A new criterion predicting the onset of scramjet unstart is proposed. The source of unstart is hypothesized to stem from the turbulent compressible flame choking altering the optimal thermal choking design location or geometry of a scramjet combustor for heat release. This criterion is useful for assessing scramjet unstart and engine tuning for optimal heat release for ideal expansion and engine performance.

A complementary experimental and computational study was performed to characterize the test section flowfield of a variable-Mach-number supersonic wind tunnel. A planar laser Rayleigh scattering technique using [Formula: see text] nanocrystals was used to capture cross-sectional flow visualizations at various test section streamwise locations. Reynolds-averaged Navier–Stokes simulations of the wind tunnel flow were performed using a [Formula: see text] shear stress transport turbulence model to compare against the flow visualizations. Mach 3.20 and 3.55 flow conditions were investigated in the wind tunnel with an empty test section. The Rayleigh signal profile near the wind tunnel walls was compared against boundary-layer thickness data from shadowgraph visualizations, pitot survey measurements, and numerical simulation predictions to guide the interpretation of the scattering results. A brief discussion regarding the nature and potential source of certain unexpected scattering patterns in the images was provided. The most distinguishable flowfield characteristics observed in the Rayleigh scattering visualizations were the asymmetric thickness of the boundary layers along the test section walls and the appearance of two floor lobes near the flow-path corners. The results of the numerical simulations showed favorable agreement with the experimental data and were used to provide insight into the mechanism that caused the boundary-layer features observed. This study demonstrated the usefulness of [Formula: see text] Rayleigh scattering to obtain flow visualization data of great value to better understand flow quality in supersonic wind tunnels with two-dimensional nozzles.

Space travel requires high-powered, efficient rocket propulsion systems for controllable launch vehicles and safe planetary entry. Interplanetary travel will rely on energy-dense propellants to produce thrust via combustion as the heat generation process to convert chemical to thermal energy. In propulsion devices, combustion can occur through deflagration or detonation, each having vastly different characteristics. Deflagration is subsonic burning at effectively constant pressure and is the main means of thermal energy generation in modern rockets. Alternatively, detonation is a supersonic combustion-driven shock offering several advantages. Detonations entail compact heat release zones at elevated local pressure and temperature. Specifically, rotating detonation rocket engines (RDREs) use detonation as the primary means of energy conversion, producing more useful available work compared to equivalent deflagration-based devices; detonation-based combustion is poised to radically improve rocket performance compared to today’s constant pressure engines, producing up to 10 % increased thrust. This new propulsion cycle will also reduce thruster size and/or weight, lower injection pressures, and are less susceptible to engine-damaging acoustic instabilities. Here we present a collective effort to benchmark performance and standardize operability of rotating detonation rocket engines to develop the RDRE technology readiness level towards a flight demonstration. Key detonation physics unique to RDREs, driving consistency and control of chamber dynamics across the engine operating envelope, are identified and addressed to drive down the variability and stochasticity observed in previous studies. This effort demonstrates an RDRE operating consistently across multiple facilities, validating this technology’s performance as the foundation of RDRE architecture for future aerospace applications.

Premixed cavity stabilized flames are experimentally investigated in a high-speed ramjet engine under the influence of varying mean pressure gradients. The ramjet cavity incorporates a backward facing step with an aft ramp for flame stabilization in high Reynolds number. The ramjet engine is subject to varying wall geometry to form converging, diverging, and nominal configurations in order to investigate the effects of mean pressure gradients on engine performance. High-speed particle image velocimetry (PIV) and chemiluminescence imaging diagnostics are simultaneously employed to capture the reacting flow fields and flame dynamics. Imposing a larger favorable pressure gradient is shown to shrink the recirculation zone and alter shear layer dynamics, which leads to increased drag on the cavity. Additionally, inducing a stronger favorable pressure gradient is shown to excite a shear layer instability mode, characterized by a Strouhal number of St=0.1. Proper orthogonal decomposition (POD) results reveal that the instability mode is comprised of large-scale oscillations that occupy the entire cavity flow region, indicating that the excited oscillations are the manifestation of a global vortex shedding instability that occurs under non-reacting conditions. The results demonstrate that the performance and stability of the ramjet cavity flame can be influenced by the mean pressure gradient, which is vital for the design of high-speed air-breathing propulsion systems such as dual-mode scramjets.

The combustion dynamics of a mixture of ethylene and air in a cavity flameholder are examined using high-resolution numerical simulations. The simulations are performed using both laminar and turbulent inflow boundary conditions to characterize the effects freestream turbulence on the flame-vortex interactions within the combustor. Experimental test results are utilized to validate the simulation's ability to capture the relevant flow and flame behavior. The turbulence intensity is shown to influence the structure of the flow field by creating a more dominant primary vortex within the combustor and decreasing the shear layer reattachment length. The integral length scales and energy spectra are calculated throughout the domain and show that the flow is driven by smaller turbulent eddies within the ramp region of the combustor. The flame-vortex interaction is further examined and revealed to cause an instability forming in the combustor under both laminar and turbulent conditions. The instability is confirmed by examining the pressure fluctuations within the combustor and it is demonstrated that the time scale of the pressure fluctuations corresponds to the frequency of the vortex shedding from the ramp wall.

A complementary experimental and computational investigation of the flow characteristics in a supersonic wind tunnel are presented. Stereoscopic planar laser Rayleigh scattering of CO2 particles, typically several nanometers to tens of nanometers in size, were used in this study allowing cross-sectional visualization of the supersonic flow and turbulent boundary layers throughout the test section of an academic scale facility. The optical diagnostic measurements show asymmetric boundary layer growth between the top and bottom walls of the test section as well as a disturbed bottom boundary layer due to the induced pressure and velocity gradient. The computational fluid dynamics simulations were used to provide additional detail and insight into the mechanisms driving these features.

Precipitation and suspended droplets in the atmosphere present a significant erosion risk to vehicles traveling and high-supersonic and hypersonic speeds. Here we lay the foundation for simulations of such phenomena by exploring and comparing common models for droplet drag and breakup. We test these models in a simulation of a sharp 10-degree wedge traveling at Mach 3. Understanding and properly modeling the interplay between the drag and breakup models will be critical to characterizing potential impact events by the size, quantity, location, and incidence angle of the droplets. We find the KHRT breakup model used in conjunction with an empirical drag model that accounts for the effects of droplet deformation to provide physically-reasonable droplet behavior; however, additional experimental data will be required to fully evaluate this model and to tailor model parameters for high-Mach number, high-Weber-number flows.

Future terrestrial and interplanetary travel will require high-speed flight and reentry in planetary atmospheres by way of robust, controllable means. This, in large part, hinges on having reliable propulsion systems for hypersonic and supersonic flight. Given the availability of fuels as propellants, we likely will rely on some form of chemical or nuclear propulsion, which means using various forms of exothermic reactions and therefore combustion waves. Such waves may be deflagrations, which are subsonic reaction waves, or detonations, which are ultrahighspeed supersonic reaction waves. Detonations are an extremely efficient, highly energetic mode of reaction generally associated with intense blast explosions and supernovas. Detonation-based propulsion systems are now of considerable interest because of their potential use for greater propulsion power compared to deflagration-based systems. An understanding of the ignition, propagation, and stability of detonation waves is critical to harnessing their propulsive potential and depends on our ability to study them in a laboratory setting. Here we present a unique experimental configuration, a hypersonic high-enthalpy reaction facility that produces a detonation that is fixed in space, which is crucial for controlling and harnessing the reaction power. A standing oblique detonation wave, stabilized on a ramp, is created in a hypersonic flow of hydrogen and air. Flow diagnostics, such as high-speed shadowgraph and chemiluminescence imaging, show detonation initiation and stabilization and are corroborated through comparison to simulations. This breakthrough in experimental analysis allows for a possible pathway to develop and integrate ultra-high-speed detonation technology enabling hypersonic propulsion and advanced power systems.

Rotating Detonation Rocket Engines (RDREs) have the potential to increase performance and lower the cost of launch vehicles by harnessing the benefits of pressure gain combustion. Thermal management and modelling in RDREs is challenging due to the high heat transfer rates and complex heat transfer process. In this study, we made quantitative heat measurements in a laboratory RDRE operating at elevated chamber pressures (CTAP up to 173 psia) with GOX-GH2 and GOX-GCH4 propellants. We found the highest heat flux near the propellant injectors (up to 25 MW/m2). The heat flux near the injectors (detonation cell region) does not follow the mass flux scaling of the traditional Bartz equation. The heat flux in the downstream combustor region does follow the mass flux scaling of the traditional Bartz equation for heat flux in constant pressure combustion rocket engines.

This research quantifies the evolution of pressure for fast burning regimes characterized by various degrees of compressibility and involving turbulent flames and shocks. The experimental exploration is conducted in a Turbulent Shock Tube facility, where the level of flame compressibility is controlled by varying the equivalence ratio of the hydrogen-air mixture. High-speed particle image velocimetry, chemiluminescence, schlieren, and pressure measurements are simultaneously acquired to capture the rise in stagnation pressure for various regimes from fast flames to shock-flame complexes. The pressure and velocity measurements are used to analyze combustion regimes on the Rankine-Hugoniot diagram that shows the flame-driven compression for a range of fast flame conditions evolving toward detonation onset. Various levels of compression are dependent on the level of shock-flame coupling and flame velocities. Lower degrees of compressibility show 52% efficiency of an ideal ZND cycle with 40% thermal efficiency, while shock-flame complexes are shown to produce 81% of the work produced by an ideal ZND cycle with 53% thermal efficiency.

Experimental evidence of controlled detonation initiation and propagation in a hypersonic flow of premixed hydrogen-air is presented. This controlled detonation initiation is created in a hypersonic facility capable of producing a Mach 5 flow of hydrogen-air. Flow diagnostics such as high-speed schlieren and OH* chemiluminescence results show that a flame deflagration-to-detonation transition occurs as a combined result of turbulent flame acceleration and shock-focusing. The experimental results define three new distinct regimes in a Mach 5 premixed flow: deflagration-to-detonation transition (DDT), unsteady compressible turbulent flames, and shock-induced combustion. A two-dimensional implicit-LES (ILES) simulation, which solves the compressible, reactive Navier-Stokes equations on an adapting grid is conducted to provide additional insight into the local physical mechanism of detonation transition and propagation.

The proposed work seeks to experimentally validate a modified Paschen law which accounts for the effects of electron-ion pair removal between two electrodes within a dynamic gas medium. A test facility is designed to produce Mach 1.5 and Mach 2 flow with pressure ranges of [309 – 760] Torr and [169 – 760] Torr, respectively. Custom designed aluminum electrodes are mounted into the test section at desired gap distances. A high voltage power supply is utilized to charge the electrodes up to 60 kV until discharge occurs. Ignition cables are used to safely transmit the high voltage from the power supply to the electrodes. The discharge voltage of the electrode is recorded over the respective range of pressures within the test section. An operational pressure range is calculated using isentropic and normal shock relations at the desired Mach numbers. Schlieren imaging is used to capture the electrostatic discharge that occurs within the supersonic flow. The measured pressure and voltage values are plotted against the modified Paschen curve as a function of Mach number and electrode gap distance as a means of validation. A linear regression approach is utilized to statistically compare the experimental data to the modified Paschen law.

Rotating detonation engines (RDEs) have been shown to be viable pressure-gain combustion devices. A variety of reactants have been tested, of which the most prominent is the non-premixed system that employs hydrogen-air mixtures due to its high detonability. Experimentally, some researchers \cite{Bykovskii_2006,Bykovskii_2012} have also evaluated two-phase mixtures such as liquid or solid particulates to produce self-sustained rotating detonations, but there is lack of numerical studies to support such findings. In this study, a Eulerian-Eulerian (EE) dense particle formulation is used to model the dispersed phase in detonation type flows. Reduced kinetics is implemented for both gas-phase mixtures and for coal particle combustion. Studies include detonation in stoichiometric and non-stoichiometric hydrogen-air mixtures, and heterogeneous detonations in solid reactive particles of aluminum and/or carbon. Comparison between the computed results and available data show reasonable agreement. Finally, a non-premixed detonation channel consisting of an unwrapped array of eight injectors as in the RDE is simulated using the EE approach to demonstrate its capability. The effects of mixing, particle volume fraction and size is examined to understand the behavior of the detonation coupling with reacting particles and the associated heat release.

Velocimetry measurements have been established inside of a 6.06” (~154 mm) radially fed Rotating Detonation Engine. The engine has a channel gap width of 0.3” (7.62 mm) making the inner diameter 5.46” (~139 mm), and an air gap height of 0.022” (0.56 mm). The fuel plenum consists of 80 discrete injection points that are 0.035” (~.89 mm) in diameter and impinge axially on the radially fed air as to assist in mixing. The oxidizer was air, and the fuel was hydrogen combusted at stoichiometric conditions with a total mass flux of 200 kg/s*m2. Particle Image Velocimetry was utilized to capture the flow field inside of the engine. The area of interest was 1” by 1” with a spatial resolution of 12 μm/px, and an interframe time of 1 μs. It was found that directly behind the transverse detonation wave (within the same axial region) velocity dropped to nearly half of the total wave speed, approximately 900 m/s. However, just above the detonation the velocity magnitudes were on the order of the transverse wave speeds, approximately 1600 m/s. These results were validated against Computational Fluid Dynamics provided by Georgia Tech using the exact same geometry.

This paper discusses the investigation of shock induced reactions in a high enthalpy Mach 5 flows. The current Mach 5 high-enthalpy combustor provides a supersonic quasi-premixed stream targeted to match the CJ detonation conditions for a propagating detonation. A hydrogen diffusion pre-burner is operated to provide total temperatures T0 ~ 1150K. High speed Schlieren and OH chemiluminescence are utilized to outline the shock induced combustion events driven by the temperature jump across the oblique shock pair formed at the nozzle exit plane without the use of an external ignition source. The addition of a ramp in the test section drives intermittent Mach stem induced combustion (MSIC).

The Deflagration-to-Detonation Transition is a complex process which involves shock reflections, deflagrations, boundary layers and their interactions with each other. Within these transient regimes of the deflagration-to-detonation process, evidence of pressure rise can be observed. This pressure rise is a result of shock-flame behavior generated by the acceleration of a fast flame. In this study, the pressure dynamics are characterized in these transient regimes as the flow evolves towards detonation. A semi-confined Turbulent Shock Tube facility with multiple configurations and optical access is used to observe the propagating shock-flame behavior. High-resolution pressure transducers are used to resolve pressure profiles which are characterized further on the Rankine-Hugoniot diagram.

It has been argued that the most suitable length scale to characterize the hydrodynamic thickness of a turbulent detonation wave is the location of the sonic surface. The sonic surface separates the standing reaction zone from the unsteady expansions behind the detonation wave. This study aims to determine the sonic surface location of a detonation structure which will be used to quantify the hydrodynamic thickness. Results show that the hydrodynamic thickness is about 6.6 the cell size which agrees with previous computational results. High-speed diagnostics such as particle image velocimetry, and tunable diode laser absorption spectroscopy are used to characterize and resolve the complex flow field behavior.

The paper presents experimental evidence of continuous detonation in a rotating detonation rocket engine (RDRE) powered by H2/O2 propellants. High-speed chemiluminescence imaging is used to characterize the detonation wave dynamics by introducing a tracer in the hydrogen fuel flow. The results show continuous five-wave co-rotating detonations at various equivalence ratios and flow rates demonstrating the potential for H2/O2 propellant based RDREs for upper-stage rocket engines.

The present work explores the mechanisms behind unconfined deflagration-to-detonation transition (DDT). Evidence of unconfined DDT is shown when a flame is exposed to extremely high turbulence levels and rapidly accelerates to produce a spontaneous detonation wave in an unbounded area. This process is hypothesized to naturally occur in various phenomena such as interstellar combustion, Type Ia supernovae, and some hydrogen explosions. Until recently, these occurrences have only been successfully explored in computational simulations due to the difficulty of replicating the experimental conditions. To experimentally examine the mechanism of unconfined DDT, a premixed, hydrogen-air flame enters an unbounded area where it encounters a spherical gas consisting of hydrogen and oxygen. Shortly after, the shock-flame and soap bubble interact creating local flame acceleration and, in some cases, a local detonation. High speed schlieren is used to visualize the varying combustion regimes and compute the shock and flame velocities.