GB1605196A

GB1605196A - Ramjet engine

Info

Publication number: GB1605196A
Application number: GB345278A
Authority: GB
Original assignee: Messerschmitt Bolkow Blohm AG
Current assignee: Airbus Defence and Space GmbH
Priority date: 1977-02-08
Filing date: 1978-01-27
Publication date: 1983-04-13
Also published as: FR2519377A1; DE2705078A1; FR2519377B1; DE2705078C2

Description

(54) A Ramjet Engine (71) We, MESSERSCHMITT-BOLKOW- BLOHM Gesellschaft mit beschränkter Haftung, of 8000 Mtinchen, German Federal Republic, a Company organised and existing under the laws of the German Federal Republic, do hereby declare the invention, for which we pray that a Patent may be granted to us, and the method by which it is to be performed, to be particularly described in and by the following statement: This invention relates to a ramjet engine primarily one for operation with fuels containing boron or a similar high-energy particulate material having high melting and high vaporisation points the engine having inlet diffusers distributed around its periphery and serving to supply ram air to a combustion chamber.

In known ramjet engines for use as cruise propulsion units for aircraft a fully integrated construction can be adopted but this still leaves room for improvement as regards the degree of burn out and stability range especially when liquid fuels are replaced by high-energy particulate fuels, such as suspensions of boron, aluminium, magnesium or the like in liquid hydrocarbons, or propulsion charges with mixtures of this kind. This drawback is due to the fact that with the propulsion unit concerned the connection of all the inlet diffusers to the combustion chamber is provided in the front part thereof.

When the total ram air is supplied to the combustion chamber in this way the production of a zone of sufficiently high temperature and a sufficiently long dwell period in the interior of the chamber for the main combustion process, combined with the moderate temperature required at the combustion chamber outlet proves difficult. It has been found that the use of inserts generally obstructs the combustion chamber and causes problems when the starting propulsion unit is accommodated inside the combustion chamber.

An object of this invention is to provide a ramjet of simple and reliable design with a more satisfactory degree of burnout and better stability range than that of known propulsion unit constructions of the same type, without the need to build bulky inserts into the combustion chamber.

According to this invention there is provided a ramjet engine for operation with a fuel formed from a combustible propulsive charge containing boron or similar high-energy particulate material with high melting and high vaporisation points, the engine having air inlet diffusers distributed around the periphery of the engine body and serving to feed ram air to a combustion chamber of the engine also fed with the fuel formed from combustion of the charge, the total ram air flow from the inlet diffusers being subdivided into combustion air which is fed to a forward combustion chamber zone forming a flame chamber and into mixing air fed to a rearward combustion chamber zone serving as a mixing chamber.

The ramjet thus has a front combustion chamber zone in which the overall cross section of the chamber is presented not to the total quantity of ram air but only to the quantity of combustion air, amounting to a small fraction of the total. The velocity retardation and thus the dwell period or (staying time) of the combustion air in the front combustion chamber zone is increased accordingly. The reaction between the combustion air and the fuel takes place at a very high temperature, which offers the advantage of a high degree of bumout and an ample stability range. This effect is particularly evident with fuels of the kind containing boron or similar high energy particulate material with a high melting point and vaporisation temperature.If the temperature were too low in the front zone of the combustion chamber a liquid layer of oxide would adhere to the surface of the particles. For the combustion of the particle material on the inside of the layer of oxide, it would first have to be reached by the oxygen required and in the space and flow conditions prevailing this would be difficult.

The high temperature, which is advantageous for the aforementioned reasons, can be achieved in the front zone of the combustion chamber and an . economical final temperature nevertheless obtained, without recourse to bulky combustion chamber inserts and this is due to the way in which the flow of mixing air is guided in accordance with the invention. An impor- tant feature of this system is the fact that the mixing air is conveyed to the location in the combustion chamber which is best for combustion. The ramjet according to the invention, makes it possible to dispense with combustion chamber inserts for the diversion of the mixing from the total air and this is also found to be an advantage when the combustion chamber is required to accommodate a rocket propulsion unit for starting purposes.

The inlet diffusers may each be followed by a channel for mixing air, situated in the outer casing and leading to the mixing chamber, laterial overflow apertures being provided for the supply of combustion air to the flame chamber. If separation of the mechanics of flow between the combustion air flow and the mixing air flow is required, this can be achieved by the simple device of providing sonic nozzles in the mixing air channels. This leads to additional pressure losses which can be avoided if the sonic nozzles in the mixing air channels are dispensed with and if the supersonic zones of the individual inlet diffusers are used for the flow separation process.This may be achieved by having a separating plate in each inlet diffuser, extending through the subsonic zone into the supersonic zone in order to subdivide it into a combustion air sector merely communicating with the overflow apertures to the flame chamber and into a mixing air sector merely communicating with the channels to the mixing chamber. If the latter separating method is adopted a simple slide valve can also be used as a means of varying the proportion of combustion air and the proportion of mixing air as required.

In the absence of separating plates, however, the respective proportions represented by the combustion air and the mixing air can be varied by very simple means and all that is required is to provided a throttle device in all or some of the mixing air channels.

The separating measures described above are unnecessary if, as a preferred feature of the invention, separate inlet diffusers for combustion air and mixing air are provided on the periphery of the propulsion unit. In this case it is advisable in the case of aircraft with a constant difference between the diffusers as regards the approach flow velocity for those receiving the less satisfactory flow to be associated with the flame chamber, the others being associated with the mixing chamber.

Several embodiments in accordance with the invention are described with reference to the accompanying drawings, wherein: - Figure 1 shows a front view of a long-range projectile intended for use against marine targets, and equipped with a ramjet engine designed for supersonic operation; Figure 2 shows a longitudinal section through the projectile of Figure 1, Figure 3 shows a front view of a hemispherical supersonic inlet diffuser differing; from that shown in Figures 1 and 2, Figure 4 shows a further projectile with a propulsion engine differing from that shown in Figure 2, and again shown in longitudinal section, and Figure 5 shows a front view of the projectile shown in Figure 4.

The projectile 1 shown in Figures 1 and 2 has a front part 2 containing the payload and systems 3 and a chamber 6 in which a gas generator 7 is located. The function of the generator is to supply combustion gas for the operation of the rear ramjet engine 8. The combustion gas reacts spontaneously with ram air, making ignition means unnecessary. In the present case the combustion gas is the reaction product of a burning propulsive charge which has a very low oxygen content and a solid particle content of about 75% by weight, for example, in which boron, aluminium, ammoniumperchlorate and binding agent account for about 40%, 10%, 25% and 25% by weight respectively which reacts with the ram air.

The combustion chamber 9 and the thrust nozzle 10 of the rear ramjet 8 are designed for long combustion times. This construction provides, where the thermal loadings make it appear advisable, for the use of refractory material 11 for heat protection purposes. As regards the combustion chamber 9 it should be noted that this, as may be seen from Figure 2, functions in the front zone 9a as a flame chamber for the stoichiometric or almost stoichiometric combustion of the combustion gas flowing through a nozzle 12 in the front end wall of the chamber using a comparatively small quantity of the total ram air from suspersonic inlet diffusers 13 which are distributed around the periphery of the projectile (Figure 1) and may, for example, be four in number.In the rear zone 9b the combustion chamber serves as a mixing chamber for the remaining quantitiy of ram air and the hot combustion gas from the flame chamber 9a to the thrust nozzle 10. For the purpose of the aforementioned subdivision of the total ram air into combustion air and mixed air each supersonic inlet diffuser 13 is followed by a channel 16 for mixing air this channel leading through the outer casing 15 to the mixing chamber 9b.

Apertures 14 are provided for the supply of combustion air to the flame chamber 9a. The mixing air channels 16, of which four are provided end with apertures 17 in the combustion chamber casing at the transition point between the flame chamber 9a and the mixing chamber 9b. In the region of the apertures 17 the channels become oblique with respect to the longitudinal axis of the propulsion unit and this, in conjunction with the relatively large channel cross sections, is found to be of advantage for the mixing as it results in jets of mixed air which are able to flow through as far as the middle of the chamber.

If importance is attached to a separation between the flow of combustion air and the flow of mixed air, from the point of view of the mechanics of flow, this can be achieved by the provision of a venturi nozzle 18 in each of the mixing air channels 16. It is possible to control the development of heat in the flame chamber 9a when required if at least some of the mixing air channels 16 are provided with a throttle flap 19 or similar device to enable the quantity of mixing air to be varied.

Referring to Figure 3 this shows a supersonic inlet diffuser 20 which is semi-circular in plan with the outer casing, central core and boundary layer scoop marked 21,22 and 23 respectively.

It differs from the diffuser design shown in Figures 1 and 2 inasmuch as the provision of a venturi nozzle in the mixed air channels is dispensed with. Instead the interior of the diffuser is subdivided into a combustion air sector 24 communicating with the flow apertures 14 in Figure 2 for the flame chamber 9a and into a mixing air sector 25 communicating with the mixing air channels 16 in Figure 2 for the mixing chamber 9b. A separating plate 26 extends through the subsonic zone into the supersonic zone. This results in the separation of the flow of combustion air from the flow of mixing air and this in contrast to the method of separation mentioned earlier using a venturi nozzle, eliminates further pressure losses.It also makes it possible by using a simple slide valve to vary the proportions of the combustion and the mixing air and thus the development of heat in the flame chamber as required. This may be necessary as a result of a tendency for compressive surges to occur in the event of excessive angles of incidence for the projectile and unsatisfactory approach air flow to the diffuser.

The propulsion unit construction shown in Figures 4 and 5, as in Figures 1 and 2, is provided with four supersonic inlet diffusers 27 and 28, which are distributed around the periphery of the projectile 1 in the manner shown in Figure 1. The difference in the present case is that one pair of diametrically opposite dif- fusers 27 are connected through the apertures 29 to the flame chamber 9a only, while the other pair of diametrically opposite diffusers 28 are connected through channels 30 to the mixing chamber 9b only.

The aforementioned arrangement of the diffusers is found to be of advantage in connection with flight manoeuvres in which the approach air acting on the diffusers 27 is always less satisfactory than that acting on the diffusers 28.

The arrangement ensures in the case of such approach air conditions that a reduction in the heat produced in the flame chamber 9a occurs thus reducing the risk of expulsion of the normal shock from the diffusers 27 with the result that the diffuser becomes subcritical. For this purpose the diffusers 27 are connected solely to the mixing chamber 9b via channels 30, while the diffusers 28 are connected solely to the flame chamber 9a when the flight manoeuvres expected are of the kind with which the diffusers 28 in all cases receive a less satisfactory approach flow than the diffusers 27.

WHAT WE CLAIM IS: 1. A ramjet engine for operation with a fuel formed from a combustible propulsive charge containing boron or similar high-energy particulate material with high melting and high vaporisation points, the engine having air inlet diffusers distributed around the periphery of the engine body and serving to feed ram air to a combustion chamber of the engine also fed with the fuel formed from combustion of the charge, the total ram air flow from the inlet diffusers being subdivided into combustion air which is fed to a forward combustion chamber zone forming a flame chamber and into mixing air fed to a rearward combustion chamber zone serving as a mixing chamber.

2. A ramjet engine in accordance with Claim 1, wherein the inlet diffusers are each followed in the direction of flow by a channel for mixing air located in an outer casing of the body, the channel feeding the mixing chamber and including lateral apertures whereby air may flow into the flame chamber.

3. A ramjet engine in accordance with Claim 2, wherein the mixing air channels each include a venturi nozzle.

4. A ramjet engine in accordance with Claim 2 or 3, wherein one or more of the mixing air channels include a throttle means whereby the quantity of mixing air flowing therethrough may be adjusted.

5. A ramjet engine in accordance with Claim 2, wherein supersonic air inlet diffusers have a separating plate extending through the subsonic zone into the supersonic zone and serving to subdivide the diffuser generally longitudinally into a combustion air part communica.ing only with the apertures feeding air to the flame chamber and into a mixing air part feeding air only to the channels feeding the mixing chamber.

6. A ramjet engine in accordance with Claim 1, wherein the periphery of the body has separate inlet diffusers for the combustion air and the mixing air.

7. A ramjet engine in accordance with Claim 6, wherein those inlet diffusers subject to the possibility of a less satisfactory approach air flow due to their position feed the flame chamber while the other diffusers feed the mixing chamber.

8. A ramjet engine contructed and arranged to function substantially as herein described with reference to and as shown in Figures 1 and 2, or 3, or 4 and 5 of the accompanying drawings.

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Claims

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mixing air to be varied.

Referring to Figure 3 this shows a supersonic inlet diffuser 20 which is semi-circular in plan with the outer casing, central core and boundary layer scoop marked 21,22 and 23 respectively.

It differs from the diffuser design shown in Figures 1 and 2 inasmuch as the provision of a venturi nozzle in the mixed air channels is dispensed with. Instead the interior of the diffuser is subdivided into a combustion air sector 24 communicating with the flow apertures 14 in Figure 2 for the flame chamber 9a and into a mixing air sector 25 communicating with the mixing air channels 16 in Figure 2 for the mixing chamber 9b. A separating plate 26 extends through the subsonic zone into the supersonic zone. This results in the separation of the flow of combustion air from the flow of mixing air and this in contrast to the method of separation mentioned earlier using a venturi nozzle, eliminates further pressure losses.It also makes it possible by using a simple slide valve to vary the proportions of the combustion and the mixing air and thus the development of heat in the flame chamber as required. This may be necessary as a result of a tendency for compressive surges to occur in the event of excessive angles of incidence for the projectile and unsatisfactory approach air flow to the diffuser.

The propulsion unit construction shown in Figures 4 and 5, as in Figures 1 and 2, is provided with four supersonic inlet diffusers 27 and 28, which are distributed around the periphery of the projectile 1 in the manner shown in Figure 1. The difference in the present case is that one pair of diametrically opposite dif- fusers 27 are connected through the apertures 29 to the flame chamber 9a only, while the other pair of diametrically opposite diffusers 28 are connected through channels 30 to the mixing chamber 9b only.

The aforementioned arrangement of the diffusers is found to be of advantage in connection with flight manoeuvres in which the approach air acting on the diffusers 27 is always less satisfactory than that acting on the diffusers 28.

The arrangement ensures in the case of such approach air conditions that a reduction in the heat produced in the flame chamber 9a occurs thus reducing the risk of expulsion of the normal shock from the diffusers 27 with the result that the diffuser becomes subcritical. For this purpose the diffusers 27 are connected solely to the mixing chamber 9b via channels 30, while the diffusers 28 are connected solely to the flame chamber 9a when the flight manoeuvres expected are of the kind with which the diffusers

28 in all cases receive a less satisfactory approach flow than the diffusers 27.

WHAT WE CLAIM IS: 1. A ramjet engine for operation with a fuel formed from a combustible propulsive charge containing boron or similar high-energy particulate material with high melting and high vaporisation points, the engine having air inlet diffusers distributed around the periphery of the engine body and serving to feed ram air to a combustion chamber of the engine also fed with the fuel formed from combustion of the charge, the total ram air flow from the inlet diffusers being subdivided into combustion air which is fed to a forward combustion chamber zone forming a flame chamber and into mixing air fed to a rearward combustion chamber zone serving as a mixing chamber.
2. A ramjet engine in accordance with Claim 1, wherein the inlet diffusers are each followed in the direction of flow by a channel for mixing air located in an outer casing of the body, the channel feeding the mixing chamber and including lateral apertures whereby air may flow into the flame chamber.
3. A ramjet engine in accordance with Claim 2, wherein the mixing air channels each include a venturi nozzle.
4. A ramjet engine in accordance with Claim 2 or 3, wherein one or more of the mixing air channels include a throttle means whereby the quantity of mixing air flowing therethrough may be adjusted.
5. A ramjet engine in accordance with Claim 2, wherein supersonic air inlet diffusers have a separating plate extending through the subsonic zone into the supersonic zone and serving to subdivide the diffuser generally longitudinally into a combustion air part communica.ing only with the apertures feeding air to the flame chamber and into a mixing air part feeding air only to the channels feeding the mixing chamber.
6. A ramjet engine in accordance with Claim 1, wherein the periphery of the body has separate inlet diffusers for the combustion air and the mixing air.
7. A ramjet engine in accordance with Claim 6, wherein those inlet diffusers subject to the possibility of a less satisfactory approach air flow due to their position feed the flame chamber while the other diffusers feed the mixing chamber.
8. A ramjet engine contructed and arranged to function substantially as herein described with reference to and as shown in Figures 1 and 2, or 3, or 4 and 5 of the accompanying drawings.