High Temperature Combustion Gas Research Articles

Study on transient thermal state of steel piston in diesel engine under cold start condition

The thermal load issue of pistons remains a topic of active research, especially the transient thermal state of steel pistons in diesel engines under cold start conditions has not been studied in depth. In this paper, a new steel piston power assembly was designed and developed on the basis of a small-bore diesel engine. A real-time temperature test system for the steel piston temperature field was developed to address the problem of the inability to obtain real-time temperature data. The test system was applied to the steel piston to investigate the transient variations in temperature field, thermal stress, and thermal deformation at different starting temperatures. The results show that the temperature field of steel pistons exhibits exponential growth in the early stage of cold start and then tends to stabilize after about 44 s. The lower the cold start temperature and the closer the monitoring points are to the high-temperature combustion gas, the faster the rate of temperature rise. The thermal stress of the steel piston initially increases rapidly, then decreases, and subsequently increases again until reaching a stable level. The lower the starting temperature, the longer the thermal stress inflection point of the monitoring point will be. The variation of thermal deformation of the steel piston follows the same trend as that of the temperature field.

Oxidation Behavior of Atmospheric Plasma Sprayed NiAl Coating at High Pressure Oxygen Atmosphere

The high-temperature and high-pressure oxygen-rich combustion gases often cause the high temperature gas conduit in rocket engine to suffer severe oxidation, resulting in a reduced service life. The present study deposited a dense NiAl coating on GH4202 superalloy using atmospheric plasma spraying and investigated its oxidation behavior in a high temperature and high-pressure oxygen atmosphere. The results showed that the NiAl coating presented excellent bonding with the superalloy substrate throughout tests. The uncoated GH4202 substrate was severely oxidized during high pressure oxygen atmosphere. It was found that only a continuous very thin Al2O3 oxide scale evolved on the NiAl coating surface and no oxidation was observed at the interface between the NiAl coating and GH4202 substrate. Moreover, no significant inner oxidation occurred to NiAl coating. Therefore, the present APS NiAl coating can provide effective protection of superalloy from oxidation at high pressure oxygen combustion gas atmosphere.

Study on Ground Experimental Method of Stagnation Point Large Heat Flux of Typical Sharp Wedge Leading Edge Structure

In this paper, aimed at the problem of large temperature gradient thermal testing with the typical sharp wedge leading edge structure of a hypersonic vehicle, a subsonic high-temperature combustion gas heating (SHCH) test device is used to conduct a series of experiments on the heat flux simulation ability of subsonic high-temperature combustion gas in the stagnation point region. Firstly, for a hypersonic vehicle with a flying height of 24 km and Mach number range of 4~6.5, the stagnation point heat flux in the head area is obtained by numerical calculation of a typical leading edge structure, which is used as the experimental target of the thermal structure test. Secondly, an experimental specimen with a Gardon heat flux meter is designed with the same shape and size as the specimen in the numerical simulations to prepare for the subsequent SHCH test. Thirdly, a method to determine the combustion gas temperature based on a Kriging surrogate model is proposed. CFD numerical simulation is conducted using the SHCH test model, and the numerical calculation results are used as the training dataset. The Kriging surrogate model is used to establish an approximate fitting relationship between the stagnation point heat flux and experimental parameters under SHCH conditions. The corresponding combustion gas temperature values are found, respectively, with the hypersonic aerodynamic heat flux at Mach 5.0~5.4 as the target value. Finally, stagnation point heat flux testing of low-speed and high-temperature combustion gas is performed at different combustion gas temperatures. The experimental and target values obtained from hypersonic aerodynamic thermal simulations are compared and analyzed to verify the heating capacity of SHCH and the feasibility of hypersonic aerothermal simulation testing.

Flame-assisted plasma modulation to improve the raw signal quality for laser-induced breakdown spectroscopy

Laser-induced breakdown spectroscopy (LIBS) is a promising multi-elemental analysis technique and has the advantage of rapid measurement with minimal sample preparation. However, due to the short lifetime of laser-induced plasma and the very small ablation mass, the detection sensitivity of LIBS is low and the signal uncertainty is relatively high. These issues severely hinder the large-scale application and commercialization of LIBS. A flame-assisted plasma modulation method is used to reduce signal uncertainty and simultaneously enhance signal intensity, based on the uncertainty generation mechanism of LIBS. During the plasma evolution, the ambient high-temperature combustion gases can provide an environment with high local sound speed, which reduces the crash between the lower part of plasma and the back-pressed frontier material. Therefore, the generated plasma morphology is stable and the signal uncertainty is lowered. Moreover, a hotter surrounding gas environment can benefit sample ablation mass and reduce heat loss of the plasma, therefore leading to higher signal intensity. The flame-assisted plasma modulation method is tested on brass samples. It reduces the relative standard deviation (RSD) from 10% to 6% approximately and shows an enhancement factor of about 2 for Cu spectral lines, which shows that the proposed method is effective to improve the raw signal quality for LIBS. Detailed analysis of the mechanisms for uncertainty reduction and intensity enhancement was further investigated for a deeper understanding.

Thermal performance evaluation and analysis of helium heat exchanger for cryogenic propellant launch vehicle

A launch vehicle that operates using cryogenic liquid propellants will require a pressurization process for a propellant tank. Usually, helium stored at cryogenic temperatures passes through the engine, is heated, and pressurizes the tank. Low-temperature helium is heated in a heat exchanger using high-temperature combustion gas from an engine. For the accurate performance prediction and weight optimization of the launch vehicle, the performance of the heat exchanger should be evaluated before launching. This study presents an evaluation method for predicting the performance of a heat exchanger in a launch vehicle. The heat exchanger performance was measured at a constant helium flow rate under actual engine operating conditions. A numerical analysis model for predicting the heat exchanger performance was developed and compared to the test results. The tests were conducted at different inlet temperatures. The test results and analysis confirmed that the effectiveness increased with the inlet temperature. Further analyses revealed that a change in the heat transfer coefficient, as well as the inlet temperature, changed the effectiveness by 6%. To accurately predict the performance of the launch vehicle, it was necessary to accurately predict the change in the heat transfer coefficient with temperature. A design method for the heat exchanger of a launch vehicle was also discussed. The heat-transfer area and weight should be optimized under the required conditions.

Effect of Spray Method and Sprayed Material on Creep Behavior of MCrAlY Coating

The blade of gas turbines used in combined cycle power plant are exposed to a high-temperature combustion gas. In order to protect it from such a harsh environment, a thermal barrier coating (TBC) is applied to the surface of the blade. TBC consists of a top coat (TC), which is a thermal barrier layer, and a bond coat (BC), which increases the adhesion strength between the TC and the substrate. Thermal stress is caused by the thermal expansion mismatch between the TC and substrate, which leads to delaminating in TC. In order to reduce such thermal stress in TC, creep deformability of BC is essential. In this study, we attempted to evaluate the creep deformation property of BC by creep indentation test method developed by our group. CoNiCrAlY and CoCrAlY are selected as BC materials and APS and HVOF are also as thermal spraying methods, and the creep indentation tests were conducted at various temperatures. It was found that material differences had little effect on creep deformation properties, and that creep properties were strongly affected by the thermal spraying method even if the material systems had the same chemical composition.

Nondestructive evaluation method of thermal barrier coatings for gas turbine by rebound type hardness test machine

Gas turbine (GT) blades are exposed to high-temperature combustion gas that exceed the thermal resistance temperature of Ni-based superalloy. Therefore, the blade surface is coated by thermal barrier coating (TBC), which consist of two layers: a ceramic top coat (TC) and a bond coat (BC). The cyclic thermal loading in GT operation causes damage to the TC layer, such as cracking and delamination. Therefore, GT blades utilized in thermal power generation are periodically recoated to prevent such damage. However, recoating costs constantly, and it is required to extend the recoating period to reduce it. For this reason, advanced technology that enable nondestructive and simple deterioration diagnosis of damage to TBCs has been actively developed globally. In this study, we focused on the rebound-type hardness test, in which microspheres are impacted on the TBC surface, and attempted to relate the hardness value (Leeb hardness) to the damage of the TBC.

Experimental and numerical study of heat and mass transfer between high temperature supersonic gas flow and liquid water in a cylindrical duct

The mixing process of a high temperature combustion gas and liquid water conveying large amounts of heat and mass transfer in a very short period of time is investigated experimentally and numerically. A solid propellant produces hot gas into a cylindrical duct where water is supplied through water injectors, facilitating instantaneous mixing accompanied by evaporation and condensation along the duct. Gas generator fire tests are carried out to confirm performance. A cylindrical duct is subsequently connected to the rear end of the gas generator with the liquid water supply system, and experiments are performed for three different amounts of water being injected into the supersonic hot gas. A numerical scheme for a homogenous mixture model with Lee's mass transfer model is applied, and comparisons are made with the experiments. Next, both the experimental and numerical results are fairly compared. The empirical coefficients of evaporation rate for the Hertz-Knudsen-Scharage (HKS) and wall film model (WFM) equations are proposed.

A novel model of liquid film flow and evaporation for thermal protection to a chamber with high temperature and high shear force

Thin liquid film evaporation is a prevailing cooling technique for thermal protection to the inner wall of a chamber with high-intensity combustion. A novel model is developed to investigate the flow and evaporation of liquid film in a rocket combustion chamber with high temperature and high shear force, in which shear stress at the interface between liquid film and high temperature combustion gas with high velocity, capillary pressure of liquid and disjoining pressure associated with liquid film evaporation are taken into account. A numerical algorithm of the backward solution is deliberately adopted to solve this model so as to eliminate the entrance effect caused by different inlet conditions of liquid film. For given temperature and hot gas rate in a chamber, a stably evolution regime of liquid film with the thickness from thick to thin is depicted corresponding to the total process of being sheared into film and being evaporated to dryness. Results indicate that the three forces mentioned above play critical parts in flowing and evaporating of liquid film during its different evolution regimes. It was shown that shear stress and capillary pressure have a positive effect on the flow and evaporation of film, as well as a negative effect from the disjoining pressure. The effects of these forces are different dependent on different initial average thickness of film, and the distribution profile of these forces are totally illustrated in detailed. The prediction of liquid film length, from being injected to being evaporated, by the proposed model is in a good agreement with available experimental data. All findings are expected to provide a fundamental guideline for the design and optimization of efficient thermal protect to the combustion chamber of space engine.

Mechanical properties degradation and microstructure evolution of 2D-SiCf / SiC composites in combustion gas environment

Based on the turbine high-temperature combustion gas simulation test platform, the long-term combustion gas environment exposure test of the 2D plain woven SiCf/BN/SiC composites under two combustion conditions was carried out. Uniaxial tensile test, fracture morphology characterization and non-destructive testing analysis revealed the degradation and microstructure evolution of composites after exposure to combustion gas environment. The results show that the degradation of 2D-SiCf/SiC composites after exposure to combustion gas environment is manifested as a decrease in static toughness, and the interphase transition is the mesoscopic cause of the decrease in static toughness of the composite.

Combustion Characteristics of Multi-Element Swirl Coaxial Jet Injectors under Varying Momentum Ratios

The combustion characteristics of a staged combustion cycle engine with an oxidizer-rich preburner were experimentally studied at different momentum ratios of multi-element injectors. Propellants were simultaneously supplied as a liquid–liquid–liquid system, and an injector was designed in which a swirl coaxial jet is sprayed. The injector burned the propellants in the inner chamber which had a temperature greater than 2000 K. To cool the combustion gas, a liquid oxidizer was supplied to the cooling channel outside the injector. To prevent the turbine blades from melting, the temperature of the combustion gas was maintained below 700 K. To confirm the combustion characteristics at different momentum ratios of the high-temperature combustion gas inside the injector and the low-temperature liquid oxidizer outside the injector, three types of injectors were designed and manufactured with different momentum ratios: MR 3.0, MR 3.3, and MR 3.7. In this study, the results of the combustion test for each type were compared for 30 s. For ORPB-A, a combustion pressure of 18.5 MPaA, fuel mass flow rate of 0.26 kg/s, oxidizer mass flow rate of 15.3 kg/s, and turbine inlet temperature of 686 K were obtained in the combustion stability period of 29.0–29.5 s. The combustion efficiency was 98% for MR 3.0 (ORPB-A), which was superior to that for other momentum ratios. In addition, during the combustion test for MR 3.0, the fluctuations in the characteristic velocity, combustion pressure, and propellant mass flow rate were low, indicating that combustion was stable. The three types of combustion instability were all less than 0.8%, thus confirming that the combustion stability was excellent.

Numerical and experimental study on high-speed hydrogen–oxygen combustion gas flow and aerodynamic heating characteristics

The need to increase the payload capacity of the rockets motivates the development of high-power rocket engines. For a chemical propulsion system, this results in an increasing thermal load on the structure, especially the combustion chamber and nozzle must be able to withstand the extreme thermal load caused by high-temperature and high-pressure combustion gas. In order to protect the structure from the effect of increasing heat flux, it is necessary to counteract such effect with more advanced thermal management technology. This requires us to accurately predict the aerodynamic heating of the structure by high-temperature and high-speed combustion gas. In this study, a high-temperature combustion gas tunnel developed in the laboratory is used to produce high-speed combustion gas. Combined with the results of numerical calculation, the flow and aerodynamic heating characteristics of air and hydrogen–oxygen combustion gas under the same total temperature and pressure are analyzed and compared. The comparison revealed that the combustion gas flow in the nozzle has higher static temperature, velocity, and smaller Mach number. When the combustion gas flows around the sphere, the shock standoff distance and stagnation pressure are smaller than those of air, and the wall heat flux is much larger than that of air. The active chemical reaction in the combustion gas makes the aerodynamic heating of the structure more severe. Finally, through the analysis of a large amount of data, a semi-empirical formula for the heat flux of the stagnation point heated by a high-speed hydrogen and oxygen equivalent ratio combustion gas is obtained.

Numerical Studies on Heat Transfer of Supersonic Combustion-Gas Jet Impingement on Perpendicular Flat Plate

This paper numerically investigates the heat transfer characteristics of supersonic high-temperature combustion-gas jet impinging to a perpendicular flat plate. First, the numerical method is validated using an experimental case and meanwhile different turbulence models are evaluated for heat transfer prediction. Then, the effects of multi-species, nozzle-to-plate distance, and nozzle pressure ratio on wall heat transfer are studied through a series of numerical experiments. Comparison of different turbulence models indicates that the variations of turbulent viscosity across the detached normal shock wave differ among the studied three models. Furthermore, if the real multi-species of combustion-gas are mixed as a single species with the same thermal-physical properties, the pressure and heat flux distributions are identical with the results of combustion-gas. Finally, when there is no stagnation bubble in the jet flow field, the heat flux on the flat plate produces a local minimum at the stagnation point for short nozzle-to-plate distance, while a local maximum appears at the stagnation point for large nozzle-to-plate distance. When the stagnation bubble exists, the other maximum of heat flux is produced at the edge of bubble.

Fatigue failure and launch safety assessment of gun barrel

Hundreds or thousands of projectiles have been fired through a gun barrel. Catastrophic fatigue failure of gun barrel due to cyclic loadings by the stresses induced by the combination of high pressure and temperature combustion gas, and projectile with the rotating band interaction with the bore surface, has not yet been understood. In this paper, laboratory hydraulic fatigue tests on simulated gun barrel were performed on MTS Land. The experimental results show that there is an inherent relation between crack propagation in thick-walled tube and the external surface hoop and axial strains. This fact suggests that fatigue failure of gun barrel should be likely prevented by monitoring the external surface strains of the gun barrel and stopping continued firing before a major crack growing up to the critical size.

Effects of Mg/PTFE pyrotechnic compositions on reignition characteristics of base bleed propellants and heating mechanism

The effects of magnesium/polytetrafluoroethylene (Mg/PTFE) pyrotechnic compositions on the coupled flow field and reignition mechanism are important aspects governing the performance and range of base bleed projectiles (BBPs). Owing to a decrease in pressure and temperature when the BBP leaves the muzzle, rapid depressurization occurs, which extinguishes the base bleed propellant. The Mg/PTFE pyrotechnic composition pressed in the igniter of the base bleed unit (BBU) provides additional energy to the BBU via a chemical reaction. Thus, the extinguished base bleed propellant is reignited under the effect of high-temperature combustion gas jets from the igniter. In this study, a numerical analysis is conducted to evaluate the effects of PTFE and Mg granularity as well as Mg/PTFE pyrotechnic compositions. Owing to the rapid depressurization, the temperature and pressure was found to decrease for different Mg/PTFE pyrotechnic compositions. However, the depressurization time increased as the PTFE granularity increased, the Mg granularity decreased, and the Mg content increased. When the pressure in the combustion chamber of the BBU decreased to the atmospheric pressure, the combustion gas jets from the igniter expand upstream (rather than downstream). However, these combustion gas jets exhibit different axial and radial expansion characteristics depending on the pyrotechnic compositions used. The results show that the reignition delay time, td, of the base bleed propellant was 377.608, 94.27, 387.243, 523.966, and 221.094 ms for cases A–E, respectively. Therefore, it was concluded that the Mg/PTFE pyrotechnic composition of case B was the most beneficial for the reignition of the base bleed propellant, with the earliest addition of energy and mass to the BBP.

Evolution of microstructure and thermal conductivity of multifunctional environmental barrier coating systems

Environmental barrier coating (EBC) systems are applied to the surface of silicon-based composites exposed to high temperature combustion gas flow paths in gas turbine engines. They reduce the rate of composite oxidation, its volatilization by reactions with water vapor, and the temperature of the composite as their thermal conductivity decreases. Current EBC systems consist of a silicon bond coat covered by an ytterbium disilicate (Yb2O3⋅2SiO2: YbDS) permeation resistant and low silica activity barrier. When applied by atmospheric-plasma spray deposition, this outer layer contains 10–15% of a secondary ytterbium monosilicate (Yb2O3⋅3SiO2: YbMS) phase. YbDS has a coefficient of thermal expansion (CTE) of 4–5×10−6 °C−1, similar to those of the silicon and the silicon-based composite, but a relatively high thermal conductivity of 5–7 W m−1 K−1. YbMS has a higher steam volatility resistance than YbDS, and it has a much lower thermal conductivity (∼2–2.5 W m−1 K−1) at ambient temperature compared to YbDS, but its higher, highly anisotropic CTE (3–11×10−6 °C−1) results in channel cracking which reduces environmental protection. Here, we use a combination of scanning electron beam and laser-based thermoreflectance methods to spatially map the distribution of the silicate phases and thermal conductivity at ambient temperature in a Si-ytterbium disilicate EBC system exposed to thermal cycling in water vapor. We show that during thermal cycling, diffusion of silica from the thermally grown oxide on the Si bond coat surface to nearby YbMS regions transforms this phase to YbDS, thereby reducing the risk of thermomechanical coating failure but decreasing its effective thermal resistance. We also show that as silica is volatilized at the water vapor–ytterbium silicate interface, YbDS is transformed to YbMS, restoring some of the thermal protection of the coating system lost by its reduction in thickness and the YbMS to YbDS transformation near the bond coat.

Thermal and flow characteristics of a cylindrical superheater with circular fins

In the present study, a cylindrical superheater for producing high-temperature superheated steam was installed in a furnace, and its heat flow characteristics were investigated. The effect of applying circular fins of various curvature angles to the inside of the superheater was analyzed to improve heat transfer performance. The steam flowing into the cylindrical superheater showed a swirling flow along the wall surface as it fell down to the bottom of the superheater, and received heat from the high-temperature combustion gas. When the numerical analysis data was compared to experimental results for verification, the hot gas temperature and superheated steam temperature were found to be similar, within 5% and 8.3%, respectively. The temperature of the steam discharged through the superheater outlet was higher when fins were applied to the superheater, compared to the case without the fins, because the heat transfer area was increased and a secondary flow was generated between the fins. As a result, compared to the superheater without fins, the pressure drop increased by 2.2% with flat fins, 2.8% with 30° fins, and 3.8% with 60° fins. The discharge steam temperature increased by 0.5% with flat fins, 1.8% with 30° fins, and 2.4% with 60° fins. Also, Nusselt numbers increased by 19.6% for the flat fins, 20.3% at 30° fins, and 21.5% at 60° fins.

Damage Evaluation of TBC by Rapid Thermal Cycling Test Utilizing a Laser Irradiation

Thermal barrier coating (TBC) is deposited onto the gas turbine blade surface in order to protect the substrate from high-temperature combustion gas. Cracks and delamination of the ceramic coating which come from high heat flux loading are serious problem in TBC. In this study, the rapid thermal cycling device utilizing laser irradiation was developed. It was then investigated how the damage progresses in the ceramic coating exposed to cyclic rapid thermal loading. As a result, a sintering layer was formed in the surface of the ceramic coating, although such phenomenon was not recognized in TBC sample tested by the conventional thermal cycling test using an electric furnace. It was also revealed from the cross-sectional observation that the vertical crack was initiated at the surface of TBC and propagated into sintering layer. Finally, mechanical factors of those damages from finite element analysis using the TBC model including sintering progress was discussed

Coupling characteristics of combustion-gas flows generated by two energetic materials in base bleed unit under rapid depressurization

The present study is devoted to investigate the physical characteristics of the coupling combusting-gas flows generated by gun propellant and pyrotechnics along with the effect of initial pressure on the extension behavior of the igniter combustion jet under the rapid depressurization process that the BBU (base bleed unit) undergoes when a BBP (base bleed projectile) is flying out of the muzzle. A semi-closed bomb is designed to reproduce the depressurization process. The near-nozzle pattern evolution of the exhaust plume is observed by using an HSC (high-speed camera). Based on the experiment, a two-dimensional axisymmetric model is established including the high-resolution upwind scheme AUSM+ and the finite-rate chemistry model (no turbulence-chemistry interaction). The cell-centered FVM (finite volume method) is employed to simulate the coupling characteristics of the two high-temperature combustion gases under sharp depressurization. The results show that as the pressure drops rapidly, the exhaust plume gradually transforms from highly underexpanded supersonic flow into a subsonic flow, in where the wave structure converts from Mach reflection to regular reflection. The periodic shock diamond and rhombic flame string are formed, and finally the flame becomes continuous. In the subsonic flow, the radial heat convection and diffusion downstream of the flame in BBU are stronger than upstream. During the depressurization process, the influence of initial NPR (nozzle pressure ratio) becomes more significant. As the initial NPR increases, the igniter flame squashes downstream more slowly, and the heat loss of the gas expansion accelerates, thus the radial heat convection and diffusion in BBU decrease.

Compact, fiber-coupled, single-ended laser-absorption-spectroscopy sensors for high-temperature environments.

The design and demonstration of a compact single-ended laser-absorption-spectroscopy sensor for measuring temperature and H2O in high-temperature combustion gases is presented. The primary novelty of this work lies in the design, demonstration, and evaluation of a sensor architecture that uses a single lens to provide single-ended, alignment-free (after initial assembly) measurements of gas properties in a combustor without windows. We demonstrate that the sensor is capable of sustaining operation at temperatures up to at least 625K and is capable of withstanding direct exposure to high-temperature (≈1000 K) flame gases for long durations (at least 30min) without compromising measurement quality. The sensor employs a fiber bundle and a 6mm diameter antireflection-coated lens mounted in a 1/8'' NPT-threaded stainless-steel body to collect laser light that is backscattered off native surfaces. Distributed-feedback tunable diode lasers (TDLs) with a wavelength near 1392nm and 1343nm were used to interrogate well-characterized H2O absorption transitions using wavelength-modulation-spectroscopy techniques. The sensor was demonstrated with measurements of gas temperature and H2O mole fraction in a propane-air burner with a measurement bandwidth up to 25kHz. In addition, this work presents an improved wavelength-modulation spectroscopy spectral-fitting technique that reduces computational time by a factor of 100 compared to previously developed techniques.