A standing normal detonation mode of combustion consisting of a normal shock coupled with heat release is realized in an experimental high-speed reacting-flow facility. The normal detonation is stabilized using a 2D ramp where the high-enthalpy freestream Mach number and reactant composition is equivalently matched to the Chapman-Jouguet (CJ) consumption speed of the detonation, at M-infinity/M-CJ = 1.06. High resolution optical measurements of OH* chemiluminescence and density gradients from schlieren clearly show the close-coupling between the normal shock and the heat release of the standing detonation. A ZND analysis have been conducted using the boundary conditions where the induction length is found to closely matches the experimentally measured induction length. The agreement between the induction length scales and the freestream Mach number to detonation CJ Mach number confirm the realization of a standing detonation mode of combustion.

The study of supersonic combustion mechanics is becoming increasingly important in the development of hypersonic propulsion systems. One approach to supersonic combustion is the use of standing oblique detonation waves. Prior experimental testing to stabilize oblique detonations has utilized a ramp on the wall of the combustor to initiate a detonation, causing potential interactions between the shock and boundary layer. A wedge positioned off the combustor wall will eliminate this shock-boundary layer interaction. Non-reacting numerical simulation of the wedge was performed.

Advancements in the study of oblique detonation wave (ODW) initiation and stabilization have primarily been within the purview of numerical simulations and limited laboratory experiments. However, the recent publication of an experimentally controlled stabilized ODW via a 30-degree ramp by the HyperReact facility at the Propulsion and Energy Research Laboratory represents a breakthrough towards developing standing detonation engines as a reliable propulsion system of the future. Additional work was conducted with ramps at 12-degrees and 20-degrees, although ODW was not achieved. Comparison of those results with the successful test cases using the 30-degree ramp can provide insight into the system. In that vein, this paper presents a parametric study of varying ramp angle configurations at low (I) to moderate (II) regime conditions, where total pressure and total temperature denote the regime where the test cases reside. For the basis of comparison, diagnostics using simultaneous high-speed shadowgraph and broadband chemiluminescence are used to visualize shock structures and combustion products within the hypersonic flow.

This paper details the design and computational analysis of a Mach 5 converging-diverging nozzle array for use in the Hypersonic Wind Tunnel at the Propulsion and Energy Research Lab at the University of Central Florida. This nozzle array is designed to create a highly turbulent, well-mixed flow for the study of turbulence-compressible reactions, including ongoing work in the study of oblique detonation waves (ODW). The nozzle array consists of a grid of axisymmetric CD nozzles, the contours for each being made using the method of characteristics. Multiple nozzle arrays with varying nozzle spacing, individual nozzle sizes, and number of nozzle instances were numerically simulated. These numerical simulations determined which design candidate produced the most well-expanded flow with the turbulence values most well-tailored to the investigation.

A fueled cavity in a supersonic crossflow was ignited via a pulse detonator (PD) producing detonation waves that were then decoupled to produce varying degrees of shock-flame separation at the exit of the PD tube. This decoupling allowed for observation of the cavity ignition mechanism, and the key parameters required for successful cavity ignition were identified. Measurements were made using high-frame-rate OH Planar Laser-Induced Fluorescence (PLIF) and schlieren and chemiluminescence imaging. It was shown that the entrainment of high-temperature intermediate species into the forward region of the cavity, immediately behind the step, is the principal criterion for cavity ignition. Both coupled and slightly decoupled detonation cases induced significant OH shedding into the step region, leading to ignition and flame stabilization within the cavity. At conditions where OH shedding into the step region did not occur, cavity ignition was not observed. In coupled and slightly decoupled cases, there is more shedding of OH behind the step due to the greater disturbances created in the flowfield. As the degree of detonation decoupling increases, there is less shedding of OH and therefore a lower likelihood of ignition. Additionally, the time required for cavity combustion to reach its steady-state condition varied with the degree of decoupling of the detonation. Coupled detonation cases were shown to be more disruptive to the cavity and thus required more time to reach steady state than the decoupled cases.

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.

This report details the in-situ use of Raman spectroscopy to determine the approximate time-averaged fuel distribution within a high-enthalpy hypersonic reacting facility (HyperReact). The HyperReact facility is being used to study Oblique Detonation Wave (ODW) stabilization, which is in part dependent upon local fueling levels, creating the need for these measurements. A mixture of high-purity hydrogen and air with flow Mach numbers ranging from Mach 4.4 to Mach 5 is produced. Measurements were made using a 532 nm laser to excite the hydrogen fuel and induce the Raman shift, which then emits at approximately 683 nm. A CCD camera filtered with a 680 nm bandpass filter (10 nm FWHM) records the response within the optically accessible test section. The intensity of the response along the beam path correlates to the local hydrogen concentration, which was used to calculate the equivalence ratios through the height of the test section at the measurement location for the conditions of interest.

In the ongoing search for stable oblique detonation waves (ODW), experimental studies are often limited in their operating temperatures and flow Mach numbers due to material and design limitations. Several past numerical studies have shown that high total temperatures (T0 > 1200 K) and freestream Mach numbers (often in excess of Mach 5) are needed to successfully initiate and sustain ODW. In this study, Mach 4.4+ flows at moderate total temperatures (800 k < T¬0 < 1200 K) are explored experimentally to determine if conditions can be tailored to allow for ignition and sustained reactions at these lower bulk temperatures. Variations in temperature, pressure, and mixture composition allowed for the study of boundary layer ignition and shock induced combustion (SIC). At total temperatures ranging from approximately 800 K to 1200 K and total pressures of approximately 3.50 MPa to 5.75 MPa, intermittent or periodic reactions were seen. The structure and behavior of these reactions, imaged using shadowgraph and broadband chemiluminescence, suggest that stabilization of an ODW may be possible with additional condition tailoring within the operational range of this test facility.

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 ignition of a cavity stabilized flame in a supersonic flow, as would be used in a scramjet, presents numerous challenges. Recent works have studied the viability of utilizing a pulse detonator as the source of energy for the ignition. The high-pressure and temperature output of such a device allows for it to successfully engage and ignite the mixture throughout the cavity under numerous conditions. However, this is contingent on the leading shock and reaction front of the detonation remaining closely coupled. A weakened or fully decoupled detonation is less likely to ignite the cavity due to decreased engagement of the detonation products with the contents of the cavity. This study makes use of multiple optical diagnostic techniques (schlieren and chemiluminescence imaging and OH PLIF measurements) in order to observe how the decoupling causes the aforementioned reduction in engagement. It is shown that weaker detonations introduce a lag time between the entrance of the leading shock into the cavity and the appearance of combustion products. This is followed by additional combustion occurring within the exhaust plume of the pulse detonator but less spreading into the cavity. Differences in the area of the cavity affected by the detonation and its combustion products at each condition are presented.

Past research into oblique detonation wave (ODW) initiation, stabilization, and behaviors in hypersonic flows have been largely computational in nature due to the complexity of experimental studies into the subject. This paper discusses an ongoing experimental investigation on this topic through the use of a supersonic wind tunnel capable of producing a highly turbulent, preheated Mach 4.4+ flow. A ramp is placed into the flow to create an oblique shock wave (OSW). Current results show potential shock induced combustion (SIC) initiated from the ramp OSW and the associated reflected shock, as well as reactions within the boundary layer on the ramp surface or immediately in front of the leading edge of the ramp. Three ramp angles were tested: 12-degrees, 20-degrees, and 30-degrees. The OSWs created by larger ramp angles produce a higher temperature rise and slower post-shock flow, creating more favorable conditions for ignition in the post-shock region. Increasing total temperature also increases the likelihood of ignition. Reactions are first seen at T0 ≈ 850 K, although they are intermittent and unstable. The current results show flow regimes in which reactions can be expected and are a key intermediary step in the process of creating a sustained ODW within this facility.

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).