Corin Segal

Subcritical to supercritical mixing

Liquid jet injection into a quiescent gaseous environment has been experimentally and analytically studied covering subcritical to supercritical conditions. The focus was placed on the influence that the surrounding gas pressure and temperature have on the jet breakup. Under subcritical conditions, the surrounding gas inertia and surface tension forces controlled this process with ligament formation from which the material broke off and the drops, later, formed. Decreasing surface tension influenced the jet surface behavior under transcritical conditions: Ligament formation was significantly reduced under these conditions with only occasional drop formation. Further increasing the pressure and temperature led to supercritical breakup modes. This manifested through a smoothening of the liquid-gas interface. Ligament formation was not observed under supercritical conditions; this indicated that surface tension did not play any role in the supercritical jet breakup. The experimental technique, using planar l...

COMBINED ROCKET AND AIRBREATHING PROPULSION SYSTEMS FOR SPACE-LAUNCH APPLICATIONS

A review of rocket-airbreathing combined-cycle propulsion systems for Earth-to-orbit applications is presented. Rocket-based combined-cycle (RBCC) engines take advantage of the synergistic interactions between the rocket and the airbreathing elements of the engine and the use of high-specie c impulse cycles to yield a mission-averaged specie c impulse that is higher than all-rocket technology can provide. An overview of the multimode operation is given, along with a review of both experimental and modeling work that has been done on this class of engines. Selected issues involved with these engines are discussed. These include engine/vehicle integration, e ow-path design for multimode operation, fuel selection, mixing enhancement and afterburning in rocket-ejector mode, thermal choking, and e ameholding. RBCC propulsion is becoming recognized as a promising technology for achieving a signie cant reduction in the cost of delivering payload to orbit.

Experimental Study of Fluid Jet Mixing at Supercritical Conditions

Subcritical and supercritical fluids were injected in an inert gaseous atmosphere. Density distribution was measured and density-gradient profiles were inferred from the experimental data. A novel method was applied for the detection of detailed structures throughout the entire jet center plane. The core lengths were measured for each of the cases and correlated with previous visualization results. An eigenvalue approach was taken to determine the location of maximum gradients. The results show a significant influence of chamber-to-injectant density ratio on the core length in the supercritical domain, unlike the subcritical conditions.

Penetration and Spreading of Liquid Jets in an External -Internal Compression Inlet

Injection of liquid fuel in the inlet of a vehicle e ying at hypersonic speed is related to the development of liquid fuel-based supersonic-combustion ramjets as a means to increasetheresidence timeand achieve partial fuel premixing priorto arrivalat thecombustion chamber. The strong liquid interactionwith theinlethigh-momentum aire owand shock wavesystem offersa mechanism forrapid jetand droplet breakup and, hence, improved mixing. The penetration and spreading of liquid jets in a two-dimensional external ‐internal compression inlet at Mach 3.5 using a noncombustible mixture that simulated the viscosity and surface tension of JP-10 was evaluated. Schlieren imaging has been used as the visualization technique for penetration studies, as well as light scattering for jet spreading. By using thin pylons to create a low-pressure region at the liquid injection station, the penetration increased compared with a nonpylon cone guration while reducing the pressure losses associated with transverse jet injection. The pylons contributed to lift the liquid from the injection surface with less lateral spreading than the nonpylon-injection case, thus avoiding the presence of a low-speed combustible mixture in the inlet/isolator boundary layers and providing a mechanism to eliminate potential e ashback.

Effects of Mixing Schemes on Kerosene Combustion in a Supersonic Airstream

A study of kerosene combustion in a supersonic vitiated aire ow at Mach 4.75 e ight enthalpy was conducted in direct-connect tests at Mach 1.8 at a stagnation temperature of 1000 K. The effects of shockand vortex-enhanced mixing mechanisms on the combustion efe ciency were evaluated. Also included in this study were the effects of fuel heating and jet penetration. The experimental conditions corresponded to the low end of the hypersonic e ight regime. The following geometric cone gurations were employed: 1) a generic, rearward-facing step, 2 ) a modie ed rearward-facing step with beveled edges to facilitate vortex-enhanced mixing, and 3 ) a rearward-facing wedge (15 or 30 deg) placed downstream of the rearward-facing step to induce shock-enhanced mixing. In all cone gurations, a gaseous hydrogen ‐ pilot jet was injected parallel to the main e ow from the base of the rearward-facing step and the liquid kerosene was injected normal to the main e ow at three or e ve step heights downstream of the step (the step height was 10 mm). Stable kerosene combustion was obtained for a maximum injected kerosene equivalence ratio of 0.86. For efe ciency evaluation, the pilot ‐ hydrogen equivalence ratio was selected between 0.02 ‐ 0.04, while the kerosene equivalence ratio was maintained at 0.325. In all experiments, locally rich stratie ed kerosene combustion took place in a layer close to the injection wall. The wedge e ameholder contributed to an increased kerosene combustion efe ciency by the generation of shock ‐ jet interactions. The beveled-edge step improved far-e eld mixing, thereby reducing the local kerosene equivalence ratio, resulting in the highest kerosene combustion efe ciency among all cone gurations tested. Fuel heating below levels required for e ash vaporization (one-third of the e ash vaporization energy, in this case ) did not contribute to increased combustion efe ciency. On the contrary, this level of heating reduced the fuel density with adverse effects on penetration and mixing.

Concentration Distribution in a Supersonic Flow Recirculation Region

Flameholding in supersonic flow depends on the local conditions in the recirculation region and the mass transfer into and out of this region. Large gradients in local gas composition and temperature exist in the recirculation region; hence, stability parameter correlations developed for premixed flames cannot be used to determine the blowout stability limits for nonpremixed flames encountered in practical devices. In the present investigation, mixture samples were extracted at different locations in the recirculation region and shear layer formed behind a rearward-facing step and were analyzed by a mass spectrometer to determine the distribution of species concentration in the region. Both nonreacting flow tests and combustion experiments were performed for a range of fuel-related parameters such as injection location, injection pressure, and fuel type. The difference between the local fuel mole fraction within the recirculation region determined from mass spectrometry and the global fuel mole fraction based on the total moles of air and fuel injected was identified. Planar-laser-induced fluorescence was used in nonreacting cases to provide a two-dimensional image of fuel distribution and complement the mass spectrometry measurements. Differences between local concentrations and estimates based on overall global injected mass flows were large, amounting to an order of magnitude. This implies significant differences in flame stability limits of a nonpremixed flame in supersonic flow compared to premixed flame.

Mixing and Chemical Kinetics Interactions in a Mach 2 Reacting Flow

An experimental study of transverse hydrogen injection and combustion behind a rearward-facing step into a Mach 2 airflow was conducted in an electrically heated (not vitiated), continuous-flow facility to evaluate the effects of initial conditions (temperature, pressure, and equivalence ratio), and analyze the interactions between mixing and combustion in supersonic, reacting flows. Neither mixing nor reaction rates dominate in this particular regime, thus the use of initial conditions to scale fuel mixing (i.e., dynamic pressure ratio) has to be modified by a descriptor that includes the effects of combustion on the flow conditions at the fuel injection station. Combustor inlet static pressure was varied from 0.25 to 0.5 atm, and total temperature from 300 to 850 K. Injector configurations include both single and staged injection with injectors of 1 and 1.5 mm diam, transverse to the airflow, behind a 5-mm rearward-facing step. Images of visible flame emission distribution at several temperatures correlated with pressure and temperature measurements are used to describe the coupling between fluid dynamics and chemical kinetics, discussed in terms of a characteristic global Damkohler number (ratio of chemical reaction rate to turbulent mixing rate). A proposed modification to mixing scaling with dynamic pressure ratio in the presence of heat release effects is presented.

Ignition Delay for Jet Propellant 10/Air and Jet Propellant 10/High-Energy Density Fuel/Air Mixtures

Jet propellant 10 [(JP-10) or exotetrahydrodicyclopentadiene] is one of the leading candidate fuels for use in pulse detonation engine applications. As such, its ignition delay characteristics have been studied previouslyin very dilute mixtures at pressures from 1 to 9 atm and temperatures from 1300 to 1670 K. The ignition delay times are studied of JP-10/air and JP-10 blended with methylated PCU alkene dimer, nitronorbornane, dinitronorbornane, and ethylhexyl nitrate in air at pressures from 10 to 25 atm and temperatures from 1200 to 2500 K using a shock tube. Ignition delays were primarily measured using CH emission and secondarily using OH emission. Ignition delays were essentially insensitive to all of the additives tested. Additionally, ignition delays for dicyclopentadiene (a suspected intermediate in the combustion of JP-10) were also tested. Above 1500 K, multiple peaks in the CH emission were found. Further tests using OH emission indicate that the main peak in CH emission at these higher temperatures is probably due to reactions involved in the approach to equilibrium and give spuriously long ignition delay times.

A Review of Fuel Pre-injection in Supersonic, Chemically Reacting Flows

Developing an efficient, supersonic combustion-based, air breathing propulsion cycle operating above Mach 3.5, especially when conventional hydrocarbon fuels are sought and particularly when liquid fuels are preferred to increase density, requires mostly effective mechanisms to improve mixing efficiency. One way to extend the time available for mixing is to inject part of the fuel upstream of the vehicle’s combustion chamber. Injection from the wall remains one of the most challenging problems in supersonic aerodynamics, including the requirement to minimize impulse losses, improve fuel-air mixing, reduce inlet∕combustor interactions, and promote flame stability. This article presents a review of studies involving liquid and, in selected cases, gaseous fuel injected in supersonic inlets or in combustor’s insulators. In all these studies, the fuel was injected from a wall in a wake of thin swept pylons at low dynamic pressure ratios (qjet∕qair=0.6–1.5), including individual pylon∕injector geometries and combinations in the inlet and combustor’s isolator, a variety of injection conditions, different injectants, and evaluated their effects on fuel plume spray, impulse losses, and mixing efficiency. This review article cites 47 references.

Heat Flux Measurements for a GO2/GH2 Single-Element, Shear Injector

A high-pressure facility was used to investigate GH2/GO2 single-element, shear, co-axial injector with operating conditions typical of rocket engines. Oxygen-tohydrogen mass flow ratios of 4 and 6 at operational chamber pressures of 6.2, 4.9, 4.5, and 2.8 MPa were investigated by (a) keeping the propellant mass flow rates constant while changing the exhaust nozzle diameter and, (b) by keeping the exhaust nozzle diameter constant and changing the propellant mass flow rates to change the chamber pressure. Axial heat fluxes, injector face temperature, exit nozzle temperature and high-frequency pressure were measured. The injector’s outer diameter was 2.7 mm and the chamber cross-section was square with a side L = 2.5 cm. Maximum heat release occurred at 2.4L from the injector face. The injector face temperatures showed little to no dependence on chamber pressure. The profiles of heat flux and chamber wall temperatures indicated no pressure dependence and only a slight dependence on propellant injection velocities when the massflows were kept constant. A scaling of heat flux values based on fuel mass flow rate, instead of chamber pressure, is, therefore, suggested. High-frequency pressure measurements indicated that most of the energy is in the first longitudinal mode. This, along with the lack of pressure dependence and only slight dependence on the propellant injection velocities suggested that the basic dynamic structures of the combusting flow were mainly dominated by the chamber geometry.