DE9003781U1 - Combustion chamber with nozzle for a hypersonic engine - Google Patents

Description

\ .: :!..·..:: : 29.03.i990, 2422A

10895 GM

Combustion chamber with nozzle for a hypersonic drive

The innovation concerns a combustion chamber with a variable nozzle and with a variable combustion chamber length for a hypersonic drive, according to the generic term of claim 1.

One such combustion chamber is, for example, part of the hypersonic engine described in B-PS 805 *18. This drive comprises a "five-cylinder engine with a turbofan, which is driven by an air-independent turbine (turborocket). The turbine housing is made of aluminum . They can be fed directly into the open air or into a combustion chamber connected downstream of the compressor. This is where the afterburning of the exhaust gases, which usually still contain unburned fuel, takes place. The combustion chamber designated with position 38 can also work in ramjet mode, for which it is equipped with its own fuel injection device. The compressor runs in "wind mode". The combustion chamber merges into a thrust nozzle with a variable cross-section at the downstream end. The narrowest nozzle cross-section is changed with the help of an axially movable displacement body (position 43) arranged at the end of a support tube. A movement of the displacement body towards the combustion chamber inlet causes an increase in the nozzle throat cross-section, a movement towards the nozzle outlet causes a reduction. In the process - more or less unintentionally - the effective combustion chamber length and thus the combustion chamber volume are also changed without to take into account adverse effects such as a deterioration in the efficiency of the combustion chamber.

In contrast, the task of the innovation is to create a combustion chamber with a variable nozzle and with a variable combustion chamber length, which either works as a ramjet combustion chamber or combined as a ramjet combustion chamber and afterburning chamber of a turbo engine, in which the dependency between the variation of the nozzle throat cross-section and the variation of the combustion chamber length and thus of the combustion chamber volume is so

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selected so that a beneficial influence on the combustion chamber efficiency and the thermal combustion chamber load is achieved.

This task is solved by the features characterized in claim 1.
Features solved.

The inner contour of the outer wall delimiting the flow channel expands continuously in the direction of flow from a plane which is given approximately by the innermost position of the largest cross-section of the displacer, i.e. the position closest to the combustion chamber inlet. Thus, the nozzle neck cross-section, i.e. the narrowest cross-section between the displacer and the outer wall, also expands continuously when the displacer is displaced in the direction of flow. In other words, an axial displacement of the displacer causes a
Simultaneous change in the combustion chamber length, the combustion chamber volume and the nozzle neck cross-section.

The switchover Mach number, i.e. the flight speed at which the engine switches from turbo mode to ramjet mode, is expected to be around Ma - 3.5.

The ramjet operated at this relatively low speed
is characterized by the following features:

Minimum combustion chamber pressure,
minimum combustion chamber temperature,
maximum combustion chamber flow Mach number,
·> maximum reaction time and combustion length.

According to the innovation, this mode of operation corresponds to the outermost position of the displacement body with maximum combustion chamber length, maximum combustion chamber volume and maximum nozzle throat cross-section. Due to the low thermal load, the large combustion chamber surface is not a problem.

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The maximum flight speed is, for example, Ma - 7.

Ramjet operation at this high speed has the following characteristics:

Maximum combustion chamber pressure,
maximum combustion chamber temperature,
minimum combustion chamber flow Mach number, -> minimum reaction time and combustion length.

According to the innovation, this mode of operation corresponds to the innermost position of the displacer with minimal combustion chamber length, minimal combustion chamber volume and minimal nozzle neck cross-section. The resulting small combustion chamber surface minimizes the absolute thermal load on the combustion chamber, which simplifies its cooling and increases its stability.

By means of the geometric conditions, in particular the length of the adjustment path of the displacement body and the course of the inner contour of the outer wall 1m adjustment range, a very good adaptation to the flow and combustion technology requirements is possible.

Even when the combustion chamber is operated as an afterburner chamber for a turbo engine, the new adjustability enables better adaptation to the operating conditions.

Claims 2 to 4 contain preferred embodiments of the combustion chamber according to claim 1.

The innovation is explained in more detail below using the drawings. These show a schematic representation:

Fig. 1 a longitudinal section through a ramjet combustion chamber with variable nozzle and with air supply through a pipe,

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Fig. 2 a longitudinal section through a ramjet combustion chamber with air supply through a double tube,

Fig. 3 is a partial longitudinal section through the combustion chamber according to Fig. 2 in a plane rotated by 90°,

Fig. 4 a cross section along line IV-IV in Fig. 3, Fig. 5 a cross section along line V-V in Fig. 3,

Fig. 6 shows a longitudinal section through a combustion chamber that functions as a combined ramjet combustion chamber and as an afterburner chamber of a turbo engine with an inlet cross-section for air and exhaust gas,

Fig. 7 is a partial longitudinal section through two combined combustion chambers, the one shown above in the middle being connected to a turbojet engine and the one shown below in the middle being connected to a turborocket engine.

The combustion chamber 1 with nozzle 7 according to Fig. 1 is intended exclusively for ramjet operation. It forms the essential element of a ramjet engine, which can be connected in parallel to one or more turbo engines as part of a hypersonic drive and possibly to one or more other ramjet engines. Furthermore, there can be engines that are independent of the outside air for operation in an airless space, such as liquid rocket engines.

For horizontally launched two-stage space transporters, it is assumed that more than half of the flight time of the lower stage is spent in ramjet mode, and that about 2/3 of the fuel of the lower stage is consumed in the process. This alone shows how important it is for the ramjet engines, i.e. in particular the ramjet combustion chambers, to function optimally.

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The arrangement according to Mg. 1 is designed to be rotationally symmetrical for various reasons (strength, thermal load, flow losses, etc.), which is not absolutely necessary. Asymmetrical configurations or those with other symmetries (square or polygonal, etc.) are also conceivable. In the spirit of the innovation, it is only that the combustion chamber length and the exhaust neck cross-section are changed in the same direction.

(The combustion chamber 1 is surrounded by the solid outer wall 10.

which 1m front part cylindrical, 1m adjustment area of the nozzle

j 7 bell-shaped, extended, 1st.

To vary the length of the combustion chamber and the nozzle neck cross-section, an axially displaceable displacement body 15 is provided (drive not shown), which is mounted on a support tube 18. The support tube 18 extends from the rear wall 25 of a central body 23 at the combustion chamber inlet. The conical central body 23 is supported by radial struts 28, which carry fuel injection elements 34, and by radial struts 29, firmly in the outer wall 10 or in the connecting housing 46. The latter forms the connection between the pipe 49 leading from the air inlet (not shown) and the fuel inlet. The inlet cross-section 39 has the shape of a circular ring, which is only interrupted by the struts 28, which may be designed as flame holders. The part of the rear wall 25 that is located radially outside the support tube 18 forms - together with the support tube root - a recirculation zone for flame retention, which allows stable, low-vibration combustion to be achieved. If other, sufficient measures for flame retention are taken (e.g. struts 28 as flame holders), the transition from the central body to the support tube can be designed to be flow-optimized (without tearing off), which allows the pressure loss caused by the recirculation zone to be eliminated or reduced. In particular, the largest diameter of the central body can be selected to be the same as the support tube diameter.

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The displacer body 15 is shown in two positions, namely in solid lines in its outermost (rearmost) position and in dashed lines in its innermost (front) position. The information L ₂ and L 3 each shows the theoretical combustion chamber length, which extends from the injection plane to the front (upstream) end of the displacer body 15. The plane E, from which the flow channel delimited by the outer wall 10 widens in the direction of flow, is located axially approximately where the largest diameter D of the displacer body is in the innermost position (combustion chamber length L 1 ). Thus - starting from L ₁ - any increase in the combustion chamber length and thus the combustion chamber volume also causes an increase in the nozzle neck cross-section. The desired relationship between combustion chamber length and nozzle neck cross-section can be influenced via the inner contour of the outer wall 10. In addition to the bell-shaped curve shown, a conical curve or a curve that widens towards the rear in a funnel shape are also conceivable.

As already explained, the combustion chamber length L 1 (short) corresponds to the conditions at the highest flight Mach number and the highest combustion chamber load, the combustion chamber length L ₂ (long) corresponds to the conditions at the lowest flight Mach number and the lowest combustion chamber load. Any intermediate position between these extremes is possible.

It should be noted that suitable fluid-mechanical extensions of the nozzle contour can be added downstream and in extension of the outer wall 10 shown, which, however, have no influence on the nature of the present innovation.

The design according to Fig. 2 also relates to a combustion chamber 2 exclusively for ramjet operation. In contrast to Fig. 1, the combustion chamber 2 is supplied with ram air via a double pipe 50. As the vertical section in Fig. 2 shows, the double pipe inflow offers advantages with regard to the height of the connecting area between the air inlet and combustion chamber, which can have a positive effect, especially when arranged in the wing area. As the horizontal section according to Fig. 3 shows, the double pipe 50 is inevitably wider than the single pipe 49 according to Fig. 1.

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The structure of the combustion chamber 2 from the circular inlet cross-section 40 to the downstream end of the displacer body 16 is practically identical to that in Fig. 1. Thus, the outer wall 11, the displacer body 16, the support tube 19 and the rear wall 26 of the central body 24 are also rotationally symmetrical. For the sake of simplicity, the axially movable displacer body 16 is only shown in its outermost (rearmost) position. The connecting housing 47 brings about the transition from the double circular cross-section of the double tube 50 to the single, large circular cross-section of the outer wall 11. This transition can be understood from Figs. 3 to 5, with Fig. 4 showing the section along line IV - IV. Fig. 5 shows the section along line V - V in Fig. 3. The components located behind the cutting planes are not shown in Figs. 4 and 5 for the sake of clarity.

Fig. 4 shows the double-circle cross-section of the double pipe 50, to which the connecting housing 47 is connected.

As shown in Figures 3 and 5, the outer contour of the connecting housing 47 in its cross-section over the vast majority of its length consists of two mirror-symmetrically adjacent circles, the imaginary centers of which approach the plane of symmetry more in the direction of flow, whereby their diameter increases. In the area of the inlet cross-section 40 of the combustion chamber 2, two halves of a circle, i.e. a full circle, with the diameter of the outer wall 11 form the outlet cross-section of the connecting housing 47, while in the connection area to the double pipe 50 two tangent, small full circles form its inlet cross-section.

The positionally symmetrical, conical central body 24 is firmly connected to the connecting housing 47 on its top and bottom over its entire length via an intermediate wall. The intermediate walls 48 also prevent the connecting housing 47 from bulging under internal pressure (suspended forces). In the area of the circular inlet cross-section 40, the central body 24 is additionally supported by radial,

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The fuel tank is supported by struts 30 distributed around the circumference. At least some of the struts 30 and possibly also the intermediate walls 48 carry fuel injection nozzles on their rear edges. If necessary, the stems are designed as fuel holders.

Fig. 6 shows a combustion chamber 3 which is operated as a combined ramjet combustion chamber and as an afterburner chamber of a turbo engine, the first case being shown below the center line, the second case being shown above the center line.

The actual combustion chamber 3 with nozzle 9, which is again rotationally symmetrical in the case shown, comprises the outer wall 12, the movable displacement body 17 with support tube 20, the ram-shaped wall 27 and the struts 31 supporting these with fuel injection elements 36. There is only one common, annular inlet cross-section 41 for the ram air, the exhaust gases and possibly the exhaust air of the turbo engine. In ramjet operation, an axially movable slide 55 closes the outlet 54 of the - not visible - turbo engine and forms a bypass channel 51 for the ram air with the outer wall. In turbo mode, the slide 55 opens the outlet 54, whereby the combustion chamber 3 can be operated as an afterburning chamber and fuel is supplied through the injection elements 36.

Finally, Fig. 7 shows two combustion chambers 4 and 5, both of which are used in combination as ramjet combustion chambers and as afterburning chambers of a turbo engine.

Combustion chamber 4 (upper half of the figure) is connected to a turbojet engine 56, preferably one with a small bypass ratio (turbojet), combustion chamber 5 (lower half of the figure) is connected to a turborocket engine 57. Instead of the turborocket engine 57, a turboexpander engine (heat exchanger instead of combustion chamber) can also be used.

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The main difference from Fig. 6 is that both the combustion chamber 4 and the combustion chamber 5 have two inlet cross-sections for the ram air on the one hand and the turbo engine exhaust gases on the other. Only the inlet side area of the combustion chamber 4 is represented by the outer wall 13, support tube 2], struts 32 and fuel injection elements 37. The bypass channel 52 for the ram air concentrically surrounds the turbo jet engine 56 and leads to the circular, outer inlet cross-section 42. During turbo operation, the bypass air of the turbo jet elevator 56 can be guided through the bypass channel 52. Of the latter, only the turbine 58 (with bearing) is shown. The exhaust gases flow into the combustion chamber 4 via the circular inner inlet cross-section 43. The fuel injection elements 37 can be arranged in series between the two inlet cross-sections (case shown).

Of combustion chamber 5, only the inlet area with support tube 22, outer wall 14, struts 33 and fuel injection elements 38 is also shown. The bypass channel 53 is intended exclusively for the ram air and leads to the external inlet cross section 44. The low-pressure compressor 60 of the turbo rocket engine 57 is driven by an externally air-independent turbine 59 via a reduction gear (not shown). The turbine propellant gases are generated in the combustion chamber 6 housed in the support tube 22. The turbine exhaust gases pass through the channel deflection 61 to the mixer 62, where they are combined with the exhaust air of the low-pressure compressor 60. The exhaust-side hand of the mixer 62 has a periodically changing radius in the circumferential direction, so that its cross-section resembles a gear or a flower. The inner inlet cross-section 45 is connected to the mixer 62. The fuel injection elements 38 can be arranged in the area of one or both inlet cross-sections (case shown).

Claims (4)

&Ggr; : : : : 29.03.1990, 2422&Agr; : ' : Pf/Th " " 10895 GM Combustion chamber with nozzle for a hypersonic drive Claims 1. Combustion chamber with variable nozzle and with variable combustion chamber length for a hypersonic drive, which works either as a ram combustion chamber or combined as a ram combustion chamber and afterburner chamber of a turbo engine, with a fixed outer wall delimiting the flow channel on the circumference, with an axially displaceable displacer body arranged in the flow channel and with a support tube holding the displacer body and extending from the combustion chamber inlet _} "characterized in that the flow channel delimited by the outer wall (10, 11, 12, 13, 14) expands continuously in the direction of flow from a plane (E) which lies close to or in the plane which is given by the innermost position of the largest cross-section (diameter D) of the displacer body (15, 16, 17). 2. Combustion chamber according to claim 1, which operates exclusively as a ramjet combustion chamber, with an outer wall that is rotationally symmetrical at least 1 m in the inlet area, characterized in that the support tube (18, 19) for the displacement body (15, 16) is arranged centrally on the flat rear wall (25, 26) of a rotationally symmetrical central body (23, 24) that delimits the combustion chamber to the front, which forms a circular inlet cross-section (39, 40) for the ram air with the outer wall (10, 11) and which is connected to the outer wall (10, 11) via radial struts (28, 30) that carry fuel injection elements (34, 35) and can be designed as flame retainers. 19. 03.1990, 24??A Pf/lh 10895 GM 3. Combustion chamber according to claim 1, which works in combination as a ramjet combustion chamber and afterburning chamber of a turbo engine, with an outer wall that is rotationally symmetrical at least in the inlet area, characterized in that the support tube (20) for the displacement body (17) is arranged on a flat wall (27) with a round outer contour that delimits the combustion chamber to the front, which forms with the outer wall (12) an annular inlet cross-section (41) for the exhaust gas and the exhaust air of the turbo lift and for the ram air, in the area of which fuel injection elements (36) are arranged. 4. Combustion chamber according to claim 1, which works in combination as a ramjet combustion chamber and afterburning chamber of a turbo engine, with an at least in the inlet area rotationally symmetrical outer wall, characterized in that the support tube (21, 22) for the displacement body is arranged at the downstream structural end of the turbo lift (turbo jet engine 56, turbo rocket engine 57) or on a structure adjacent thereto, with an inner circular inlet cross section (43, 45) for the exhaust gas or for the exhaust gas and the exhaust air of the turbo engine and an outer circular ring-shaped inlet cross section (43, 45) for the exhaust gas or for the exhaust gas and the exhaust air of the turbo engine between the support tube (21, 22) and the outer wall (13, 14). an inlet cross-section (42, 44) adjacent to the outer wall (13, 14) for the ram air or for the ram air and part of the exhaust air of the turbo lift are present, and wherein fuel injection elements (37, 38) are arranged at least in the region of one of the two inlet cross-sections (42, 43; 44, 45).