Air-breathing Engines Research Articles

Experimental Investigation of Nozzle Effects on Thrust and Inlet Pressure of an Air-breathing Pulse Detonation Engine

Nozzle effects on thrust and inlet pressure of a multi-cycle air-breathing pulse detonation engine (APDE) are investigated experimentally. An APDE with 68 mm in diameter and 2 050 mm in length is operated using gasoline/air mixture. Straight nozzle, converging nozzle, converging-diverging nozzle and diverging nozzle are tested. The results show that thrust augmentation of converging-diverging nozzle, diverging nozzle or straight nozzle is better than that of converging nozzle on the whole. Thrust augmentation of straight nozzle is worse than those of converging-diverging nozzle and diverging nozzle. Thrust augmentations of diverging nozzle with larger expansion ratio and converging-diverging nozzle with larger throat area range from 20% to 40% on tested frequencies and are better than those of congeneric other nozzles respectively. Nozzle effects on inlet pressure are also researched. At each frequency it is indicated that filling pressures and average peak pressures of inlet with diverging nozzle and converging-diverging nozzle with large throat cross section area are higher than those with straight nozzle and converging nozzle. Pressures near thrust wall increase in an increase order from without nozzle, with diverging nozzle, straight nozzle and converging-diverging nozzle to converging nozzle.

How Much Compression Should a Scramjet Inlet Do?

The supersonic combustion ramjet, or scramjet, is the engine cycle most suitable for sustained hypersonic flight in the atmosphere. This paper examines a key decision in the design of the inlet or intake of these hypersonic airbreathing engines, namely, the level of compression to be performed. Too much compression can lead to onerous system level issues including the need for bleed or variable geometry, while too little compression can result in low cycle efficiency and poor combustion of fuel. An analysis of the important factors that affect the choice of scramjet inlet compression ratio has been performed for hydrogen-fueled scramjets at Mach 6, 8, 10, and 12. It was found that contrary to classical thermodynamic analyses, scramjet cycle efficiency reaches an optimum at a relatively low compression ratio between 50 and 100 for all Mach numbers. Practical constraints related to nonequilibrium flow effects, inlet starting, and boundary-layer separation were also shown to prompt a desire for low compression ratio. The lower limit on compression was found to be set by the need to complete the combustion reaction in the available engine length and is therefore dependent on engine scale. On the basis of these factors it is recommended that scramjet inlet compression ratio be set to the minimum that satisfies the robust combustion requirement, with the caveat that it not be below 50 in order to maintain high cycle efficiency. For typical wind-tunnel-scale engines, this results in a requirement for the inlet to compress airflow entering the combustor to a pressure of approximately 1/2 atm, regardless of the flight Mach number.

Recent Advances in Turbine Heat Transfer—With A View of Transition to Coal-Gas Based Systems

The performance goal of modern gas turbine engines, both land-base and air-breathing engines, can be achieved by increasing the turbine inlet temperature (TIT). The level of TIT in the near future can reach as high as 1700 °C for utility turbines and over 1900 °C for advanced military engines. To ensure the turbine airfoil component integrity operated under such a condition, advanced cooling capacity by both external and internal means was necessary to remove the excessive heat load from the turbine airfoil. This paper discusses state-of-the-art airfoil cooling technologies along with the associated thermal transport issues. Discussion is given based on five key regions on and around an airfoil: leading edge, main body, trailing edge, endwall, and near-tip. Potential implications and challenges of near-term developments in coal-gas based turbines on the cooling technologies are identified. A literature survey focusing primarily on the past 4–5years since the last International Heat Transfer Conference has also been performed.

Hidden Mass Boson

The evidences of the hidden mass boson existence are presented following the fruitful ideas of M. Planck and A. Einstein and using empirical data of modern physics. Within this article main parameters of this mass particle are predicted and its possible structure is analyzed. Moreover, the close system of nonlinear conservative equations and the spread system of Maxwell linear equations are written in the frame of phenomenological description of the hidden mass continuous medium. The displacement current, the Umov-Pointing vector and the physical vacuum polarization have been described adequately in our paper. We discuss some applications of our methodology for simulations of nature and technical device processes. In particular, numerical solutions for cosmic jets and air breathing engines are shown.

Effect of Local Flow Parameters on Skin-Friction Coefficient in Hypersonic Airbreathing Engines

Effects of Rex, Re , Tw=Taw, and Me on the local skin-friction coefficient (Cf ) in a hypersonic airbreathing combustor with various combustion modes were examined. Measurements were performed in the hypersonic airbreathing combustor with airflowMach number and unit Reynolds numbers at the entrance of combustor of 2.5 and 1:6 10 3:5 10 m , respectively. The ranges of total temperature, total pressure of airflow and fuel equivalence ratio were varied, i.e., 1000–2000 K, 1.0–1.8 MPa, and 0–0.4, respectively, to change Rex, Re , Tw=Taw, and Me. Direct skin-friction measurements were conducted. The state of flow was also surveyed in the streamwise locationwhere skin frictionwasmeasured to attain core-flowparameters, velocity profile, andmomentum thickness. When boundary-layer transition occurred due to combustion, evaluation ofCf by the vanDriestmethod using either by Rex or by Re was not effective. When there was separation with a long oblique shock train and a pseudo shock wave system in the precombustion region due to combustion, evaluation of Cf by the method using Rex was not effective because the origin of the turbulent boundary layer was changed in the downstream combustor.

Effect of Local Flow Parameters on Skin-Friction Coefficient in Hypersonic Airbreathing Engines

Effect of Local Flow Parameters on Skin-Friction Coefficient in Hypersonic Airbreathing Engines

Improved Air Turbo Rocket for Space Applications Application to Orbital Vehicles and Reentry

An Air Turbo Rocket (ATR) is a propulsion system which combines a turbo jet with a rocket engine. Currently it is being touted as a propulsion system for future missile systems, as these engines have a higher thrust density when compared to other air breathing engines. This paper explores the possibility of modifying the ATR for use in space application as well as during spacecraft re-entry. Such modified ATR’s could be used to power space vehicles up to the Low Earth Orbit (LEO) to dock with the International Space Station (ISS). In addition, thrust reversal techniques on the ATR systems could be used to improve the accuracy of Ballistic Missiles and hypersonic space planes upon Re-entry. Challenges faced would be in this type of air breathing engine would be operating at different atmospheric conditions. This paper will explore an ATR design, which will operate at different modes namely conventional mode, which will be used during below absolute ceiling, and the mission mode, which will be employed during flight in vacuum. Lastly, the reentry mode, which can be used for lessening the entry velocity of a vehicle to reduce the risks associated with reentry. The paper will try to emphasize the advantages of ATR as an affordable launch system for space shuttles and satellites with high maneuverability.

고성능 비행체 엔진을 위한 분출냉각의 연구동향

Transpiration cooling is the most effective cooling technique for the high-performance liquid rockets and air-breathing engines operating in aggressive environments with higher pressures and temperatures. When applying transpiration cooling, combustor liners and turbine blades/vanes are cooled by the coolant(air or fuel) passing through their porous walls and also the exit coolant acting as an insulating film. Practical implementation of the cooling technique has been hampered by the limitations of available porous materials. But advances in metal-joining techniques have led to the development of multi-laminate porous structures such as Lamilloy fabricated from several diffusion-bonded, etched metal thin sheets. And also with the availability of lightweight, ceramic matrix composites(CMC), transpiration cooling now seems to be a promising technique for high-performance engine cooling. This paper reviews recent research activities of transpiration cooling and its applications to gas turbines, liquid rockets, and the engines for hypersonic vehicles.

Numerical simulation of the operation process and thrust performance of an air-breathing pulse detonation engine in supersonic flight conditions

Multidimensional simulations of the unsteady gasdynamic flow in the duct of an air-breathing pulse detonation engine (ABPDE) operating on propane gas and the flow around it in supersonic flight at Mach numbers M of 3.0 and an altitude of 9.3 and 16 km are performed. It is shown that, at a length and diameter of the duct of 2.12 m and 83 mm, respectively, an ABPDE with an air intake and a nozzle can operate in a cyclic mode at a repetition frequency of 48 Hz, with a rapid deflagration-to-detonation transition (DDT) occurring at a distance of 5–6 combustion chamber diameters. To determine the thrust performance of the ABPDE in flight conditions, a series of working cycles were simulated with consideration given to the external flow around the engine. Calculations showed that the specific impulse of the ABPDE is approximately 1700 s. This value is much higher than the specific impulse typical of ramjet engines operating on conventional combustion (1200–1500 s) and substantially lower than the specific impulse obtained for the atmospheric conditions at sea level at zero flight velocity (∼2500 s).

Experimental study of an air-breathing pulse detonation engine ejector

Experimental studies were performed in order to improve the understanding of the performance of ejector driven by an air-breathing pulse detonation engine (PDE) with a convergent nozzle. This research utilized a gasoline-air PDE at four different operating frequencies of 8 Hz, 10 Hz, 12 Hz and 15 Hz. The performance of PDE-ejector was quantified by thrust measurements. The effects of single ejector length and axial location on thrust augmentation were investigated. It was found that the single ejector with L/D of 2 showed the best performance and the maximum thrust augmentation occurred at a downstream placement of +1 tube diameter. The performances of two-stage and three-stage ejectors were also investigated. The results indicated that both the overlap ratio and the flow area between two stages should not be too large. The performance of the two-stage ejector was not as sensitive as single-stage ejector to axial position in current conditions. The three-stage ejector behaved better than the two-stage ejector but worse than the single-stage ejector in this work. A maximum thrust augmentation of 1.8 was obtained with an L/D of 2 at a downstream placement of +1 position and 15 Hz operating frequency.

利用TDLAS技术开展吸气式发动机来流热试实验参数测量

利用TDLAS技术开展吸气式发动机来流热试实验参数测量

Numerical Analysis of Supersonic Combustion by Strut Flat Duct Length with S-A Turbulence Model

In this numeri cal study, of hydrogen has been presented with strut flat duct length. The combustor has a single fuel injection parallel to the main flow from the base. Finite rate chemistry model with S-A Model have been used for modeling of combustion. In this paper strut at 60 with flat duct length analyzed for without hydrogen injection, with hydrogen injection and hydrogen injection with combustion. The investigation of the shock system produced by a strut injector with subsequent mixing, ignition and flame stabilization showed that the characteristics of the ignition process is considerably different from the shock induced ignition which has been postulated in the past. The ignition is achieved through the irreversible total pressure loss in the wake of Strut due to flat duct length. Flat duct lengths allow the ignition delay to occur before the pressure and temperature is decreased by the expansion. The presence of a boundary layer in the chamber changes the shock system and the ignition conditions in the chamber considerably. Index Term—supersonic combustion, S-A model, finite rate, stagnation temperature, flame speed, Mach number. I. INTRODUCTION Superson ic is the key enabling technology for sustained hypersonic flights. In scramjet engines of current interest, the combustor length is typically of the order of 1 m, and the residence time of the mixture is of the order of Milliseconds. Due to the high flow speed in the chamber, problems arise in the mixing of the reactants, flame anchoring and stability and completion of within the limited combustor length. The flow field in the scramjet combustor is highly complex. It is shown that when the flight speed is low, the kinetic energy of the air is not enough to be used for the optimal compression. Further compression by machines is needed in order to obtain a higher efficiency. For example, a turbojet employs a turbine machine for further compression. When the flight speed is higher than a certain value, the air flow entering a combustor will remain to be after the optimal compression. With a further compression ( i. e. deceleration), the efficiency of the engine will decrease. Therefore the has to take place under the flow condition. This kind of air-breathing engine, which works under hypersonic flight condition, is called the ramjet (Scramjet). The term of supersonic combustion applied here means the

A two-level system of modeling air-breathing engine afterburners

Basic aspects of constructing multilevel, multidiscipline systems for modeling the airbreathing engine (ABE) units are presented. A two-level, two-discipline system of afterburner modeling is considered and the numerical modeling results and experimental data are given for comparison.

Analysis of efficiency of using hybrid propulsion for accelerating small-size rockets starting from the earth surface

The flight trajectories of a hypothetical small-size rocket starting from the Earth surface and driven by hybrid propulsion, which involves a solid-propellant rocket motor (SRM) and a ramjet-type air-breathing engine (ramjet), are calculated and analyzed. The basic varied parameter is the ratio of the fuel masses used for operation of various engine types with the total fuel mass being unchanged. The objective function is the flight range of the rocket under different initial conditions. Based on a model problem considered as an example, it is demonstrated that the flight range can be substantially increased by using hybrid propulsion (SRM and ramjet), as compared with the vehicle having identical mass and size characteristics and using the SRM only.

CFD analysis of the HyShot II scramjet combustor

The development of novel air-breathing engines such as supersonic combustion ramjets (scramjets) depends on the understanding of supersonic mixing, self-ignition and combustion. These aerothermochemical processes occur together in a scramjet engine and are notoriously difficult to understand. In the present study, we aim at analyzing the HyShot II scramjet combustor mounted in the High Enthalpy Shock Tunnel Göttingen (HEG) by using Reynolds Averaged Navier Stokes (RANS) and Large Eddy Simulation (LES) models with detailed and reduced chemistry. To account for the complicated flow in the HEG facility a zonal approach is adopted in which RANS is used to simulate the flow in the HEG nozzle and test-section, providing the necessary inflow boundary conditions for more detailed RANS and LES of the reacting flow in the HyShot combustor. Comparison of predicted wall pressures and heat fluxes with experimental data show good agreement, and in particular does the LES agree well with the experimental data. The LES results are used to elucidate the flow, mixing, self-ignition and subsequent combustion processes in the combustor. The combustor flow can be separated into the mixing zone, in which turbulent mixing from the jet-in-cross flow injectors dominates, the self-ignition zone, in which self-ignition rapidly takes place, and the turbulent combustion zone, located towards the end of the combustor, in which most of the heat release and volumetric expansion takes place. Self-ignition occurs at some distance downstream of the injectors, resulting in a distinct pressure rise further downstream due to the volumetric expansion as observed in the experiments. The jet penetration is about 30% of the combustor height and the combustion efficiency is found to be around 83%.

A STUDY ON FRACTURE BEHAVIORS OF CERAMIC USING SHOCK COMPRESSIVE WAVE

An experimental investigation was carried out to evaluate the fracture pressure and behaviors of ceramic materials for a dome port cover of an air breathing engine. The experiments were performed in a shock tube, which had a working section of 70 mm in diameter and a total length of 6 m in the shock tube. The response pressure transducers were used to measure expansion pressure and reflection pressure of working section near the end edge of the shock tube. Fractured specimens collected from the dump tank were investigated for the fracture phenomenon of ceramics after testing. The fracture pressure increases as a specimen thickness increases, and as the specimen diameter decreases, respectively. The driven gas pressures P 2 obtained from theory and experiment agree well. The reflection wave pressures P 5 agree well comparatively with both results of equation and experiment at low pressure of P 4, but they do not agree well at high pressure of P 4. The fracture phenomena of the plate and dome specimens at the same diameter are broken into small particles in various thicknesses, but the fracture phenomenon of the plate specimen is not broken into the particles as specimen diameter decreases.

A study on fracture phenomena of glass filled ceramics using shock tube

AbstractFracture phenomena of plate and dome shaped specimens using a shock tube for glass‐filled ceramics were investigated. Glass‐filled ceramics have been considered a promising candidate material for the dome port cover of an air breathing engine. This part of the air breathing engine has an important role in separating solid and liquid fuel, and needs a frangible characteristic that the fracture of a part should not affect the internal components of combustion. The objective of this study is to evaluate the fracture pressures for various thicknesses and diameters of the shock impact area. Also, fracture phenomena of separated membrane using a shock tube are compared with a numerical method. The experimental apparatus consists of a driver, a driven section and a dump tank. The used material is a glass‐filled ceramic made from Corning Company. Specimens are used with 3, 4.5 and 6 mm thicknesses, respectively. It is expected that the results obtained from this study can be used in the basic data for the dome port cover design of an air breathing engine.

Mixing characteristics of pulsed air-assist liquid jet into an internal subsonic cross-flow

Penetration depth, spray dispersion angle, droplet sizes in breakup processes and atomization processes are very important parameters in combustor of air-breathing engine. These processes will enhance air/fuel mixing inside the combustor. Experimental results from the pulsed air-assist liquid jet injected into a cross-flow are investigated. And experiments were conducted to a range of cross-flow velocities from 42∼136 m/s. Air is injected with 0∼300kPa, with air-assist pulsation frequency of 0∼20Hz. Pulsation frequency was modulated by solenoid valve. Phase Doppler Particle Analyzer(PDPA) was utilized to quantitatively measuring droplet characteristics. High-speed CCD camera was used to obtain injected spray structure. Pulsed air-assist liquid jet will offer rapid mixing and good liquid jet penetration. Air-assist makes a very fine droplet which generated mist-like spray. Pulsed air-assist liquid jet will introduce additional supplementary turbulent mixing and control of penetration depth into a cross-flow field. The results show that pulsation frequency has an effect on penetration, transverse velocities and droplet sizes. The experimental data generated in these studies are used for a development of active control strategies to optimize the liquid jet penetration in subsonic cross-flow conditions and predict combustion low frequency instability.

Rotordynamic Force Coefficients of a Hybrid Brush Seal: Measurements and Predictions

Brush seals effectively control leakage in air breathing engines, albeit only applied for relatively low-pressure differentials. Hybrid brush seals (HBS) are an alternative to resolve poor reliability resulting from bristle tip wear while also allowing for reverse shaft rotation operations. A HBS incorporates pads contacting the shaft on assembly; and which under rotor spinning, lift off due to the generation of a hydrodynamic pressure. The ensuing gas film prevents intermittent contact, reducing wear, and thermal distortions. This paper presents rotordynamic measurements conducted on a test rig for evaluation of HBS technology. Single frequency shaker loads are exerted on a test rotor holding a hybrid brush seal, and measurements of rotor displacements follow for operating conditions with increasing gas supply pressures and two rotor speeds. A frequency domain identification method delivers the test system stiffness and damping coefficients. The HBS stiffness coefficients are not affected by rotor speed though the seal viscous damping shows a strong frequency dependency. The identified HBS direct stiffness decreases ∼15% as the supply/discharge pressure increases Pr=1.7–2.4. The HBS cross-coupled stiffnesses are insignificant, at least one order of magnitude smaller than the direct stiffnesses. A structural loss factor (γ) and dry-friction coefficient (μ) represent the energy dissipated in a HBS by the bristle-to-bristle and bristle-to-pad interactions. Predictions of HBS stiffness and damping coefficients correlate well with the test derived parameters. Both model predictions and test results show the dramatic reduction in the seal equivalent viscous damping coefficients as the excitation whirl frequency increases.

충격파관을 이용한 세라믹 돔의 파괴 특성에 관한 연구

Fracture characteristics for plate and dome shapes of glass filled ceramics using shock tube were carried out. Glass filled ceramics have been considered as a promising candidate material for the dome port cover of air breathing engine. This part of the air breathing engine has an important role as separated membrane between combustion and external air, and needs the frangible characteristics that the particles of fractured glass filled ceramics should not affect the internal components of combustion. The objectives of this study are to evaluate the fracture pressures for various thicknesses and diameters of shock impact area. Also fracture phenomena of separated membrane using a shock tube compare with analytical method. The experimental apparatus consists of a driver, a driven section and a dump tank. The used material is glass filled ceramic made from Corning company. Specimens have the thickness of 3, 4.5 and 6㎜. It is expected that the results obtained from this study can be used in the basic data for the dome port cover design of an air breathing engine.