Combustion module for a turbomachine

A combustion module for a turbomachine of an aircraft, the combustion module comprising:

• an annular casing extending around a longitudinal axis;
• an annular combustion chamber located inside the annular casing and having coaxial annular walls, respectively being an internal wall and an external wall connected to each other by an annular bottom chamber, wherein:
• the internal wall, the external wall, and the annular bottom chamber are formed by at least two annular envelopes;
• the at least two annular envelopes are made of a ceramic matrix composite material and have annular edges fitted together one inside the other and in axial abutment on each other;
• and
• the at least two annular envelopes include a first envelope and a second envelope;
• and
• anti-disengagement devices configured to maintain the annular edges in axial abutment with each other, the anti-disengagement devices being carried by the annular combustion chamber and/or the annular casing,
• wherein the anti-disengagement devices are abutment members which are projectably formed on the annular casing and each of the abutment members comprises a free end configured to come in abutment on the combustion chamber,
• wherein the annular casing carries an annular row of fuel injectors angularly distributed around the longitudinal axis, and wherein each of the fuel injectors extends along an axis perpendicular to the longitudinal axis,
• wherein the fuel injectors of the annular row of fuel injectors are engaged in orifices of the first envelope of the at least two annular envelopes,
• wherein the anti-disengagement devices comprise first anti-disengagement devices being the abutment members, and second anti-disengagement devices having washers,
• wherein the washers are mounted around the fuel injectors and tightened radially against the first envelope, and
• wherein the washers have lugs configured to cooperate by abutment and/or hooking with the second envelope of the at least two annular envelopes.

A combustion module for a turbomachine of an aircraft, the combustion module comprising:

• an annular casing extending around a longitudinal axis;
• an annular combustion chamber located inside the annular casing and having coaxial annular walls, respectively being an internal wall and an external wall, connected to each other by an annular bottom chamber, wherein:
• the internal wall, the external wall, and the annular bottom chamber being formed by at least two annular envelopes;
• the at least two annular envelopes being made of a ceramic matrix composite material and having annular edges fitted together one inside the other and in axial abutment on each other, wherein the annular edges are also made of a ceramic matrix composite;
• the at least two annular envelopes include a first envelope and a second envelope, wherein the first envelope defines the internal wall and the annular bottom chamber, and wherein the second envelope defines the external wall;
• the anti-disengagement devices configured to maintain the annular edges in axial abutment with each other, the anti-disengagement devices being carried by the annular combustion chamber and/or the annular casing;
• the anti-disengagement devices carried by the annular casing are abutment members which are projectably formed on the annular casing and each of the abutment members comprises a free end configured to come directly in abutment on the annular bottom chamber, and
• the abutment members are made of the same material as the annular casing, each of the abutment members having a first thickness that is reduced compared to a second thickness of the annular casing.