared to that of the simplified cases to reduce a computational cost. The considered simplifications are a presence of reaction and a number of injector rows. At first, these simplifications are validated in the simulation of sub-scale chambers. The reaction is essential for the prediction of heat flux because it makes change the species distribution in a thermal boundary layer on a thrust chamber wall. A heat flux using a combustion simulation with only outermost injectors shows a good agreement with that with an original configuration near a face plate. On the other hand, it overestimates the heat flux around nozzle and throat parts. It is clarified that this overestimate comes from the shortage of unburned hydrogen near a chamber wall in the simplified method. Next, the simplification of the number of injector rows are applied to the simulation of full-scale thrust chambers. The effectiveness of this simplification for the prediction of wall heat flux is revealed. The optimal solution by using of the simplification is proven to be effective for the prediction of heat flux in a full-scale thrust chamber.