Hybrid Aircraft Propulsion System Synthesis (HAPSS) is a computer program that sizes and analyzes pure-series hybrid electric propulsion systems for aircraft. The development of this program began during a NASA SBIR contract, in conjunction with Empirical Systems Aerospace (ESAero), with the creation of a propulsion fan design tool. Since the completion of this contract in July 2010, the HAPSS program has been expanded to combine the many aspects of a hybrid propulsion system such as the propulsive fans, electric motors, generators, and controllers, and the internal combustion engines. This thesis describes the benefits and drawbacks of aircraft hybrid propulsion systems to reveal the usefulness of a program of this nature. The methodology behind HAPSS, the creation of the program, its operation, and its many applications are also discussed in detail. Finally, this thesis includes a brief example in which HAPSS is used to analyze a hybrid propulsion system for a commercial transport aircraft. This example demonstrated the usefulness of the program and revealed interesting behavior and trends unique to hybrid propulsion. To date, the HAPSS program has been utilized on several different contract projects in which an aircraft hybrid propulsion system was designed. In the summer of 2012, a government organization in conjunction with ESAero will begin funding a contract to continue the development of HAPSS by adding functionality and improving accuracy while making the tool available to other government agencies.

Experiments were conducted on the Cal Poly air augmented rocket (AAR) in order to characterize two-dimensional flowfield phenomenon occurring in the mixing duct. The testing utilized a direct connect system where high pressure nitrogen is fed into the combustion chamber, to form a primary flow. The high pressure nitrogen is then expanded through a nozzle, with an area ratio of 22 and an exit area of 0.75 in2, up to Mach 4.3. Secondary air is entrained from a plenum chamber which is used to create a lower stagnation pressure for the secondary flow. The two flows mix in a duct that has a cross sectional area of 2.06 in2. The maximum pressure ratio, the ratio of primary to secondary stagnation pressure, achieved during testing was 132. The stagnation pressures of the primary and secondary flows are transient throughout the test. The quasi-steady portion of each run increased with increasing pressure ratio. Pressure and temperature measurements were collected from ten test runs. Shadowgraph images were taken of the mixing duct during testing in order to image the interactions between the primary and secondary flows. The images show an oblique shock forming in the primary flow. The angle of the shock matches theoretical predictions to within 8.41%. The oblique shock begins at a distance of 1.5 inches downstream of nozzle exit when the AAR is operating in the Fabri choked condition. The images also show the mixing region which forms between the primary and secondary flows. The mixing region represents as much as 25% of the cross-sectional area of the flow field in the mixing duct two inches downstream of the nozzle exit. An analysis of the secondary Mach number in the mixing duct shows that Fabri choking is occurring during testing. The secondary Mach number decreases as pressure ratio increases, in the Fabri choked condition. The transition to Fabri choking occurs at a pressure ratio of 100, suggesting that this is the pressure ratio of the saturated case. The shape of the primary plume was compared to results from a 2-D simulation developed to predict the flow field inside the Cal Poly AAR. Although, the simulation is unable to predict the entire flowfield, modifications made it able to predict the velocity of the secondary, entrained, flow within 3.7%. The modified simulation also predicts the that the primary plume will have expanded 98% of its total distance from the centerline of the mixing duct 1.7 inches downstream of the primary nozzle exit. Pressure data taken along the wall of the mixing duct was used to identify the location of Fabri choking in the mixing duct. Tests showed that Fabri choking is occurring between 1 inch and 2.5 inches downstream of the nozzle exit. The location of Fabri choking moves farther downstream of the nozzle as pressure ratio increases.

A bio-derived fuel, lard, was successfully tested using a laboratory scale hybrid rocket motor and a static test stand at the University of Tennessee-Knoxville. The experimental setup used gaseous oxygen as the oxidizer. Twenty-three experimental tests were successfully conducted with lard and oxygen. The nine-inch fuel grains used in the current investigation produced a measured thrust ranging from 70-145 Newtons (15-33 pounds) with calculated specific impulses ranging from 122-181 seconds. All the tests conducted were intensely fuel rich, and had equivalence ratios ranging from 0.2 to 0.45. The low equivalence ratios are partially due to unburned fuel particles that exit the nozzle. The tests conducted have shown that the regression rate of the lard was higher than that of other fuels, such as HTPB (hydroxyl-terminated polybutadiene), used in hybrid rocket motors. Lard has produced results similar to those obtained by researchers at Stanford University using paraffin. This investigation has provided sufficient evidence to indicate that lard merits further study as a fuel for hybrid rocket motors.

A bio-derived fuel, lard, was successfully tested using a laboratory scale hybrid rocket motor and a static test stand at the University of Tennessee-Knoxville. The experimental setup used gaseous oxygen as the oxidizer. Twenty-three experimental tests were successfully conducted with lard and oxygen. The nine-inch fuel grains used in the current investigation produced a measured thrust ranging from 70-145 Newtons (15-33 pounds) with calculated specific impulses ranging from 122-181 seconds. All the tests conducted were intensely fuel rich, and had equivalence ratios ranging from 0.2 to 0.45. The low equivalence ratios are partially due to unburned fuel particles that exit the nozzle. The tests conducted have shown that the regression rate of the lard was higher than that of other fuels, such as HTPB (hydroxyl-terminated polybutadiene), used in hybrid rocket motors. Lard has produced results similar to those obtained by researchers at Stanford University using paraffin. This investigation has provided sufficient evidence to indicate that lard merits further study as a fuel for hybrid rocket motors.

The current research investigates a next-generation method of combustion called detonations for propulsive applications which has support from the Combustion Services Branch at the Air Force Research Laboratory (AFRL) of the Rocket Propulsion Division. The overall initiative seeks to explore the viability of detonation-based systems to augment or replace current rocket propulsion methods in the coming five years. The present work progresses this vision by providing proof of a liquid RP-2 based Rotating Detonation Rocket Engine (RDRE) as well as the fundamentals associated with the physics involved to further the understanding of the phenomenon and its potential uses. Phase one of the research starts with development of a theoretical model based on the Zel'dovich-Neumann-Doring (ZND) detonation model and the well established D2 droplet burning model to establish a droplet sizing parameter for varying detonation wavespeeds. Phase two then documents the interaction of a millimeter droplet colliding with varying supersonic waves where the goal is to document the breakup and transient evolution of the droplet when imparted with the supersonic flow velocity. Phase three consists of characterization of an unlike-doublet impinging injector using Liquid-RP2 and Air to record the incipient spray behavior relative to varying operating conditions of interest. Phase four comprises of the aforementioned spray interacting with a fully developed detonation wave where Schlieren, CH*, Formaldehyde PLIF and Mie scatter were used to archive the reactions that ensues. Finally, phase five consists of testing a RDRE fueled with liquid RP-2 and O2 as the reactants where thrust measurements and back-end imaging analysis of the wavespeeds provide insight into the operation and sustenance of the combustion phenomenon occurring inside the engine.

Liquid jet in crossflow systems are often used as lightweight and efficient mechanisms of atomizing fuel prior to entertainment in the flame holder and combustion, making them integral components of liquid fueled engines. Unfortunately, such systems are susceptible to deviations in both trajectory and breakup rate, depending primarily on the Weber number and momentum flux ratio of the injected jet. In these studies, the effects of solid obstructions, called pintiles, on the variability of liquid jet in cross flow trajectory and breakup are investigated. Initial investigations looked at the impacts of broad geometric parameters on flow independence, using Mie scatter imaging and phase Doppler particle analysis. The results of that investigation yielded an optimal overarching geometry for pintiles. This knowledge was then refined by looking at specific face characteristics of the obstructions, primarily investigating face angle and concavity. Spray characteristics were spatially resolved using LIF/Mie particle sizing techniques, revealing that modest convex surfaces yielded the most consistent breakup characteristics across space, while simultaneously improving the average breakup distance of the liquid jet. Finally, this progression of pintile characteristics is investigated on the effects pintiles have on overarching flame properties, using C2*/CH* chemiluminescence ratios to determine spatially resolved equivalence ratio distributions across a wide range of Weber numbers and momentum flux ratios encompassing breakup regimes from the enhanced capillary modes through to shear breakup modes. Results from these studies demonstrate significant improvement of combustion properties from the introduction of the pintiles.

Detonation-based propulsion systems are desired due to their potential for thermodynamic cycle efficiencies much higher than traditional, constant pressure deflagration-based systems. Additionally, the ability to propagate in high-Mach number flow regimes makes detonations promising for supersonic combustion applications. Oblique detonation waves (ODW) are a subset of detonations that have been proposed for use in a number of detonation-based propulsion concepts. However, most studies into initiation and stabilization of ODWs have been numerical studies without sufficient experimental validation. Experimental setups, such as expansion tubes, typically have extremely limited run times (< 100 ms) due to practical difficulties in attaining the necessary conditions in a continuous flow ( > 1 second) system. Those difficulties include the need for high pressures, temperatures, and the correct chemical composition in order to create a supersonic flow that is favorable for detonations. The work described within this dissertation focuses on the facility design and investigations into supersonic combustion and detonation stabilization using continuous flow facilities, including the HyperReact Facility. Using HyperReact, tests were conducted with varied pressures, temperatures, geometries, and flow compositions. From those tests, the operability map of the system was created, from which 3 major reaction regimes were defined. Those regimes include: intermittent, low-intensity reactions (regime I), intermittent, higher-intensity reactions (regime II), and quasi-stable reactions (regime III). Multiple diagnostics were performed, including high-speed shadowgraph and chemiluminescence imaging, Raman spectroscopy, as well as pressure and temperature measurements at multiple locations.

Conventional engines are limited by the efficiency of their combustion mode. Compared to present constant pressure deflagration-based engines, detonation-based systems can realize a higher thermodynamic cycle efficiency, making them an attractive candidate for next generation propulsion systems that will take humanity to hypersonic speeds and even to Mars. For all its performance gains, detonation engines are still far off from implementation. One system, the rotating detonation engine (RDE) is promising as a detonation-based engine concept for its stability, simplicity, and versatility. For these reasons, RDEs have been the subject of studies internationally in efforts to understand their operation and integration into conventional technology. RDEs are on the cusp of field use, considered at technology readiness level 5 with prototype demonstrations occurring today; however, there are still significant barriers holding back this technology from widespread adoption. The work of this dissertation confronts each of these barriers with experimental methods. Using multiple different RDE test facilities, investigations into injection, fueling, exhaust, detonability, and integration were conducted, targeting research gaps in each barrier. As a result, many novel advancements have been made from these studies such as the first demonstration of hydrogen and oxygen rotating detonations, the detonability of sustainable solid particle fuels, and the effect of fuel stratification on rotating detonation propagation. Altogether, the work presented depicts the RDE from a complete perspective by advancing current RDE research through multiple channels with the intention of advancing the technology readiness level of RDEs.

Propulsion systems are influenced by the efficiency of combustion systems. One approach to substantially improve combustion efficiency is through pressure gain combustion or detonation-based engines. Detonations exhibit attractive features such as increased stagnation pressure and rapid heat release; however, their highly unsteady and three-dimensional nature makes them difficult to characterize. In addition, the deflagration state prior to detonation is not well defined experimentally. Detonations can be achieved via the deflagration-to-detonation transition (DDT), where a deflagration that propagates on the order of 1 – 10 m/s is accelerated to a detonation that propagates on the order of 2000 m/s. The DDT process is highly dynamic and can occur through several mechanisms such as the Zeldovich reactivity-gradient mechanism where hot spots are created by Mach stem reflections, localized vorticial explosions, boundary layer effects, or turbulence. This work focuses on transient compressible flame regimes within the turbulent DDT (tDDT) process which causes a flame to undergo various burning modes. These burning modes can be categorized into four regimes: (1) slow deflagrations, (2) fast deflagrations, (3) shock-flame complex, and (4) detonation. To achieve each burning mode, turbulence levels and propagation velocities are tailored using perforated plates and various fuel-oxidizer compositions. The primary goal of this dissertation is to characterize the relationship between the turbulent flame speed (ST) and Chapman-Jouguet (CJ) deflagration speed (SCJ) using high-speed optical diagnostics in a turbulent shock tube facility. This work will: (1) further validate and classify the turbulence-compressibility characteristics associated with fast flames that lead to detonation onset in a highly turbulent environment, (2) quantify local ST for fast flames, and (3) investigate the flow field conditions of flame modes relating to the SCJ criteria, from slow deflagrations to shock-flame complexes.

The work presented herein consist of first studying the instantaneous properties of the detonation waves in a rotating detonation rocket engine by tracking each individual wave and recording its position, velocity, and peak intensity as it travels around the annulus. Results for a steady portion of a test performed on a rotating detonation rocket engine show that the wave properties exhibit oscillatory behavior. Results obtained from the rotating detonation rocket engine show that the properties are highly dependent on the azimuthal position. In an attempt to understanding the cause of such a behavior, similar investigations were performed on an air-breathing rotating detonation engine with a different injection design to see if the behavior persists. Results show that air-breathing rotating detonation engines do indeed exhibit this behavior in a more attenuated fashion. It is demonstrated that the pre-detonation hole might be the reason for the observed combustion instabilities. After establishing the steady state behavior of a single mode in a rotating detonation rocket engine, transient analyses of multiple tests were performed in order to capture the relative wave speeds between the modes. Wave speeds and operational frequency plots showcasing the range of operation of each mode (single and counter-propagating) were constructed. Moreover, operating maps of the engine were built and clearly demonstrate where each mode resides. The mode transition instability phenomenon observed in rotating detonation rocket engines is then studied. Each mode transition is distinguished by different mechanics and behavior requiring different diagnostic tools and research techniques to analyse. In this investigation, five possible mode transitions in rotating detonation engines have been identified and are Types AS, DS, AO, DO and SO and their behavior is discussed. Also, the counter-propagation wave behavior within an intermidiate period for Type _O mode transition have been discussed.