Antonella Ingenito

Using Aluminum for Space Propulsion

The combination of aluminum and water was theoretically analyzed to assess its performance potential for space propulsion, in particular for microrocket applications and whenever a compact package is desirable. Heat of reaction, impulse density, and handling safety are features making this combination interesting for chemical thrusters, especially because thrust is higher than typical of satellite electric thrusters. Ideal specific impulse I s p , thrust coefficient, adiabatic flame temperature, and combustion products were calculated for chamber pressures 1-10 atm, nozzle area ratios 25-100, and mixture ratios O/F 0.4-8.0. I s p reaches up to 3500 m/s. Also, the effect of hydrogen peroxide addition to aluminum and water on performance was explored. This combination improves performance slightly at the expense of simplicity, making it less attractive for microrocket engines. Ignition delay times were conservatively estimated assuming that aluminum was coated with its oxide and ignition occurred after the melting of the aluminum oxide. For this purpose heating and kinetics times were evaluated, the first by a one-dimensional physical model, the second by a reduced scheme. Results indicate that the heating time of a 0.1-μm-diameter aluminum particle may be of order 0.4 μs, whereas overall kinetics takes 10 μs: thus, the Al/water combination looks practical in principle for microrocket chambers characterized by short residence times.

Physics and Regimes of Supersonic Combustion

Understanding the physics of supersonic combustion is the key to design a performing engine for scramjet-powered vehicles. Despite studies on supersonic combustion dating back to the 1950s, there are still numerous uncertainties and misunderstandings on this topic. The following questions need to be answered: How does compressibility affect mixing, flame anchoring, and combustion efficiency? How long must a combustor be to ensure complete mixing and combustion while avoiding prohibitive performance losses? How can reacting turbulent and compressible flows be modeled? Experimental results in the past have shown that supersonic combustion of hydrogen and air is feasible and takes place in a reasonable distance, which is a necessary requirement in actual hypersonic vehicles powered by supersonic combustion ramjets. These results are explained based on a theoretical analysis of the physical mechanisms driving mixing and combustion in supersonic airstreams, where they are found to be different from those in the incompressible regime. In particular, the classic Kolmogorov scaling is shown to be no longer strictly valid, and the flame regime is predicted to be significantly affected by compressibility and different from that of subsonic flames. This analysis is also supported by the results of the numerical simulations presented, showing that by generating sufficiently intense turbulence, a supersonic combustion flame is short and can indeed anchor within a small distance from fuel injectors, with the flame typically burning in the so-called flamelets-in-eddies regime.

LES of the HyShot scramjet combustor

With the overall goal to clarify the physics of com pressible (supersonic) combustion, a 3D LES of the HyShot supersonic combustor has been performed and is reported in this paper. HyShot is an (originally) Australian p rogram to assess feasibility of supersonic combustion by means of a ballistic test flight. The HyShot combustion chamber is shaped as a box 75x9.8 mm in cross section and 300 mm long. Hydrogen is injected at 90 degrees with respect to the superson ic airstream 40 mm downstream from the combustor inlet by means of four 2 mm diameter choked orifices. Air enters the channel at a Mach number that, in the actual te st, depended on the flight trajectory; in this simulation, the trajectory poi nt is that at height = 28 km, where the Mach number was 2.79, P=82.11 kPa and T = 1229 K. A structured grid of about 14x10 6 nodes discretizes the actual combustor shape, wher e hydrogen-air combustion is treated by means of a detailed chemical kinetics model including 9 species and 37 reactions. Numerical results indicate that hydrogen penetrates in the air stream generating 3D bow shock structures upstream of the injection orifices as seen in experiments. In these regions recirculation zones u pstream and downstream of the fuel injection orifices are observed as expected; the OH predicted by LES indicates that a flame starts already in the upstream recircu lation zone. Interactions among the essentially 1D airstream entering the combustor, th e heat released and the 3D jets produce large vorticity rates and therefore enhance and accelerate turbulent mixing. Combustion is predicted very fast and efficient: on ly 0.5% of hydrogen is found unburned at the combustor exit.

The role of the baroclinic term in supersonic fuel/air mixing enhancement

Understanding of physics of supersonic combustion is mandatory for future hypersonic air-breathing propulsion systems. In fact, in order to overcome Mach 5 flight speeds, a supersonic combustion ramjet (SCRJ), where combustion takes place in supersonic conditions is required. In a scramjet, the air entering the combustor is supersonic: this means that in a residence time of about a millisecond air and fuel must mix and react. Hence, mixing in supersonic combustion plays a critical role on the the combustion efficiency and its understanding is critical to properly address a working engine. In this paper a rigorous analysis of the vorticity generation and transport in supersonic flows has been done in order to understand the key parameters to improve mixing and combustion. This 3D LES of the HyShot supersonic combustor performed by means of a fairly dense grid of 50Mnodes showed that the baroclinic term is the primary responsible of vorticity generation. In fact, interactions among the airstream entering the combustor and the H2 crossflow jet, the heat released and the shock waves produce a vorticity rate of order of 10 Hz. This vorticity generation is mainly due to the baroclinic term that creates spanwise vortices just upstream the H2 injection. These vortexes are afterwards tilted and stretched by the vortex stretching in the streamwise direction. LES predicts a very fast and efficient combustion: only 0.2% of hydrogen is found unburned at the combustor exit. Comparison of pressures distribution along the wall centerline at 1.32 ms shows a good agreement, mostly in the first part of combustor, where the grid is much more refined.

LES Modeling of Scramjet Combustion

The physics of supersonic combustion is analyzed in order to derive a new subgrid scale model for Large Eddy Simulation. Anisotropy associated to the directional Mach number typical of supersonic flows (i. e., M > 1 in only one direction) is explicitly considered by means of non-dimensionalized Navier-Stokes equations. The study shows that high Mach number flows experience mainly streamwise vorticity and consequently maximum helicity. Both affect mixing and may alter the turbulent kinetic energy decay by decreasing its rate, i.e., decreasing its characteristic spectral slope below that predicted by Kolmogorov scaling. Furthermore, it is analytically predicted that transversal pressure gradients increase vorticity, thus plausibly explaining the improved mixing realized by certain injectors. The supersonic regime is also found to affect the combustion regime: analysis of characteristic acoustic and convective times shows that while subsonic combustion takes place at approximately constant pressure, supersonic combustion takes place at approximately constant volume. Furthermore, collisional frequency is shown to increase due to local dilatation, resulting in faster kinetics and shorter ignition delay times. This effect could explain flame anchoring observed in some SCRJ combustor experiments. All the physical features and aspects of supersonic combustion found are used as ingredients to build a new subgrid scale model. In particular, micro-scale physics has been included by means of a subgrid kinetic energy equation that is algebraically modeled to provide the velocity fluctuation needed by the eddy viscosity SGS closure. Numerical simulations of a supersonic combustion NASA – Langley test case provide qualitative validation of the proposed model.

Shock/Boundary Layer/Heat Release Interaction in the HyShot II Scramjet Combustor

Previous computational work by these authors showed a very complex structure arising within the HyShot II combustor due to the interaction between the sonic crossflow injection and the airstream flowing at M = 2.78. In that work, a 3D Large Eddy Simulation (LES) of the HyShot II combustor was performed with a 14 ¤ 10 6 nodes structured grid. In the present work, in order to analyze in higher details the structures occurring within the HyShot II combustor, a fairly refined grid of 52 ¤ 10 6 nodes has been adopted. The LES simulations have been performed by means of a in-house code (S-HeaRT, Supersonic Heat Release and Turbulence): here, the LES model is based on a high order with lowdiffusion numerical schemes to accurately reproduce complex shock interactions, contact surfaces without artificially damping resolvable scales of turbulence. A hybrid method has been implemented to properly capture shocks while at the same time solving the transport equations away from discontinuities via a low dissipation, central scheme with fourth order accuracy. Simulations performed by means of the fairly accurated grid predicts very complex 3D flow structures arising within the flowfield due to the blockage induced by the H2 transverse injection. Ahead of each H2 injector a bow shock forms: the interaction of the bow shock and the boundary layer leads to a boundary layer separation zone where recirculation of H2 is allowed. In this recirculation zone, the presence of OH radical is predicted by LES indicating that a flame starts already in the injectors upstream the recirculation zone, downstream the flow separation. LES predicts the formation of barrel shock due to the H2 expansion: here the H2 jet expands untill it is definitively recompressed through the Mach disc. These flow structures, such us the bow and the barrel shocks, the Mach disk, the jet vortices and the horseshow vortices are in very good agreement with experimental results. Interactions among the airstream entering the combustor, the heat released and the shock waves produce a large vorticity rate that enhances and accelerates turbulent mixing. The vortex shedding, merging and tilting has also been analyzed, pointing out the contribution of the baroclinic term to the vortex generation and intensification. LES predicts a very fast and efficient combustion: only0.5% of hydrogen is found unburned at the combustor exit. Comparison of pressures distribution along the wall centerline at 1.2 ms shows a good agreement, mostly in the first part of combustor, where the grid is much more refined.

Mixing and Combustion in Supersonic Reactive Flows

Supersonic combustion is a key issue in any future plan to develop supersonic combustion (SC) ramjets (SCRJ) and rocket based combined cycles (RBCC) vehicles. Past experience has shown that among the major fundamental problems to be solved in SC are the need for rapid (i.e., < 1 ms) mixing and combustion efficiency. In this context numerical simulations, in particular LES, can help in improving the understandi ng of these issues. Current LES subgrid models developed for subsonic and adapted to supersonic combustion do not predict well or at all experimental results such as flame anchoring, whilst past experimental results with hydrogen injected at Mach 2.5 in Mach 2 airstreams showed combustion taking pl ace in about 2 ft. In fact, theoretical analysis shows that at high Mach number mixing and combustion are driven not only by transfer of kinetic energy by vortex stretching, as in subsonic reacting flows, but also by compressibili ty and baroclinic effects. Compressibility favours combustion by increasing reaction rates, as supersonic combustion occurs locally at about constant volume. Thus, when modelling mixing and combustion at small scales using LES, all these effects must be accounted before attempti ng to reproduce experimental results and to predict performance. To this purpose, a novel sub-grid scale (SGS) model (called henceforth ISC model, or ISCM for short) including these physical effects has been developed. This model has been validated so far by means of two experimental test cases. The first is the cross-flow injection of a sonic gaseous hydrogen jet in a Mach 2 airstream in the combustor built at the University of Tokyo; the second consists in 30° oblique injection of gaseous hydrogen in a Mach 2.5 airflow, an experiment performed in the supersonic combustion facility at NASA Langley Research Centre . For brevity, only the second test case validation is reported here. LES simulations using the well known Smagorinsky-Lilly SGS closure have been also performed for comparison. Results show that the ISCM is in better agreement with experimental data. In fact, while the Smagorinsky-Lilly model predicts neither combustion nor vortex structures, ISCM predicts flame anchoring, streamwise vorticity and temperatures close to those observed in the NASA-Langley experiments.

LES OF SUPERSONIC COMBUSTION OF H2/VITIATED AIR

The aim of this work is to numerically analyze supersonic combustion in a representative scramjet combustion geometry, focusing attention on turbulent mixing, flame stabilization and fuel conversion rate. The wind tunnel built at the University of Tokyo, consisting in a channel with a rectangular cross section and a step on the lower wall, followed by a divergent channel has been simulated. Hydrogen fuel is injected in cross-stream past the step and ahead of the divergent channel. Vitiated air enters the channel at Mach 2, P=49.7 kPa, T = 1145 K; hydrogen enters at Mac h 1, P=171 kPa, T=248 K. In order to show the effect of different subgrid scale models on results, 3-D unsteady LES simulations have been simulated by using the Smagorinsky and FM models. The Smagorinsky model is coupled with the EDC model to treat turbulence– combustion interaction. Hydrogen oxidation is treated by means of a one-step global reaction. Numerical results point out no substantial differences by using the two models, and show a different flame anchoring mechanism with respect to that observed in the experiments.

LAPCAT-II : conceptual design of a Mach 8 TBCC civil aircraft, enforced by full Navier-Stokes 3D nose-to -tail computation

Within the frame of the LAPCAT II project (Long-term Advanced Propulsion Concepts and Technologies, funded by the European Commission as part of the 7 th Frame Program and involving 16 European research labs and industries), four concurrent system studies have be en carried out to design a Mach 8 vehicle able to c arry passengers on the Brussels-Sydney route. The required range is above 18000 km, and the defined number of passengers is 300, which correspond to a 60 tons / 1400 m 3 payload.

Advance in Supersonic Combustion Modeling and Simulations

Mixing and combustion of supersonic reacting flows are currently under investigation for scramjet engines. Experimental results at NASA-Langley RC with single-jet hydrogen injection at Mach 2.5 in a Mach 2 airstreams showed combustion taking place in a very short lenght: this indicates that supersonic combustion may indeed be very fast. Numerical simulations of the same test case have also done using LES and the subgrid scale model, ISCM, developed purposely for supersonic combustion. This model accounts for the effects of compressibility on reaction rate and on mixing. Numerical simulations indicate that the flame is unsteady: it anchors at about 15 cm from the injector, develops downstream and lifts off. Periodical ignition and quenching have been investigated by means of frequency analysis.