Showing papers on "Afterburner published in 2013"

Exergetic analysis of an aircraft turbojet engine with an afterburner

Abstract: An exergy analysis is reported of a J85-GE-21 turbojet engine and its components for two altitudes: sea level and 11,000 meters. The turbojet engine with afterburning operates on the Brayton cycle and includes six main parts: diffuser, compressor, combustion chamber, turbine, afterburner and nozzle. Aircraft data are utilized in the analysis with simulation data. The highest component exergy efficiency at sea level is observed to be for the compressor, at 96.7%, followed by the nozzle and turbine with exergy efficiencies of 93.7 and 92.3%, respectively. At both considered heights, reducing of engine intake air speed leads to a reduction in the exergy efficiencies of all engine components and overall engine. The exergy efficiency of the turbojet engine is found to decrease by 0.45% for every 1°C increase in inlet air temperature.

Gas turbine and gas turbine afterburner

Abstract: A gas turbine afterburner includes a gutter electrode that helps to hold an afterburner flame. A charge source applies a majority charge to be carried by a turbine exhaust gas. Electrical attraction between the majority charge and the gutter electrode helps to hold the afterburner flame.

Flameholding in Converging and Turning Channels over Cavities with Periodic Port Injection

Abstract: The deliberate continuation of the combustion in the turbine passages of a gas-turbine engine has the potential to increase the efficiency and the specific thrust or power of current gas-turbine engines. This concept, known as a turbine burner, introduces certain challenges concerning the injection, mixing, ignition, and burning of fuel within a short residence time in a turbine passage characterized by large three-dimensional accelerations. Here, the injection of the fuel into a cavity adjacent to the modeled turbine passage is examined, which creates a low-speed zone for mixing and flameholding. The turbine passage is modeled as a converging and curving channel flow of high-temperature vitiated air adjacent to a cavity. To give a broader understanding of the cavity channel-flow coupling, both constant-area and converging channels with both straight and curving centerlines are modeled. Three-dimensional unsteady calculations with periodic port injection are performed, examining the effects of channel con...

Variable-cycle air turbine combined engine of rocket

Abstract: The invention discloses a variable-cycle air turbine combined engine of a rocket. The variable-cycle air turbine combined engine of the rocket comprises an axial compressor, a shaft, a main combustion chamber, an air duct, a turbine, an afterburner and an exhaust nozzle, wherein the axial compressor is arranged on the side, with an air inlet, of the main combustion chamber, the turbine is arranged on the side, with an air outlet, of the main combustion chamber, the axial compressor is connected with the turbine through the shaft, the air duct is arranged on the outer side of the main combustion chamber, and is used for communicating the axial compressor with the afterburner, the afterburner is arranged on the exhaust side of the turbine, and the exhaust nozzle is communicated with the afterburner. The variable-cycle air turbine combined engine of the rocket further comprises a fuel gas generator and a second turbine, wherein the fuel gas generator is arranged between the air duct and the main combustion chamber, the second turbine is arranged at the blade tip position of the turbine, and is connected with the axial compressor through the shaft, and an outlet of the fuel gas generator exactly faces the second turbine. The variable-cycle air turbine combined engine of the rocket provides a power system solution scheme for a spacecraft, wherein according to the power system solution scheme, the spacecraft can carry out the long-time economical cruise, and can carry out high-speed maneuver in a short time in a high comprehensive performance mode, and the problem that an existing aviation turbine engine can not carry out self-starting can be solved.

Development of combustion driven heating system for semi-free jet test rig

Abstract: National Aerospace Laboratories, CSIR has recently established the Aero-propulsion Mach 3.6 Tunnel for the experimental research on 'High Speed Combustors. It is a blow down type test facility and simulates the scramjet inlet conditions of Mach number and temperature corresponding to the flight Mach number of 6 and altitude of 25 km. The flow conditions obtained at the test section are – Mach number 3.6, total temperature 1700 K, pressure 18 bar, mass flow rates of 25 kg/s with 10 seconds of test duration. The combustion driven heating method with upstream injection of replenishment oxygen has been adapted. The heating system consists of two combustors namely Pre-Heater -01 and Pre-Heater -02. Pre-heater -01 is an aero engine ‘can type’ combustor which heats air from 300 K to 800 K. The liners and fuel injectors from time expired R-11 aero engine were modified to suit to the requirements. Pre-Heater-02 is an ‘afterburner type’ combustion chamber developed for the facility, which heats the air from 800 K to 1700 K. This paper discusses about the development of the pre-heaters and its performances

Design and Evaluation of a Turbojet Exhaust Simulator, Utilizing a Solid-Propellant Rocket Motor, for Use in Free-Flight Aerodynamic Research Models

Abstract: A method has been developed for modifying a rocket motor so that its exhaust characteristics simulate those of a turbojet engine. The analysis necessary to the design is presented along with tests from which the designs are evaluated. Simulation was found to be best if the exhaust characteristics to be duplicated were those of a turbojet engine at high altitudes and with the afterburner operative.

Numerical Investigation of Aero Engine Afterburner Fuel Manifold

Abstract: In this paper, Method of CFD is used to simulate afterburner fuel manifold for aero engine, and investigate the relationship between the boundary layer of nozzle hole and mass flow. At the same time the numerical simulation of permitted maximum and minimum of nozzle hole was carried out and mass flow for calculation was also compared with the design value. The calculation results show the effect of machining quality of nozzle holes on mass flow of afterburner fuel manifold, for which the diameter of nozzle hole is so small that the effects of machining precision and boundary layer thickness on mass flow is very large. Therefore the jet fuel must be machined according to tolerance level I so as to ensure that the mass flow meets the design requirements. Investigation of boundary layer of small hole shows that only the boundary layer factors are fully considered, the numerical calculation accuracy may be enhanced.

Flame tube head of afterburner with flame stabiliser of variable geometry

Abstract: FIELD: engines and pumpsSUBSTANCE: flame tube head of afterburner of gas-turbine engine includes radial cooled shutters Cooled shutters have aerodynamic streamlined shaped parts arranged immediately in flow and have the possibility of being turned about their axis; high-drag body in the form of V-shaped stabiliser is formed at synchronous rotation in opposite directions to a certain angle and at joining of a pair of neighbouring radial cooled shutters Aerodynamically streamlined shaped parts arranged immediately in the flow smoothly pass to cylindrical axes fixed in housing of flame tube head Turning device in the form of a rotating ring is located outside afterburner and has the possibility of being blown with ambient air during flight Various internal cavities and radial channels are made inside radial cooled shutters Internal cavities are connected to flow part by means of cooling openings located on inner surface of V-shaped stabiliser Radial channels are connected to flow part by means of injectors for spray of afterburner fuel, which are located on outer surface of V-shaped stabiliser at certain angle to the flowEFFECT: invention allows reducing total pressure losses in afterburner of aircraft gas-turbine engine in no-afterburning modes2 dwg

Gas turbine plant with an afterburner device for separating co2

Abstract: A gas turbine plant (1) comprising at least one gas turbine (10) with a compressor stage (11) and an expansion stage (12), an exhaust gas line (20) to guide the flow of exhaust gas exiting from the expansion stage (12) during operation of the gas turbine (10) and a CO2 separation device (30) connected by fluid technical means to the exhaust gas line (20) for the separation of at least a part of the CO2 contained in the guided exhaust gas, wherein an afterburner device (40) is connected into the exhaust line (20) between the gas turbine (10) and CO2 separation device (30) by fluid technical means, which device is designed to afterburn fuel, particularly fossil fuel, with the addition of exhaust gas, and wherein the afterburner device (40) is a gas turbine (40) which has an expansion stage (42) which permits only expansion to a pressure level which lies above the pressure level of the exhaust gas introduced into the compressor stage (41) of the gas turbine (40) concerned.

Rotating granular fuel burner with steps and air lifting fuel

Abstract: A combustion device for granular solid fuel, e.g. wood pellets, has a rotating cylindrical combustion chamber 102 with an outer wall 104, the inner surface of the combustion chamber having steps 138 and primary air apertures 140, the steps 138 and apertures 140 lifting fuel when the combustion chamber is rotated. Primary through apertures 140 can move parallel or circumferentially in the chamber. Secondary combustion air is provided through apertures (204, fig 2) in a centre plate 125, for combustion in an afterburner 114 with a smaller radius than the chamber 102. Rotation of the chamber can be pulsed, and adjusted to suit the fuel. A helical feeding device 116, blast apparatus 112, and electric lighting means 118 can be provided. The steps 138 can have a ceramic coating e.g. titanium nitride (TiN).

Steady State Thermal Analysis of an Afterburner Liner

Abstract: Military aero engines employ afterburner system for increasing the reheat thrust required during combat and takeoff. During reheat the gas temperature in the afterburner is of the order of 2100K.The afterburner liner has to be cooled with the available bypass air to maintain metal temperature within allowable limits. The liner has cooling rings at the rear to cool the liner with tangential film cooling.This paper discusses the methodology of afterburner liner metal temperature prediction and comparison with measured metal temperature during aero engine testing at reheat condition. All the modes of heat transfer are considered for thermal analysis, radiation due to higher level of gas temperature during reheat, conduction due to presence of low conductivity thermal barrier coating and convection due to higher gas velocities are considered.At different steady state reheat conditions metal temperature are predicted and compared with measured data during aero engine testing. The predicted skin temperatures and measured temperatures are in good agreement. Empirical correlations are used for estimating the heat loads coming on the liner and adiabatic film temperature near screech holes and cooling rings. Metal temperature and thermal loads coming onto the liner are predicted with 1D code. The estimated thermal loads are applied on 3D FE model to obtain nodal temperature distribution. The thermal Analysis is carried using ANSYS software in which thermal barrier coating is also modeled.The parameters like gas temperature, thermal barrier coating thickness, coating conductivity, and coolant mass flow distribution are considered for carrying out a sensitivity analysis of liner metal temperature.Copyright © 2013 by ASME

Design and development of major balance of plant components in solid oxide fuel cell system

Abstract: The balance of plant (BOP) of a Solid Oxide Fuel Cell (SOFC) system with a 2 kW stack and an electric efficiency of 40% is optimized using commercial GCTool software. The simulation results provide a detailed understanding of the optimal operating temperature, pressure and mass flow rate in all of the major BOP components, i.e., the gas distributor, the afterburner, the reformer and the heat exchanger. A series of experimental trials are performed to validate the simulation results. Overall, the results presented in this study not only indicate an appropriate set of operating conditions for the SOFC power system, but also suggest potential design improvements for several of the BOP components.

Jet turbine engine

Abstract: FIELD: engines and pumps.SUBSTANCE: jet turbine engine includes a housing, in which a turbine with an exhaust cone, a mixer, an afterburner duct and a nozzle are located in series. The exhaust cone includes cooled and non-cooled parts. The cooled part is provided with an annular channel formed with an inner shell and a perforated surface of the exhaust cone. The mixer forms together with the housing a cold air channel, and with the non-cooled part of the exhaust cone a hot air channel. The afterburner duct includes annular flame stabilisers. A small annular flame stabiliser is of a cooled type and installed so that it screens the non-cooled part of the exhaust cone on the nozzle side.EFFECT: invention allows reducing the infrared emission level of the jet turbine engine to rear semi-sphere of the aircraft and full pressure losses, and reducing overall dimensions and weight of the engine exhaust unit.5 cl, 3 dwg

Simulation of twin spool turbojet engine

Abstract: Simulation of performance of gas turbine engine fitted in fighter class aircraft is tried both for steady and transient operations. Performance of the engine is calculated for the inputs of air and fuel mass flow rates, altitude conditions and throttle positions for sea level and combinations of varying Mach number and altitude conditions. Methodology consists of usage of a series of equations for various modules of engine and all are sequentially evaluated to get the output namely thrust, specific fuel consumption and station values like pressures and temperatures. Dynamic simulation consists of varying the throttle position, when the engine is on ground and obtaining various performance parameters of the engine as outputs. In order to do this a series of equations are developed in order to arrive at the transfer function of the spools of the engine. For estimating the stability of the engine step inputs in the form of fuel flow rate for HP spool and exit area variation for LP spool are tried and found to be as per the expectations. 1. INTRODUCTION Gas turbine engines have come into existence replacing piston engines because piston engines could not be operated at higher altitudes and higher forward Mach numbers. Despite the fact that gas turbine engines are seriously inefficient compared to piston engines, this aspect is overlooked and many modifications were carried out on turbojets to improve the efficiency aspects. As, thrust to weight ratio improves, the generation of engines is also parallely raised. This paved the way for twin spool turbojet engine with afterburner. Engines for application in fighter class aircrafts are always in the developmental phase to meet the avaracious demands of the pilot. While doing so, simulation of the engine performance occupies a large chunk of developmental exercise. Since fighter class engines demand for transient dynamics due to deep throttle requirements, it is necessary to carry out the simulation for transient dynamics also. In order to carry out the static simulation a series of equations, which are available in open literature is used barring aside afterburner performance. The static simulation that is carried out on the engine is covered by the boundary conditions like ISA sea level static, various RPM conditions of the engine, variations in altitude and Mach number combinations. The platform that is used for calculation is Matlab and the results that are obtained are compared with the information available in gas turbine literature. As said earlier, since dynamic simulation is also to be carried out, a set of equations are derived and used for this purpose. Both afterburner operations and altitude and Mach number combination are safely neglected because the error accumulation in the dynamic simulation is observed to be prohibitive. Transient dynamics since requires transfer functions for identifying the response, an attempt is made by using step input for both LP and HP RPMs. The step input is fuel flow rate for HP spool while variation in exhaust area for the LP spool (8). Crossover frequencies are also required in order to identify the stability of the engine either during acceleration or deceleration, for locating the margin availability.

Fuel cell system e.g. mobile fuel cell system, for use in e.g. motor vehicle, has gas conveyer unit supplying oxygen-containing gaseous mixture to afterburner and fuel cell and standing in direct gas-conductive contact with afterburner

Abstract: The system has a fuel cell (2) i.e. solid oxide fuel cell (SOFC), comprising an anode (3) and a cathode (4) between which an electrolyte (5) is arranged. An afterburner (25) is provided for burning fuel, which is contained in anode exhaust gas, using an oxygen-containing gaseous mixture i.e. air. An anode exhaust system (26) conducts the anode exhaust gas to the afterburner, and a gas conveyer unit (17) e.g. blower, supplies the oxygen-containing gaseous mixture to the afterburner and optionally the fuel cell and stands in direct gas-conductive contact with the afterburner. An independent claim is also included for a method for operating a fuel cell system.

Fuel cell system with afterburner for anode waste gas

Abstract: A fuel cell system comprises a fuel cell 2 having an anode 3 and a cathode 4, wherein an electrolyte 5 is arranged between the anode 3 and the cathode 4, and an afterburner 25 for burning fuels contained in anode waste gas using an oxygen-containing gas mixture. The fuel cell system also includes an anode gas channel 26 for carrying anode waste gas to the afterburner 25, and at least one gas conveying unit 17, (28, figure 3) which feeds an oxygen-containing gas mixture, in particular air, to the afterburner 25, and optionally to the fuel cell 2, and is in direct gas-carrying contact 35 with the afterburner 25. Such a fuel cell system allows the working point of the afterburner to be adjusted independently of the cathode gas requirement of the fuel cell. A method of operating a fuel cell system is also disclosed.

A Numerical Study on Effects of Displacement of a Variable Area Nozzle on Flow and Thrust in a Jet Engine

Abstract: Variable area nozzle, where both throat and exit area vary, is required for optimal expansion and optimal nozzle shape upon operation of after-burner. Steady-state and transient analyses are carried out for each condition with and without afterburner operation and as a function of the location of the nozzle flap. Effects of that nozzle displacement on flow and thrust characteristics are analyzed from numerical results. With variable area nozzle adopted, the combustion field is variable in time, leading to periodically variable thrust. For off-design conditions, flow separation shows up due to over expansion at the flap tips and shock wave does in the nozzle due to under expansion. The undesirable phenomena can be solved by control of variable area nozzle.

Propulsion Models Implemented in Finite Element CFD

Abstract: Fast and efficient models are presented for three means of propulsion: Rocket, turbojet, and ram/scramjet systems. These propulsion models are desirable for evaluating vehicle performance, stability, and control, especially dur ing the early design stages. Rocket and turbojet boundary conditions along with quasi-combustion terms are investigated using an inviscid finite element approach. Results are pres ented for rocket nozzle, three turbojet inlets, coupled turbojet, simple afterburner, and g eneric hypersonic vehicle with a scramjet. These results show the application of the rocket, t urbojet, and quasi-combustion terms. The simplicity of the models lend themselves to concept ual and early trade studies. Such applications are discussed briefly as they apply as an early design tool.

Fuel cell system for motor car, has bypass guide connected with cathode exhaust gas guide, through which cathode exhaust gas is conductible, and conveyor device is provided downstream of fuel cell for conveying cathode exhaust gas

Abstract: The system (1) has an afterburner (25) which is provided for burning fuel contained in anode exhaust gas using cathode exhaust gas as oxidants. An anode exhaust gas guide (26) is provided for guiding the anode exhaust gas to afterburner, and cathode exhaust gas guide (23) is provided for guiding cathode gas to afterburner. A bypass guide (27) is connected with cathode exhaust gas guide downstream of fuel cell (2) and upstream of afterburner, through which cathode exhaust gas is conductible. A conveyor device (28) is provided downstream of fuel cell for conveying cathode exhaust gas. An independent claim is included for method for operating fuel cell system.

Utilisation method of residual components of rocket fuel in spent stages of carrier rockets

Abstract: FIELD: engines and pumpsSUBSTANCE: utilisation method of residual liquid components of rocket fuel in spent stages of carrier rockets by their simultaneous supply to an afterburner at the step of lowering the stage with afterburning of residual fuel, after oxidiser ends, with air oxygen and by ejection of combustion products through side surface of the stage housing, according to the invention, ejection of combustion products is performed through convergent-divergent nozzles located in structural planes of the stage in the directions perpendicular to its longitudinal axis, and reaction forces occurring from convergent-divergent nozzles are used for changing angles of attack and wobbling of the stage housing and control of its lowering due to its aerodynamic quality of the stage housingEFFECT: invention provides utilisation of rocket fuel components by using their energy for control of the stage lowering2 dwg