Axial compressor

About: Axial compressor is a research topic. Over the lifetime, 12035 publications have been published within this topic receiving 127766 citations.

Effects of Inlet Flow Field Conditions on the Performance of Centrifugal Compressor Diffusers: Part 2—Straight-Channel Diffuser

Abstract: This is Part 2 of an examination of the influence of inlet flow conditions on the performance and operating range of centrifugal compressor vaned diffusers. The paper describes tests of a straight-channel type diffuser, sometimes called a wedge-vane diffuser, and compares the results with those from the discrete-passage diffusers described in Part 1. Effects of diffuser inlet Mach number, flow angle, blockage, and axial flow nonuniformity on diffuser pressure recovery and operating range are addressed. The straight-channel diffuser investigated has 30 vanes and was designed for the same aerodynamic duty as the discrete-passage diffuser described in Part 1. The ranges of the overall pressure recovery coefficients were 0.50-0.78 for the straight-channel diffuser and 0.50-0.70 for the discrete-passage diffuser, except when the diffuser was choked. In other words, the maximum pressure recovery of the straight-channel diffuser was found to be roughly 10 percent higher than that of the discrete-passage diffuser investigated. The two types of diffuser showed similar behavior regarding the dependence of pressure recovery on diffuser inlet flow angle and the insensitivity of the performance to inlet flow field axial distortion and Mach number. The operating range of the straight-channel diffuser, as for the discrete-passage diffusers, was limited by the onset of rotating stall at a fixed momentum-averaged flow angle into the diffuser, which was for the straight-channel diffuser, α crit = 70 ± 0.5 deg. The background, nomenclature, and description of the facility and method are all given in Part 1.

Fully integrated environmental and secondary power system

Abstract: An aircraft subsystem provides the aircraft with all its electrical and conditioned air requirements without requiring the extraction of either shaft power and/or pressurized air from the aircraft's main engines. At the core of this subsystem are two rotating assemblies journaled on non-oil lubricated bearings to a housing. One assembly includes a cooling turbine, a starter/generator, a core compressor, and a high pressure stage of a two stage axial turbine, all mounted on a single shaft. A combustor is disposed between the core compressor and the turbine. The other assembly is comprised of the low pressure stage of the axial turbine coupled to a load compressor via a second shaft. The high pressure turbine stage and cooling turbine are sized to drive the core compressor and starter/generator which provides all the aircraft's electrical needs both on the ground and inflight. The low pressure turbine stage drives the load compressor, which produces pressurized air that is conditioned through heat exchange components and then expanded and cooled across the cooling turbine before being delivered to the aircraft cabin.

Experimental Investigation of the Flow Field in a Transonic, Axial Flow Compressor with Respect to the Development of Blockage and Loss

Abstract: A detailed experimental investigation to understand and quantify the development of loss and blockage in the flow field of a transonic, axial flow compressor rotor has been undertaken Detailed laser anemometer measurements were acquired upstream, within, and downstream of a transonic, axial compressor rotor operating at design and off-design conditions The rotor was operated at 100%, 85%, 80%, and 60% of design speed which provided inlet relative Mach numbers at the blade tip of 148, 126, 118, and 089 respectively At design speed the blockage is evaluated ahead of the rotor passage shock, downstream of the rotor passage shock, and near the trailing edge of the blade row The blockage is evaluated in the core flow area as well as in the casing endwall region Similarly at pm speed conditions for the cases of (1) where the rotor passage shock is much weaker than that at design speed and (2) where there is no rotor passage shock, the blockage and loss are evaluated and compared to the results at design speed Specifically, the impact of the rotor passage shock on the blockage and loss development, pertaining to both the shock/boundary layer interactions and the shock/tip clearance flow interactions, is discussed In addition, the blockage evaluated from the experimental data is compared to (1) an existing correlation of blockage development which was based on computational results, and (2) computational results on a limited basis The results indicate that for this rotor the blockage in the endwall region is 2-3 times that of the core flow region and the blockage in the core flow region more than doubles when the shock strength is sufficient to separate the suction surface boundary layer The distribution of losses in the care flow region indicate that the total loss is primarily comprised of the shock loss when the shock strength is not sufficient to separate the suction surface boundary layer However, when the shock strength is sufficient to separate the suction surface boundary layer, the profile loss is comparable to the shock loss and can exceed the shock loss

Gas turbine engine with improved compressor casing for permitting higher air flow and pressure ratios before surge

Abstract: A plurality of skewed slots of a particular shape are provided within a compressor casing adjacent at least one stage of compressor blade tips, the slots having an axial length greater than that of the adjacent blade tips. The slots are provided such that upon occurrence of compressor surge or stall, the stagnating air occurring about the blade row may be directed by the slots downstream of the compressor blade row back into the main stream of fluid passing through the compressor. By such an arrangement, the slots provide a compressor in which the air flow and pressure ratio may be increased before reaching compressor stall or surge.