Turbochargers play a key role in producing more power for internal combustion engines of small aircrafts operated at altitudes with low air density. However, turbochargers can experience catastrophic failure at specific operating conditions where aeroelastic instability or resonance occurs at extremely high speed, putting the aircrafts at high risk. Therefore, it is critical to avoid turbocharger shaft speeds at these dangerous conditions, requiring accurate and stable turbocharger speed measurements. Currently, turbocharger speed sensors such as eddy-current sensors based on electromagnetic fields are often used for this purpose. However, these devices have limited capabilities such as applicable impeller blade material, maximum measurable speed, and small mounting distance between the sensor and impeller blade. Therefore, the present study focuses on development of a new turbocharger speed measurement method using a Fiber Optic Coupled Optical Speed (FOCOS) sensor and assessment of its capabilities. The FOCOS sensor was installed in the compressor housing of a high-speed turbocharger mounted on an internal combustion engine subjected to altitude conditions. These results were also compared to the performance of eddy-current speed sensors. It was concluded that the FOCOS sensor reliably measured turbocharger speeds over 250 kRPM for a range of impeller blade materials including aluminum and titanium, while the eddy-current speed sensor showed significant limitations.

In the present work, an optical sensor was developed and calibrated for the purpose of non-intrusive equivalence ratio measurements in combustion systems. The sensor incorporates a unique four-line, single-sensor chemiluminescence imaging-based approach, which relies on the ratio of C2* and CH* radical-species intensities to obtain measurements of equivalence ratios. The advantage of the four-line sensor is the use of additional filtering to mitigate broadband luminescence signals, and its improvements over conventional two-line chemiluminescence diagnostics are discussed. The sensor was calibrated using a premixed bluff-body jet burner with a propane–air flame operating over a wide range of equivalence ratios. The results showed that the four-line processing technique improved the signal-to-noise ratio of the chemiluminescence images for all test cases. Calibrations of C2*/CH* intensity ratio to equivalence ratio were developed for both the four-line and two-line techniques. The calibrations were then used to create maps of local equivalence ratios in the flame-holding region. The maps revealed a non-uniform field of equivalence ratios due to the nature of the radical-species intensity profiles within the flame. Therefore, special consideration is required for calibration in order to accurately quantify equivalence ratios and apply these to diffusion flames.

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.

Many practical flows in propulsion and energy are dynamic and three-dimensional. To study these flows researchers are continuously extending the dimensionality, speed, and resolution of imaging diagnostics. Tomographic Particle Image Velocimetry (Tomo PIV) is an example of an advanced diagnostic technique involving measurement of three-component velocity vectors (3C) over a three-dimensional volume (3D). Tomo PIV setups are complex and typically involve four or more cameras, a seeding system, and a laser for illumination. To design a Tomo PIV setup and experiment the user must define a large number of parameters such as: camera count, resolution, viewing angle, lens focal length, aperture, and standoff distance; laser energy and illumination volume; and seeding requirements. Due to the complexity involved, it can be difficult to plan and execute an optimized experiment without significant experience or trial-and-error experimentation. We present a software tool for planning and simulation of advanced 3D imaging setups such as Tomo PIV. The tool involves virtual experiments using synthetic flow data to compute the signal level, velocity error, and signal-to-noise ratio for a given setup. The tool can be used to plan or establish requirements for new setups, or to analyze and interpret error in existing setups. A description of the software tool is presented, followed by a demonstration of the tool involving application to some simple flows.

Lean flame blowout is experimentally investigated using a turbulent bluff-body combustor facility and high-speed 4D measurements. Bluff-body flames are stabilized at an equivalence ratio of Ф = 0.7 and lean blowout is induced through a rapid reduction of the fuel flow rate. Tomographic particle imaging velocimetry (TPIV) and CH* chemiluminescence imaging are simultaneously employed to capture the 3D flow field and 3D flame topology throughout the duration of blowout. Flame-vortex interaction dynamics are found to be the primary driving mechanism of extinction. As the equivalence ratio is reduced, localized extinctions are found to occur in spatial regions of high vorticity magnitudes. Localized extinctions are followed by large-scale fluctuations within the bluff-body wake and lead to global blowout.

The physical phenomena that control bluff body stabilized flames have been extensively studied because of their widespread use in many combustion technologies. Much of the understanding of bluff body physics has been acquired through experimental and anylitacal studies. The flow fields associated have been described with techniques such as chemiluminescence imaging, planar laser induced fluorescence (PLIF), particle image velocimetry (PIV), etc. Most of these techniques are two-dimensional in nature and rely on assumptions or negligence of the third component. Consequently, to validate previous findings, it is necessary to capture three-dimensional and time resolved data of the flow field associated with bluff body stabilized flames. This work describes the mechanisms of a reacting flow field associated with a cylindrical bluff body three-dimensionally using time resolved tomographic PIV and tomographic chemiluminescence imaging.

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 Rotating Detonation Engine (RDE) has been seen to exhibit lateral exhaust velocities, reducing useful thrust and limiting efficient power extraction. A previous study explored nozzle configurations on a RDE for rocket application with the intent to attenuate the exhaust swirl, thereby minimizing these lateral velocities. The work in this paper continues that research on a 5th Order Polynomial contour spike nozzle by numerous velocimetry analyses on the exhaust plume of the RDE by side- and back-end imaging. Through repetitive testing over a range of well-studied operating conditions, changes in the exhaust fluctuations with differing configurations of the spike nozzle integration assess nozzle influence on exhaust swirl.

Previous works have shown that the time averaged azimuthal velocity component exhibit overall low velocity magnitudes. On an instantaneous basis however, mitigation of local swirl in the exhaust of rotating detonation engines has proven necessary, which has made engine integration and optimum power extraction from the engine more difficult. The work in this paper outlines the design process for several novel back-end nozzle configurations for a rocket rotating detonation engine to dampen the fluctuating azimuthal velocity component. Velocimetry measurements are conducted in the exhaust plume of the RDE to assess nozzle influence on flow straightening.

Measurements are reported of the heat release profiles, the flame lengths, flame structure and other properties of a reacting jet-in-cross-flow (JICF) for two fuels. The air was heated to a static temperature of 1390K, which is above the autoignition temperature, and the air velocity was 468m/s, which is much larger than values that were considered previously. Aerodynamic strain rates are so large that the flame was expected to fall into either the “distributed reaction”, “thickened flamelet”, or “shredded flamelet” regimes. Fluorescence images of CH, OH and formaldehyde identified the flame structure. The jet-in-cross-flow is a unit physics problem that occurs in turbojets and scramjets. While scaling relations are known for the non-reacting case, more information about the reacting case is needed, especially when autoignition and strain rates become important. Three regions were identified. In the liftoff region autoignition reactions occur which create a strong formaldehyde PLIF signal. However, flames and heat release do not occur in the liftoff region since CH and CH∗ signals were negligible. The second region is the lifted flame base, which has the character of a premixed flame, as evidenced by a very rapid rise in the heat release rate as indicated by the CH∗ and OH∗ signals. The third region contains a turbulent non-premixed flame and the CH images indicate the presence of thickened and shredded flamelets. The 2–3mm thickness of each CH layer is more than 10times the laminar flamelet thickness. In the third region the heat release rate decays slowly downstream, which is typical of a non-premixed flame. Because both upstream autoignition and downstream thickened flamelets were observed, we classify this combustion to be an “autoignition-assisted flame”. Flame lengths increase linearly with fuel mass flow rate, indicating that mixing is controlled by the air velocity rather than the fuel velocity.

PLIF imaging is used to identify the internal structure of a partially-premixed lifted turbulent jet flame within a cross-flow of air that is heated to a temperature of 1364 K, which is above the autoignition temperature. Large aerodynamic strain rates were imposed since the velocity of the cross-flow air is 480 m/s. A jet flame-in-a-cross-flow is a simple unit physics problem for which scaling relation are known for the non-reacting case. However, the present work addresses the roles of two factors: autoignition chemistry and large aerodynamic strain rates. This unit physics problem is important in propulsion devices, including ramjets and in internal combustion engines. Formaldehyde PLIF is used identify the distributed auto-ignition regions, while CH and OH PLIF identifies the (downstream) primary reaction regions. The primary reactions are shown by the CH images to have the structure of broken and thickened flamelets under these high-strain rate conditions. That is, the CH occurs within a 2-3 mm thick wrinkled interface that is not continuous but contains many holes due to local extinction. The formaldehyde PLIF images show that there is a broad 10 mm thick region of formaldehyde upstream of the primary thickened flamelets. This broad formaldehyde region indicates that distributed (non-flamelet) reactions occur early in the chemical reaction process due to auto-ignition. Therefore we classify this combustion to be an “auto-ignition-assisted flame”. The measured profiles of heat release rate are presented, which are useful in assessing modeling efforts.

To model the performance and operability limits of a dual-mode scramjet engine, the heat release distribution must be accurately predicted. This distribution controls the thermal choking point, and thus the proflles of pressure, Mach number, and heat transfer in the engine. The current research efiort consists of two parts: measurement of heat release distributions from OH* and CH* in a dual-mode combustor, and development of a 1-D scramjet engine performance model. A dual-mode combustor with transverse wall fuel injection and a cavity ∞ameholder is investigated experimentally for air stagnation temperatures of 1270-1520K. The upstream region of the heat release distribution depends on the ∞ame stabilization and spreading. The ∞ame length and downstream region of heat release distribution appear to be mixing limited for all cases. This result is used to develop a combustion model for a quasi-1-D scramjet combustor code.

This paper provides details of the combustion and inlet submodels used in the Michigan-Air Force Scramjet In Vehicle (MASIV) model.The model solves conservation equations in 1-D, using several modeling techniques to retain some of the fidelity of higher-order simulations.Inlet wave interactions, fuel mixing and finite-rate chemistry are considered.The order of the problem is reduced by physics-based, experimentally-verified algebraic scaling laws, which retains the required physics but reduces the computation time of the problem to seconds, instead of the several days required by computational fluid dynamics (CFD).Scaling coefficients and assumptions are given.The model is used to compute the performance of an experimental configuration for which real data are available.

Combustion characteristics of a laboratory dual-mode ramjet/scramjet combustor were studied experimentally. The combustor consists of a sonic fuel jet injected into a supersonic crossflow upstream of a wall cavity pilot flame. These fundamental components are contained in many dual-mode combustor designs. Experiments were performed with an isolator entrance Mach number of 2.2. Air stagnation temperatures were varied from 1040 to 1490K, which correspond to flight Mach numbers of 4.3–5.4. Both pure hydrogen and a mixture of hydrogen and ethylene fuels were used. High speed imaging of the flame luminosity was performed along with measurements of the isolator and combustor wall pressures. For ramjet mode operation, two distinct combustion stabilization locations were found for fuel injection a sufficient distance upstream of the cavity. At low T0, the combustion was anchored at the leading edge of the cavity by heat release in the cavity shear layer. At high T0, the combustion was stabilized a short distance downstream of the fuel injection jet in the jet-wake. For an intermediate range of T0, the reaction zone oscillated between the jet-wake and cavity stabilization locations. Wall pressure measurements showed that cavity stabilized combustion was the steadiest, followed by jet-wake stabilized, and the oscillatory case. For fuel injection close to the cavity, a hybrid stabilization mode was found in which the reaction zone locations for the two stabilization modes overlapped. For this hybrid stabilization, cavity fueling rate was an important factor in the steadiness of the flow field. Scramjet mode combustion was found to only exist in the cavity stabilized location for the conditions studied.

A new scramjet engine model has been developed to support hypersonic vehicle design studies and flight dynamics and control system analysis.This paper explains the methodology and the governing equations for the new propulsion system model that is suitable for use with a control oriented dynamic model of a hypersonic vehicle.Previous propulsion models used for this purpose were based on simple Rayleigh flow for the combustion process, but despite this, captured the propulsion system interactions with the vehicle aerodynamics and structural dynamics.A new, higher fidelity propulsion system model is constructed that simulates numerous phenomena that were neglected in the Rayleigh flow approach.The new model is of higher fidelity, and therefore it is not designed to calculate the flow physics on a timescale that is suitable for dynamics and control simulations.Instead it will be used as a truth model and the starting point for the derivation of a reduced-order model.Specific phenomena that are included in the new model are: a pre-combustion shock train within the isolator and its interactions with the combustor, the loss of stagnation pressure due to gas dissociation and recombination, wall heat transfer and skin friction, a fuel-air mixing submodel, and a finite-rate chemistry and autoignition reaction mechanism.It is shown that the new propulsion system model expands the operability envelope as compared to the previous model by accommodating ramjet combustion, which occurs at high supersonic/low hypersonic flight Mach numbers.