Coherent anti-Stokes Raman scattering (CARS) was used to perform temperature measurements, over short durations, in a model rocket combustor operated at elevated pressures. The combustor uses an oxidizer centered single shear-coaxial injector. Experiments were performed with gaseous hydrogen as fuel and oxygen as the oxidizer. We investigated H2/O2 flames with adiabatic flame temperatures in excess of 3000 K and chamber pressures in excess of 7 MPa.

Advanced gas turbine combustion strategies, such as axially staging the fuel, are of great interest due to their potential to increase cycle efficiency while maintaining low levels of pollutants. In our previous work within a staged gas turbine model combustor, we used exhaust gas emissions measurements to demonstrate a significant NO X reduction by increasing the combustor exit Mach number, even at a constant residence time. In this paper, the development of an optically accessible secondary combustion zone (SCZ) to further study the injection of a reacting jet into a high-speed vitiated crossflow is described. Measurements were targeted for a nominally 1700 K vitiated crossflow, a premixed jet at an unburnt temperature of nominally 500 K, and a combustor pressure of 500 kPa. Key aspects of this design challenge include the high-speed and high temperature crossflow leading to relatively high convective heat flux at the inner surface of the windows and the necessary use of a narrow channel for the combustion gas. Emphasis is placed on the critical design features: a double-windowed design, an air-cooling scheme based on forced convection of air between the inner and outer windows, and intricate water-cooling circuits for the metal hardware. These design features have enabled long-duration, steady-state operation despite elevated pressure, high combustion gas temperatures, and high-speed reacting flows. The SCZ has survived operation at MW thermal powers over several hours of continuous operation and over a dozen test days to date; a set of windows retained good transparency without discoloration for typically 3–4 test days. Survivability of the windows through the air-cooling design has enabled us to study the reacting jet-in-crossflow at the desired high-speed conditions, without risking disturbing the physics with a window film-cooling flow. The capability to acquire useful measurements is illustrated using chemiluminescence imaging, pressure measurements, and emissions sampling.

The development and performance of a perforated plate burner (PPB) operating using premixed natural gas and air at engine-relevant inlet temperatures and combustor pressures with thermal powers up to 1 MW is discussed. A significant benefit of using burners with simplified flow fields, such as the PPB, for experimental studies in the laboratory is the potential for decoupling the complex fluid dynamics in typical combustors from the chemical kinetics. The primary motivation for developing this burner was to use it as a source of vitiated flow with negligible swirl for reacting jet in vitiated crossflow experiments. The design methodology for the PPB is described, including plate geometry selection and flashback mitigation features. The stable operation of the PPB within a high-pressure test rig was validated: successful ignition, effective use of red-lines for flashback mitigation, and long duration steady-state operation in both piloted and nonpiloted modes were all observed. Exhaust gas emissions measured using a Fourier-transform infrared (FTIR) spectrometer showed very good performance of the PPB in terms of the combustion efficiency (based on measured CO and UHC), and a stability diagram of the PPB was developed as a function of the equivalence ratio and the PPB hole velocity. FTIR measurements also showed very low levels of NOX in nonpiloted operation that were generally within 3 ppm (reported dry and referenced to 15% O-2). The capability for steady-state operation, high combustion efficiency, and low levels of NOX makes this PPB an excellent burner candidate for combustion experiments in the laboratory.

Nitrous oxide (N2O) has gained popularity as a unique oxidizer for propulsion applications due to its ability to decompose exothermically, producing nitrogen and oxygen. In the current work, the flame acceleration, deflagration-to-detonation transition, and detonation properties of bipropellant mixtures with N2O as the oxidizer are studied for potential applications in pulsed blowdown and detonation-driven thrusters. These properties are compared with those in mixtures with oxygen (O2) or nitrogen tetroxide (N2O4) as the oxidizer. The performance of N2O versus O2/N2O4 for detonation engine applications is investigated using theoretical Chapman-Jouguet detonation calculations of bipropellant systems with ethylene (C2H4) and acetylene (C2H2) as fuels. A critical requirement for the application of bipropellant mixtures to pulsed propulsion systems is rapid flame acceleration to achieve significant chamber pressure rise in a short distance with the potential for a prompt transition to detonation. This deflagration-to-detonation transition behavior of mixtures using C2H4 and C2H2 with N2O and O2 is investigated for increasing initial pressures in the experimental portion of this work. While C2H2 is a highly energetic fuel with theoretically high performance, it presents serious practical storage concerns when considered for propulsion applications. These practical issues motivate investigation of C2H4 as a potential alternative fuel, which is relatively easy to manage. The precompression of the bipropellant mixtures during flame acceleration is also estimated and compared.

Advanced combustion technologies to limit NOx production are needed to meet rigorous emissions standards, which are required due to the harmful effect of NOx on the environment. One such advanced concept involves axially staging the fuel to create a staged combustion system. This paper reports on emissions measurements obtained for premixed natural gas and air reacting jets into vitiated crossflow with negligible swirl at conditions corresponding to elevated inlet temperatures of 500-600K and an elevated combustor pressure of 500 kPa. A significant NOx reduction was achieved when the staged combustor exit Mach number was increased and the axial residence time was reduced. Based on this preliminary investigation, a test matrix was developed to independently vary the exit Mach number for a constant axial residence time by utilizing modular rig hardware to change the length of the axial combustor. Up to 70% reduction in NOx produced by the axial stage was observed when the combustor exit Mach number was increased from about 0.26 to 0.66 at a constant residence time of 1.4 ms. NOx reduction based on variation in the Mach number and at a constant residence time has not been previously reported in the literature to the best of our knowledge. This decrease in NOx is hypothesized to be due to the lower static temperature of a compressible flow and potentially better mixing of the jet with the crossflow due to the interaction occurring at high speeds. The effects of axial residence times on NOx emissions were also investigated at a constant Mach number of 0.44 for axial residence times between 0.8 and 2.3 ms. Varying the axial residence time was observed to have a strong effect on NOx emissions for exit total temperatures greater than 1900K, which agrees well with existing trends. (C) 2019 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

Large-scale extraction of oil and natural gas requires an effective method of generating a high surface area network of fractures, or the stimulation of existing fractures, in a formation in order to increase permeability. Conventional hydraulic fracturing has limited utility in this application. In this work, Sandia National Laboratories is exploring high rate pressurization techniques employing tailored energetic materials systems to control both pressure rise rate and peak pressure in order to optimally stimulate potential rock formations. Rapid pressurization, at rates far exceeding quasi-static conventional hydraulic rates, can generate multiple radial well bore fractures and potentially provide a mechanism to induce shear destabilization within the formation that enables the fractures to be self-propping. Multiple fractures from the well bore allow efficient coupling to the existing formation fracture network and increase near field well bore permeability. Furthermore, these techniques can allow for repeated stimulations and produce energetic events within the fractures thereby allowing fractures to be extended further. Controlled rate pressurization is a useful tool for the efficient generation of fracture networks and has potential application to increase oil and gas production. This paper provides an overview of the concept of controlled rate pressurization, laboratory experiments and field trials that are being conducted. The present work investigates detonations of a stoichiometric mixture of ethylene and nitrous oxide (C2H4 + 6N(2)O) at high initial pressures ranging from 0.862 to 2.068 MPa as a method of fracturing rock below the ground surface. The experiments investigate the fracture generation as a function of the initial pressure of the mixture. In the current configuration, the combustion reaction is initiated by an electrically fired igniter at the ground surface and quickly transitions to a detonation. The experimental setup accommodates one high pressure (690 MPa) transducer, placed downstream of the igniter, to measure peak pressure. The pressure transducer recorded peak pressures, which are 2.3-2.6 times in excess of the Chapman-Jouguet (CJ) values as a result of pre-compression of the unburned gas mixture during the flame acceleration prior to deflagration-to-detonation transition (DDT). Analysis of the data indicated an increase in rock permeability due to detonations and this was confirmed by the core drilled sections at the test site.

Following three decades of research in short duration facilities, Purdue University has developed an alternative turbine facility in view of the modern technology in computational fluid mechanics, structural analysis, manufacturing, heating, control and electronics. The proposed turbine facility can perform both short transients and long duration tests, suited for precise heat flux, efficiency and optical measurement techniques to advance turbine aero-thermo-structural engineering. The facility has two different test sections, linear and annular, to service both fundamental and applied research. The linear test section is completely transparent for visible spectra, aimed at TRL 1 and 2. The annular test section was designed with optical access to perform proof of concepts as well as validation of turbine components at the relevant non-dimensional parameters in small engine cores, TRL 3 to 4. The large mass flow (28 kg/s) combined with a minimum hub radius to tip radius of 0.85 allows high spatial resolution. The Reynolds (Re) number extends from 60,000 to 3,000,000, based on the vane outlet flow with an axial chord of 0.06 m and a turning angle of 72 deg. The pressure ratio can be independently adjusted, allowing for testing from low subsonic to Mach 3.2. To ensure that the thermal boundary layer is fully developed the test duration can range from milliseconds to minutes. The manuscript provides a detailed description of the sequential design methodology from zero -dimensional to threedimensional unsteady analysis as well as of the measurement techniques available in this turbine facility.

I njector spray characteristics have a significant influence on the combustion performance in a gas turbine engine, including an impact on dynamics, emissions and component life. Furthermore, commercial aviation faces fuel cost, environmental, and energy security challenges that arise from the use of petroleum based jet fuels. Sustainable alternative jet fuels can help address these challenges and need to be characterized in their spray performance. The present work describes the detailed characterization of several alternative fuels using a hybrid airblast atomizer on the basis of spray shape, droplet size and velocity distribution at a range of operating conditions including fuel temperature, injector pressure drop and spray chamber pressure and temperature. The characterization is done using optical patternation, phase Doppler anemometry (dual-PDPA) and high speed back-lit imaging. The measurements obtained as part of this work provide the validation data-set for computational modeling of the spray behavior which forms a critical part of the broader project. The results show a strong influence of the fuel temperature on the spray, with lower temperature (290 K to 240 K) decreasing the atomization quality by 14%, while the effect of fuel injection pressure on the spray is minimal. A large effect of pressure drop across the injector is seen on the spray, with a change from 2% to 6% leading to a decrease in drop size of up to 36%, which can result of a shift in the secondary breakup regime of the spray.

A test article has been specifically designed in order to investigate the methane behaviour inside rocket engine cooling channels. The test article is composed of a suitable copper-alloy block warmed up by cartridge heaters and of a single channel with rectangular cross section, which is fed with transcritical methane flow. Steady-state conditions and channel dimensions are representative of a typical rocket engine cooling channel. Several tests have been conducted with mass flow rate ranging from 10 to 25 g/s, exit pressure from 60 to 150 bar, and inlet temperature of about 130-140 K. The maximum provided heat flux at channel bottom is 20 MW/m(2). Measurements of channel inlet and exit temperature and pressure, mass flow rate, and wall temperature at different channel locations have provided data useful for the evaluation of heat transfer and pressure loss. In particular, the channel surface roughness induced by the manufacturing process has been estimated and a peculiar Nusselt number correlation has been obtained. This correlation is suitable to describe the thermal behaviour of the rectangular cooling channel including both methane flow and wall. (C) 2016 Elsevier Ltd. All rights reserved.

Undesirable hot surface ignition of flammable liquids is one of the hazards in ground and air transportation vehicles, which primarily occurs in the engine compartment. In order to evaluate the safety and sustainability of candidate replacement fuels with respect to hot surface ignition, a baseline low lead fuel (Avgas 100 LL) and four experimental unleaded aviation fuels recommended for reciprocating aviation engines were considered. In addition, hot surface ignition properties of the gas turbine fuels Jet-A, JP-8, and JP-5 were measured. A test apparatus capable of providing reproducible data was designed and fabricated to experimentally investigate the hot surface ignition characteristics. A uniform surface temperature stainless steel plate simulating the wall of a typical exhaust manifold of an aircraft engine was used as the hot surface. Temperature uniformity of ±5°C was achieved on the stainless steel plate by virtue of its being bolted to a copper plate in which five automatically controlled 1000 W electrical cartridge heaters were inserted. A programmable syringe pump was used to dispense ~25 μL fuel drops onto the hot surface. Testing was performed in a quiescent environment with the exception of a mild upward flow created by an exhaust fan aiding the buoyant plume created by the hot plate. Ignition and flame propagation events were recorded using visible and midinfrared still and video imaging. The ignition and flame propagation events are transient and occur at randomly distributed locations on the hot surface. To characterize the ignition event statistically, the surface temperature leading to at least one ignition out of the number of drops and the surface temperature resulting in the ignition of all of the drops were recorded. The results of the experiment confirmed that the experimental variations in the drop size, drop velocity, plume characteristics, surface properties including temperature changes, and the nonlinear dependence of temperature of the chemical reaction rate lead to the probabilistic nature of the ignition event. The results of the experiment are of practical value in designing vehicular ignition and safety systems.

The use of the Methane as coolant in a regenerative liquid rocket engine (LRE) presents some difficulties since transcritical fluidynamics operating conditions occur in the cooling channels. Transcritical conditions cause large fluid properties variation that strongly influences the coolant performance. The HYPROB program is carried out by CIRA under contract by the Italian Ministry of Research with the main objective to improve National system and technology capabilities on liquid rocket engines for future space applications, with specific regard to LOx/LCH4 technology. Its main objective is to develop an test a LOX/LCH4 demonstrator. In order to match this objective a specific breadboard, the Methane Thermal Properties (MTP) breadboard has been manufactured and test. It is based on an electrical heating of a single representative cooling channel that has the aim to validate numerical methodologies and to improve the understanding of relevant physics of methane thermal properties in transcritical conditions. The experimental test campaign has been succsesfull performed at Maurice J. Zucrow Laboratories in Purdue University and the paper presents the main results.