Compressor for an aircraft engine
AssigneeRolls-Royce Deutschland Ltd & Co KG
InventorErik Johann
Life of this patent
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Abstract
On a compressor with compressor blades, a flow transition fixation mechanism (4) is provided on the suction side (2), approximately parallel to the leading edge (3) and upstream of the compression shocks acting upon the blade, which prevents the transition point from the laminar to the turbulent boundary layer flow from oscillating, thus suppressing oscillation of the compression shocks and their coupling effect with the natural frequencies of the compressor blades.
Claims34
An aircraft engine compressor having compressor blades attached to a compressor disk, which compressor blades, under operating conditions, are loaded by natural frequencies and by compression shocks on suction sides thereof at a certain distance from leading edges thereof, with a flow transition fixation mechanism being provided on the suction side of each compressor blade in a portion of the compressor blade located upstream of an area of action of the compression shocks essentially vertical to a boundary layer flow extending at a compressor blade surface which limits oscillation of a flow transition point between an initially laminar end subsequently turbulent boundary layer flow and, thus, oscillation of the compression shocks, and suppresses their reaction on first and second bending modes as well as on a first torsional mode of the compressor blade occurring under certain conditions of flight.
A compressor blade for an aircraft engine compressor, which, under operating conditions, is loaded by natural frequencies and by compression shocks on a suction side thereof at a certain distance from a leading edge thereof, with a flow transition fixation mechanism being provided on the suction side in a portion of the compressor blade located upstream of an area of action of the compression shocks essentially vertical to a boundary layer flow extending at a compressor blade surface which limits oscillation of a flow transition point between an initially laminar and subsequently turbulent boundary layer flow and, thus, oscillation of the compression sbocks, and suppresses their reaction on first and second bending modes as well as on a first torsional mode of the compressor blade occurring under certain conditions of flight.
A compressor blade for an aircraft engine compressor, comprising a flow transition fixation mechanism on a suction side thereof in a portion of the compressor blade located upstream of an area of action of compression shocks on the compressor blade essentially vertical to a boundary layer flow extending at a compressor blade surface, the flow transition fixation mechanism limiting oscillation of a flow transition point between an initially laminar and subsequently turbulent boundary layer flow and, thus, oscillation of the compression shocks.
A compressor blade for a gas torbine, comprising a flow transition fixation mechanism on a suction side thereof in a portion of the compressor blade located upstream of an area of action of compression shocks on the compressor blade essentially vertical to a boundary layer flow extending at a compressor blade surface, the flow transition fixation mechanism limiting oscillation of a flow transition point between an initially laminar and subsequently turbulent boundary layer flow and, thus, oscillation of the compression shocks.
Dependencies are inferred from the claim text.