Showing papers on "Combustion chamber published in 1969"

Research on combustion instability in liquid propellant rockets

Abstract: Nonlinear combustion instabilities in liquid propellant rockets, considering various combustion models and experimental techniques

Vortex mixing for supersonic combustion

Abstract: Fundamental concepts of mixing and combustion theory are examined in order to define an optimum system. By analogy with the familiar parameter “combustion intensity,” a “mixing intensity” is defined as the total mass of fuel and air mixed per unit time, volume, and density. Mixing intensity is related to the mean kinetic energy of turbulence generated in a mixing system. The efficiency of generation of turbulence is defined by the ratio of the kinetic energy of turbulenee generated to the kinetic energy loss from the mean flow. Bluff body mixers are compared with turbulent jet mixers, and turbulent intensities are determined as a function of fuel/air ratio and jet-to-stream-velocity ratio. For turbulence generated by a baffle system or by jets, the mixing intensity is shown to be proportitional to the rms velocity and the square root of the number of sources. Substantial increases in mixing rates can be obtained by applying a swirling motion to the central fuel jet. Radial and axial pressure gradients are set up which, in the case of strong swirl, result in reverse flow along the axis. Empirical equations for velocity decay, mass entrainment, and angle of jet expansion are given as functions of the degree of swirl. However, further work is needed on the behavior of these parameters in an external flow field. For the combustion chamber of a scramjet, the degree of swirl imparted to the fuel can be altered readily by varying the proportion of fuel introduced axially and tangentially into the swirl chamber. The rate of mixing, angle of jet expansion, and the size of the internal reverse-flow zone can all be controlled by variation of the swirl number. With the ability to initiate and remove the reverse-flow region, an optimized combustion system could operate efficiently from low-subsonic to high-supersonic Mach numbers.

Comparison of combustion instabilities found in various types of combustion chambers

Abstract: Combustion instabilities in motors and furnaces are contrasted under three separate headings: System instabilities involve, in an essential way, interactions between processes occuring in the combustion chamber and processes occurring in at least one other component of the system. Combustion chamber instabilities consist of unsteady phenomena that are localized within the combustion chamber. Intrinsic instabilities are inherent in the reactants themselves and would be observed if combustion were to occur in the absence of any external influences. The second category, combustion-chamber instabilities, is again divided into three subcategories: Acoustic instabilities involve, in an essential way, the propagation of acoustic waves in the combustion chamber. Shock instabilities are characterized by the presence of steep-fronted finite-amplitude shock or detonation waves in the combustion chamber. Fluid-dynamic instabilities are associated with the establisment of special kinds of flow patterns, such as vortices, in the chamber. Experimental and theoretical aspects of each of these types of instabilities are discussed separately. Characteristics of each type of instability are described, and the kinds of chambers in which each has been observed are indicated. Attention is given to surface-burning systems, such as solid-propellant rockets and hybrid rocket motors, and also to volume-burning systems, such as liquid-propellant rocket engines and both gas-fired and oil-fired industrial burners. Experimental procedures for investigating each type of instability are reviewed along with selected experimental results, and the character and state of development of the theoretical methods that have been used for describing each type of instability are compared. The review is an attempt to bring about discussions between specialists who study combustion instabilities in different systems. Classification according to the type of instability rather than the type of system can emphasize the similarities and differences that exist among combustion instability problems encountered in different systems. One resulting observation is that examples of system instabilities and of acoustic instabilities are common to all kinds of combustion systems.

Combustion chamber construction

Abstract: THE SUBJECT COMBUSTION CHAMBER CONSTRUCTION HAS AN OUTER CASING ENCLOSING AN INNER CASING TO FORM AN ANNULAR COMBUSTION CHAMBER WITH AN AIR PLENUM OR PASSAGE EXTENDING ALONG THE OUTSIDE, RADIALLY INWARDLY AT THE REAR END, AND AXIALLY ALONG THE INNER SIDE AIR FROM THE ENGINE COMPRESSOR ENTERS THE PLENUM AT THE FRONT END AND FLOWS THROUGH THE PASSAGE INTO THE COMBUSTION CHAMBER THROUGH OPENINGS IN THE INNER WALL THEREOF, SUCH AIR THEN FLOWING TANGENTIALLY FROM THE INNER WALL TO FORM A TOROIDAL FLOW PATTERN FUEL IS SPRAYED INTO COMBUSTION CHAMBER FROM TANGENTIALLY DIRECTED JETS TO MIX AND BURN WITH THE AIR IN THE TOROIDAL FLOW PATTERN, THE RESULTING GASES DISCHARGING FROM THE CHAMBER VIA AN ANNULAR RADIALLY INWARDLY DIRECTED NOZZLE ADJACENT THE AIR-INLET THE TOROIDAL FLOW PATTERN IN THE ANNULAR CHAMBER PROLONGS THE AIR/FUEL MIXING AND BURNING PERIOD

Rocket burner with flame pattern control

Abstract: An oxygen-fuel burner of the rocker burner type comprising a cylindrical combustion chamber having an open discharge end and a burner plate with separate oxygen and fuel ports constituting the opposite end of the chamber; the projected longitudinal axes of the oxygen ports extending in converging directions towards the longitudinal axis of the chamber but in off-set, non-intersecting relation thereto, so that points on the respective axis that most closely approach the chamber axis define a transversely positioned plane between the burner plate and the chamber exhaust; the projected longitudinal axes of the fuel ports being substantially parallel to the chamber axis for mixing of oxygen and fuel at and beyond the plane of closest approach, and means for adjusting the longitudinal position of the burner plate on the chamber axis and thereby locating the plane of closet approach in relation to the chamber exhaust for determining the pattern of the burner discharge flames

Piston and piston rings unit for an internal combustion engine

Abstract: The invention has for its object, on the one hand, all means which protect, at least partially, the internal face of the sealing joint placed between the lower adjacent faces of the piston ring and the piston groove against the high pressures which, from the combustion chamber pass between the compression ring and the bottom of the groove and also into the crosssectional or gap clearance of the compression ring, and on the other hand, a support for the said joint, which is not affected by creep, is annular, is not fixed to the piston and is arranged at least under the periphery of the said joint between the joint and the lower face of the piston groove, the peripheral face of the support being maintained at least in proximity to the cylinder and having as a maximum narrow discontinuities, the maximum clearances of this proximity and of these discontinuities being smaller than those in which the material of the said joint begins to creep, under the conditions of temperatures and pressures reached during the operation of the engine.

Rocket engine combustion chamber

Abstract: A combustion chamber construction includes an inner wall having longitudinally extending cooling channels defined in the exterior thereof and an outer wall which is bonded to the inner wall by galvanizing, both the inner wall and outer wall are made of a single piece of an oxygen-free copper or equivalent material such as silver or molybdenum. The cooling channels are advantageously cut in accordance with the method of the invention such as by machining and formed with the least wall thickness in the area of the thrust nozzle and with the relatively greatest wall thickness in the area of the discharge of the combustion chamber. The internal wall thickness at the head portion of the combustion chamber is made of medium thickness. After the cooling channels are machined into the inner wall portion the channels are filled with a filler material which is electrically conductive and has a low melting point so that upon galvanizing it is melted out. In some instances an intermediate layer is positioned between an outer relatively thick and strong layer and inner wall of the combustion chamber, and the three layers are galvanized together.

Nonlinear Longitudinal Instability in Rocket Motors with Concentrated Combustion

Abstract: Liquid rocket motors with concentrated combustion, studying nonlinear longitudinal instabilities with shock waves in combustion chambers

Ignition system for rotary piston internal combustion engine

Abstract: A rotary piston internal combustion engine carries two ignition plugs exposed to a common combustion chamber with ignition circuits corresponding to the two plugs including means for delaying the ignition time for either of the two ignition plugs in response to low speed and/or low load operation of the engine.

Composite gas turbine ramjet engine

Abstract: A composite turbo-ramjet engine comprising, in combination: TWO DUCTS SEPARATED FROM EACH OTHER BY A WALL AND SUPPLIED WITH AIR FROM A COMMON AIR INTAKE, THE TWO DUCTS CONSISTING OF, FIRSTLY, A PRIMARY TURBOJET DUCT WHICH INCLUDES A GAS-TURBINETYPE GAS GENERATOR OF HOT GASES, A PRIMARY DISCHARGE NOZZLE, AND ADJUSTABLE MEANS TO BLOCK OFF THE SAID PRIMARY NOZZLE, AND SECONDLY, A SECONDARY DUCT INCLUDING A SECONDARY COMBUSTION CHAMBER FOR RAMJET OPERATION AND A SECONDARY DISCHARGE NOZZLE; AT LEAST ONE CONNECTING PASSAGE BETWEEN THE PRIMARY AND THE SECONDARY DUCTS, THE SAID PASSAGE BEING FORMED THROUGH THE SAID WALL DOWNSTREAM OF THE TURBINE AND UPSTREAM OF THE SECONDARY COMBUSTION CHAMBER; ADJUSTABLE OBTURATION MEANS FOR THE SAID CONNECTING PASSAGE, AND ADJUSTABLE OBTURATION MEANS FOR THE SECONDARY RAMJET DUCT UPSTREAM OF THE SECONDARY COMBUSTION CHAMBER. The present invention relates to a composite propulsion device for airborne machines, and more especially for aircraft, that is capable of supplying with good efficiency a high thrust over the full range of speeds included between the takeoff of the aircraft and a cruising speed well into supersonic figures. Composite propulsion devices are already known, constituted by the association of a gas turbine engine with a ramjet engine. These devices put to good use the differing and to some extent complementary qualities of these two types of engine, of which it may be said, in summary form, that the first-named is suitable for propulsion at relatively low, subsonic or slightly supersonic speeds, whereas the second-named makes it possible to achieve, with good efficiency and a low fuel consumption, speeds which extend far into the supersonic range. The invention has for its object an improved composite gas turbine and ramjet engine, conveniently termed a turbo-ramjet engine, which makes it possible to obtain excellent performance figures over a wider range of flight parameters than in the case of already known combinations. As will be seen hereinafter, the said composite engine brings about the synthesis of the qualities of reheat-equipped gas turbine jet engines of the single-flow or of the dual-flow type (bypass engines) at subsonic or slightly supersonic speeds, and of the qualities of ramjets with high supersonic speeds, and in other respects it possesses a series of advantages, more especially with reference to the possibility, after the extinction of the gas turbine combustion chamber when high speeds have been attained, of allowing the moving parts of the said gas turbine engine to windmill under the effect of the relative wind velocity. The composite engine according to the invention includes in combination: TWO DUCTS SEPARATED FROM EACH OTHER BY A WALL AND SUPPLIED WITH AIR FROM A COMMON AIR INTAKE, THE TWO DUCTS CONSISTING OF, FIRSTLY, A PRIMARY DUCT WHICH INCLUDES A GAS-TURBINE TYPE GAS GENERATOR, A PRIMARY DISCHARGE NOZZLE, AND ADJUSTABLE MEANS TO BLOCK OFF THE SAID PRIMARY NOZZLE, AND SECONDLY, A SECONDARY DUCT COMPRISING A SECONDARY COMBUSTION CHAMBER AND A SECONDARY DISCHARGE NOZZLE; AT LEAST ONE CONNECTING PASSAGE BETWEEN THE PRIMARY AND THE SECONDARY DUCTS, THE SAID PASSAGE BEING FORMED THROUGH THE SAID WALL DOWNSTREAM OF THE TURBINE AND UPSTREAM OF THE SECONDARY COMBUSTION CHAMBER; ADJUSTABLE OBTURATION MEANS FOR THE SAID CONNECTING PASSAGE, AND OBTURATION MEANS FOR THE SECONDARY DUCT UPSTREAM OF THE SECONDARY COMBUSTION CHAMBER. The two ducts are preferably coaxial, the secondary duct surrounding the primary duct. According to one arrangement of the invention, the means for adjustable obturation of the connecting passage between the two ducts comprises at least one flap hinged on the separating wall between the two ducts, the said flat in one of its extreme positions blocking off the said connecting passage. As regards the obturation of the secondary duct, the design may be such that the said duct comprises an upstream and a downstream section separated from each other by at least one obturatable aperture arranged in the vicinity of the connecting passage between the two ducts, the said aperture being able to be blocked off by the selfsame flap in the other of the extreme positions of the said flap. Other means, such as a sliding valve body, may, however, be provided to block off the secondary duct. According to one arrangement of the invention and applicable to the case in which the gas generator forms part of a gas-turbine engine with a dual flow, that is, an inner hot flow and an outer cold flow, the inner flow passes along the primary duct, whereas the outer flow is injected into the secondary duct, downstream of the obturating means of the said duct, via supplementary connecting passages between the two ducts, these passages being, if so required, of adjustable cross section. The arrangement set out above renders it possible, as will be seen hereinafter, to make the composite power plant operate according to several possible configurations, with the option of continuous transition from one to another: a pure turbojet, a turbo-ramjet with a partially mixed flow, a turbo-ramjet with separate flows, and a pure ramjet, each of the said configurations possessing its own advantages, depending on flight conditions. The following description with reference to the accompanying drawing and given by way of nonlimitative example will bring out how the invention may be put into effect.

Internal combustion engine including maximum firing pressure-limiting means

Abstract: A variable compression ratio internal combustion engine wherein the engine crankshaft together with pistons and connecting rods are moved in such a way to increase or decrease the compression ratio (decrease or increase combustion chamber volume) during operation to thereby limit maximum firing pressure to a predetermined valve for all engine loadings. The crankshaft is suspended in bearings in support members extending transverse to the crankshaft. The support members are pivotally supported at one end in the engine frame and individually suspended at the opposite end from an auxiliary crankshaft or the like.

Development of an Advanced Annular Combustor

Abstract: The objective of the effort described in this report was to determine the structural durability of a full-scale advanced annular turbojet combustor using ASTM A-1 type fuel and operating at conditions typical of advanced supersonic aircraft. A full-scale annular combustor of the ram-induction type was fabricated and subjected to a 325-hour cyclic endurance test at conditions representative of operation in a Mach 3.0 aircraft. The combustor exhibited extensive cracking and scoop burning at the end of the test program. But these defects had no appreciable effect on combustor performance, as performance remained at a high level throughout the endurance program. Most performance goals were achieved with pressure loss values near 6% and 8%, and temperature rise variation ratio (deltaTVR) values near 1.25 and l.22 at takeoff and cruise conditions, respectively. Combustion efficiencies approached l004 and the exit radial temperature profiles were approximately as desired.

Reciprocating internal combustion engine with constant pressure combustion

Abstract: A reciprocating internal combustion engine adapted to provide substantially constant pressure combustion in a combustion chamber separated from the compression and expansion cylinders of the engine is provided.

Combustion chambers for gas turbine engines

Abstract: A combustion chamber for a gas turbine engine has a primary air inlet and a number of additional air nozzles at intervals along the chamber. Each of the air inlets has one or more holes in the inlet wall, so that air may be injected into the inlets or extracted therefrom. Air injection effectively reduces the inlet areas and extraction increases the inlet areas. The airflow through each inlet, or group of inlets, may thus be controlled, the arrangement being such that, irrespective of the proportion of the total airflow which enters each inlet, the resistance to airflow through the combustion chamber does not vary. The holes in the inlet walls may be tangential to create a vortex within the inlet.

Dual mode supersonic combustion ramjet engine

Abstract: A supersonic ramjet engine having a fixed-geometry combustion chamber. The ramjet engine is operated in the subsonic mode by injecting fuel in fuel injectors located in a uniform cross-section portion of the combustion chamber and in the supersonic mode by injecting fuel in fuel injectors located upstream of the subsonic injectors in an adjacent expanding portion of the combustion chamber.

Nonlinear Periodic Oscillations in Rocket Motors with Distributed Combustion

Abstract: Transverse and longitudinal combustion chamber oscillations in rocket motors with distributed mass and energy sources

An investigation of the swirling flow in a spinning end-burning rocket

Abstract: The purpose of this study was to determine the swirling flowfield which develops at various spin rates in the chamber of a spinning end-burning rocket. Cold flow experiments employing a porous plate to represent the propellant surface were conducted. The experiments were designed to simulate dynamically the conditions in a typical cylindrical rocket combustion chamber. Pressure measurements at the plate and in the flow were correlated with smoke traces to arrive at a physical description of the flowfield in which solid-body rotation gives way to a high-speed vortex in the chamber at sufficiently high motor spin rates. The experimental correlation of the onset of this flow transition with a critical value of the Rosby number is in agreement with a theoretical analysis, and it is concluded that the vortex development is a characteristic of the inviscid rotating flow. The experiments also indicate that significant interaction between the vortex and the propellant surface is little affected by chamber length or nozzle entrance shape.

Burners having a pulsating mode of operation

Abstract: 1,254,453. Pulsating burner processes. BRITISH PETROLEUM CO. Ltd. 22 Aug., 1969 [16 Sept., 1968], No. 43938/68. Heading C2C. [Also in Divisions C1 and F1] A burner having a pulsating mode of operation comprises a combustion chamber with grossly rough walls and a gaseous oxygen/fuel inlet system which has a low resistance to gaseous flow and which is arranged to mix the fuel and oxygen at one end of the combustion chamber so that in operation a series of explosion waves is produced by repeated ignition of an explosive mixture formed in the combustion chamber. In Fig. 2 the wall of the combustion chamber 10 has a spiral metal rod 27 secured thereto. At the inlet end of the chamber is a mixing chamber 24 to which fuel is supplied through duct 21 and air through annular duct 22. A sparking plug 26 projects into the chamber. In Fig. 3 air is supplied through manifold 33 and ducts 29 and fuel is supplied through manifold 28 and ducts 30, an igniter device 26 projecting through the end wall of the chamber. In Fig. 6 an antechamber 24 is provided, an arm 40 having a vent 41 communicating with the chamber. A glow filament 42 is disposed adjacent the vent and serves to ignite the charges. Alternatively a pilot flame could be provided adjacent the vent 41. The burner may be used to obtain formaldehyde by combustion of methane, to telomerize ethylene with carbon tetrachloride and to form 1-naphthol from naphthalene.

Rotary piston machine

Abstract: This invention is an internal combustion engine having arcuate pistons inside a rotating annular cylinder block which is divided into two combustion chambers by a pair of diametrically opposed cylinder heads. Each combustion chamber has an air intake port and an exhaust outlet for burned gases. The arcuate pistons form within the cylinder heads two combustion chambers and two compressors or pumps. The relative rotation of the cylinder and the arcuate pistons is controlled by mechanical scissor-action-type eccentric cranks connected to an eccentrically rotatable ring.

Method and apparatus for reducing pollutants in the exhaust gas of an internal combustion engine

Abstract: A method and apparatus for reducing pollutants in the exhaust gas of an internal combustion engine of the type wherein air-fuel mixture is supplied to the engine combustion chamber through an air suction pipe having a carburettor with a main nozzle and a throttle valve, and comprising means for feeding oxygen to the air-fuel mixture in said carburettor when the throttle is between its closed and one-fourth open positions and the engine is operating at light load, means for stopping said oxygen feed and feeding inactive gas to the air-fuel mixture in the carburettor when the throttle valve is approximately between one-fourth and three-fourths open and the engine is operating at medium load condition, and means for stopping said inactive gas feeding and again feeding oxygen to the air-fuel mixture in the carburettor when the throttle valve is between approximately three-fourths and fully open, and the engine is operating at full load condition.

Pollution control system for internal combustion engines

Abstract: A pollution control system for internal combustion engines including a crankcase unit for scavenging the blowby gases from the oil pan and an exhaust unit for scavenging the exhaust gases in the engine and exhaust manifold.

Rotary piston injection-type internal combustion engine

Abstract: A rotary piston injection-type internal combustion engine, especially of trochoidal type of construction in which a throttle valve mounted on a throttle valve shaft is arranged in the inlet channel to control the amount of combustion air flowing through the inlet channel; the throttle valve and/or the throttle valve shaft are provided with one or several bores for the flow of combustion air through the same when the throttle valve is in the closed or partially open position.

Method for in-situ utilization of fuels by combustion

Abstract: Subsurface combustible material, such as coal, can be burned in situ by providing a number of passages which extend downwardly from the surface and which extend to, and generally horizontally through, the combustible material, by forming a combustion chamber within said combustible material adjacent to and in communication with the ends of said passages, and by using one of said passages to supply air to said combustion chamber while using another of said passages to remove products of combustion from said combustion chamber. The wall of said combustion chamber, which is adjacent the inner ends of said passages, will progressively burn away, and thus cause said combustion chamber to ''''move'''' toward said passages; and the overburden adjacent said wall will provide a self-supporting overhang which will enable air to continue to enter said combustion chamber from said one passage and will enable products of combustion to move to and outwardly through said other of said passages - even though the overburden adjacent the opposite wall of said combustion chamber slumps down into said combustion chamber adjacent to said opposite wall, reach the area where that combustible material is being burned.

Improvements in liquid fuel vaporizing combustion systems

Abstract: A T-shaped liquid fuel prevaporizing device for a combustion chamber, said device comprising a tubular shank projecting into the combustion chamber and two arms projecting from the downstream section of said shank to open into the combustion chamber, wherein overheating of the device is minimized by compartmenting the downstream section of the tubular shank by longitudinal partitions to form separate but intercommunicating passages which are selectively fed with combustion air alone and with fuel/air mixture.

Oscillating internal combustion engine

Abstract: The invention hereinafter disclosed is a power unit which includes an internal or external combustion engine having an annular cylinder, arcuate pistons which oscillate in that cylinder and improved mechanism for translating the power developed in the engine to a shaft to be driven. The engine may be supplied with fuel either by injection or carburetion or by steam or fluid or gas under pressure and the arrangement is such that vibrations due to explosions of the fuel mixture or of power strokes are minimized and the output shaft is powerfully driven at uniform speed.

Penetration and Vaporization of Diesel Fuel Sprays

Abstract: In the diesel engine, fuel is injected into the hot, compressed air in the combustion chamber. Thus the process of diesel combustion is essentially inhomogeneous, and the mixing of the fuel and air in the combustion chamber dominates the whole combustion process. Since fuel–air mixing is so important the distribution of the injected fuel has a major effect on combustion performance. This is particularly true of direct-injection diesel engines which have relatively low rates of air movement.In all diesel engines, fuel is injected into the combustion chamber at high pressure through small nozzles. The high-velocity liquid jet atomizes, after emerging from the nozzle, into a spray of liquid droplets. The penetration, distribution, and vaporization of the sprays, together with the air movement, govern the mixing of fuel and air.The penetration of fuel sprays is dealt with in Part 1 of the paper; Part 2 describes a study of the vaporization of fuel sprays.