Multistage turbomachine compressor
View Patent ↗A multistage compressor for a turbomachine, in particular an airplane turboprop or turbojet, the compressor comprising a double-walled casing having an inner wall made up of shrouds surrounding respective annular rows of moving blades and annular rows of straightening stator vanes, said shrouds being connected to the outer wall of the casing by independent suspension means enabling the radial clearances between the outer ends of the moving blades and the shrouds of the inner wall of the casing to be adjusted independently from one compression stage to another.
1. A multistage turbomachine compressor comprising annular rows of moving blades rotating inside a double-walled casing, and annular rows of straightening stator vanes carried by an inner wall of the double-walled casing, wherein the inner wall of the casing comprises a plurality of shrouds substantially in end-to-end alignment and suspended independently of one another from the outer wall of the casing by suspension members that are flexible or deformable, some of the shrouds each surrounding an annular row of moving blades and other shrouds each carrying an annular row of straightening stator vanes, said suspension members having a substantially C-shape cross section and being configured to adjust the radial clearances between radially outer-ends of the moving blades and the corresponding shrouds independently of one another.
2. A compressor according to claim 1 , wherein each annular row of moving blades is surrounded by a shroud suspended from the outer wall of the casing independently from the other shrouds surrounding the other annular rows of moving blades.
3. A compressor according to claim 1 , wherein two shrouds of the inner wall of the casing, surrounding two successive annular rows of moving blades are securely connected together, the moving blades of these two successive rows having the same dimensions.
4. A compressor according to claim 3 , wherein a shroud carrying an annular row of straightening stator vanes situated between the two shrouds surrounding annular rows of moving blades is carried by said two shrouds and is securely connects them together.
5. A compressor according to claim 1 , wherein at least one of the shrouds surrounding an annular row of moving blades is adjacent to a second shroud carrying an annular row of straightening stator vanes and is connected in gastight manner to said second shroud.
6. A compressor according to claim 5 , wherein the second shroud is connected to a third shroud surrounding an annular row of moving blades by means for providing sealing against the flow of gas passing through the compressor.
7. A compressor according to claim 6 , wherein the second shroud is connected to the outer wall of the casing by means for suspending said second shroud which are independent of the suspension means for suspending the other shrouds.
8. A compressor according to claim 5 , wherein the second shroud is axially spaced apart from a third shroud surrounding an annular row of moving blades and is connected by means for suspending said second shroud to the outer wall of the casing.
9. A compressor according to claim 8 , wherein annular clearance axially separating the second shroud and the third shroud constitutes a passage allowing air to penetrate into a cavity that is formed in the casing between the means for suspending the second shroud and the means for suspending the third shroud, and into which there open out means for taking off air, which are carried by the outer wall of the casing.
10. A compressor according to claim 1 , wherein the radial modes of vibration of at least some of the shrouds surrounding the annular rows of moving blades are adjusted independently from one shroud to another, in order to optimize the radial clearances between said shrouds and the annular rows of moving blades rotating inside said shrouds.
11. An airplane turbojet or turboprop, including a compressor of the type defined in claim 1 .