Thrust chamber device and method for operating a thrust chamber device

An engine for a flying object or an aircraft, comprising:

• a thrust chamber device, which comprises:
• a thrust chamber with a thrust space having a first portion, a second portion adjacent thereto, and a third portion adjacent to the second portion,
• the thrust space being delimited in all three portions by an outer nozzle wall having an outer thrust space surface, which outer thrust space surface continuously tapers, over an entire extent of the first and second portions, toward the third portion, widens in the third portion away from the second portion, and has a narrowest point at the transition from the second portion to the third portion,
• the first portion being delimited by an inner nozzle wall with an inner thrust space surface, which continuously tapers, over an entire extent of the first portion, toward the second portion, an annular combustion chamber being formed between the inner thrust space surface and the outer thrust space surface and extending over the first portion,
• the first portion being further delimited on a side facing away from the second portion by an injection wall, which connects the inner nozzle wall and the outer nozzle wall to one another,
• a plurality of first propellant inlets for a first propellant component and a plurality of second propellant inlets for a second propellant component,
• wherein:
• the outer nozzle wall comprises at least a part of the plurality of first propellant inlets,
• the plurality of second propellant inlets are arranged or formed only in the injection wall,
• the inner nozzle wall and the outer nozzle wall have only first propellant inlets, and
• the outer thrust space surface is continuously convexly curved or substantially convexly curved in a direction of the inner thrust space surface over an entire extent of the outer thrust space surface.

A flying object or aircraft, comprising:

• a first propellant store for at least one first propellant component,
• a second propellant store for at least one second propellant component, and an engine comprising a thrust chamber device, the thrust chamber device comprising:
• a thrust chamber with a thrust space having a first portion, a second portion adjacent thereto, and a third portion adjacent to the second portion,
• the thrust space being delimited in all three portions by an outer nozzle wall having an outer thrust space surface, which outer thrust space surface continuously tapers, over an entire extent of the first and second portions, toward the third portion, widens in the third portion away from the second portion, and has a narrowest point at the transition from the second portion to the third portion,
• the first portion being delimited by an inner nozzle wall with an inner thrust space surface, which continuously tapers, over an entire extent of the first portion, toward the second portion, an annular combustion chamber being formed between the inner thrust space surface and the outer thrust space surface and extending over the first portion,
• the first portion being further delimited on a side facing away from the second portion by an injection wall, which connects the inner nozzle wall and the outer nozzle wall to one another, and
• a plurality of first propellant inlets for a first propellant component and a plurality of second propellant inlets for a second propellant component,
• wherein:
• the outer nozzle wall comprises at least a part of the plurality of first propellant inlets,
• the plurality of second propellant inlets are arranged or formed only in the injection wall,
• the inner nozzle wall and the outer nozzle wall have only first propellant inlets, and
• the outer thrust space surface is continuously convexly curved or substantially convexly curved in a direction of the inner thrust space surface over an entire extent of the outer thrust space surface.

A method for operating a thrust chamber device which comprises a thrust chamber with a thrust space having a first portion, a second portion adjacent thereto, and a third portion adjacent to the second portion, the thrust space being delimited in all three portions by an outer nozzle wall having an outer thrust space surface, which outer thrust space surface continuously tapers, over an entire extent of the first and second portions, toward the third portion, widens in the third portion away from the second portion, and has a narrowest point at the transition from the second portion to the third portion, the first portion being delimited by an inner nozzle wall with an inner thrust space surface, which continuously tapers, over an entire extent of the first portion, toward the second portion, an annular combustion chamber being formed between the inner thrust space surface and the outer thrust space surface and extending over the first portion, the first portion being further delimited on a side facing away from the second portion by an injection wall, which connects the inner nozzle wall and the outer nozzle wall to one another, and a plurality of first propellant inlets for a first propellant component and a plurality of second propellant inlets for a second propellant component, wherein the outer nozzle wall comprises at least a part of the plurality of first propellant inlets and the plurality of second propellant inlets are arranged or formed only in the injection wall,

• wherein:
• the inner nozzle wall and the outer nozzle wall have only first propellant inlets, and
• the outer thrust space surface is continuously convexly curved or substantially convexly curved in a direction of the inner thrust space surface over an entire extent of the outer thrust space surface,
• the method comprising injecting a greater amount of the first propellant component than is necessary for cooling the inner nozzle wall and/or the outer nozzle wall into the thrust space through a part of the plurality of first propellant inlets.