Propulsion Axial Flow Compressor Fan Aerospace Engineering International

Propulsion: Axial Flow Compressor & Fan Aerospace Engineering, International School of Engineering (ISE) Academic year : 2012 -2013 (August – December, 2012) Jeerasak Pitakarnnop , Ph. D. Jeerasak. p@chula. ac. th jeerasak@nimt. or. th October 27, 2012 Aircraft Propulsion 2

Introduction • First rotating component that the fluid encounters. • Basic function: – Impart kinetic energy to the working fluid by means of rotating blades, then – Convert the increase in energy to an increase in total pressure. “The pressure ratio increases, the required fuel flow decreases and the extracted power increases” October 27, 2012 Aircraft Propulsion 3

Introduction • Design of an efficient axial flow fan/compressor is a complex process which often involve success or failure of an engine CFD tools can efficiently be use for complex 3 D analysis and design. • Addition functions: – A small portion of the air is bled to provide some cockpit and electronic environmental control. – A small portion is bled to provide pressurized air for inlet antiicing. – Some of the high-pressure “cool” air is directed to the turbine and used to reduce the temperature of hot turbine blades. October 27, 2012 Aircraft Propulsion 4

Geometry Tip and housing diameter are approximately constant through a compressor. Rotor blades: do the work on fluid. Stationary blades: do not input any energy but necessary for guiding the flow. October 27, 2012 Aircraft Propulsion 5

Geometry Disk October 27, 2012 Aircraft Propulsion 6

Inlet & Exit Guide Vanes • Inlet Guide Vanes (IGV): – Same function as stator vanes but the design is quite different. – They turn the incoming air, which is in the axial direction in a direction of 1 st rotor blades incident free. • Exit Guide Vanes (EGV) – A set of stator vanes after the last stage that readies the flow for entrance to the combustor. – Design to add swirl to the flow, which aids in mixing within the combustor. October 27, 2012 Aircraft Propulsion 7

Compressor Stage Operation A stage: a rotor wheel carrying blades + a stator assembly carrying stationary blades or vanes Proper flow direction from IGV or stator of previous stage Investigation of stage aerodynamics is usually carried out in a cascade tunnel, an experimental setup where single or multi-stage cascades are tested under simulated flow conditions. Conceptual unwrapping of middle section Consider the flow at some midspan blade section (between hub and tip). Absolute velocity V as seen by an external observer standing next to the engine. Circumferential velocity U depending on rotational speed (rpm) and radial position. Relative velocity Vrel as seen by an observer sitting on the rotating blade and moving with it. 3 D effects occur in an actual compressor but for study purposes flow in a cascade is considered to be 2 D. October 27, 2012 Aircraft Propulsion 8

2 D Simulation in CFD October 27, 2012 Aircraft Propulsion 9

Velocity Polygon or Triangles October 27, 2012 Aircraft Propulsion 10

Inlet Guide Vanes The axial flow velocity relative to the engine flame (absolute velocity) at the IGV inlet is c 0. The inlet flow to IGV is typically aligned with the axis of the engine Exit blade angle relative to the axis of the engine. . The absolute flow velocity (velocity in the non rotating flame) at the IGV exit is c 1. October 27, 2012 Exit flow angle relative to the axis of the engine. . Aircraft Propulsion 11

1 st Stage Rotor Relate the velocities in the stationary ref. frame to those in rotating frame Resulting velocity in a rotating frame w 1. has a flow angle β 1. If β 1 = β’ 1 incidence angle is 0. Difficult to happen in offdesign conditions The difference at tailing edge β 2 and β’ 2 is called the deviation. Abs. vel. c 1 has component in tangential direction cu 1 In axial direction ca 1 Rel. vel. w 1 has component in tangential direction wu 1 In axial direction wa 1 October 27, 2012 Aircraft Propulsion 12

1 st Stage Stator The relative flow velocity (velocity in the rotating flame) at the rotor exit is w 1. The absolute flow velocity (velocity in the stationary flame) at the stator inlet is c 1. The absolute flow velocity (velocity in the stationary flame) at the stator exit is c 2. October 27, 2012 Aircraft Propulsion 13

2 nd Stage Rotor October 27, 2012 Aircraft Propulsion 14

2 nd Stage Stator October 27, 2012 Aircraft Propulsion 15

Side View of First Stage Blade Height Radius of the blade October 27, 2012 Aircraft Propulsion 16

Single-Stage Energy Analysis • Relate the velocity from polygons to the pressure rise and other • Axial Flow Compressor component Trends October 27, 2012 Aircraft Propulsion 17

Total Pressure Ratio The equations is derived for a single stage (rotor and stator) using 2 D planar mean line c. v. approach. “Midway between hub and tip” • Power Input to the Shaft • Total Pressure Ratio of the Stage Control Volume definition for compressor stage October 27, 2012 Aircraft Propulsion 18

Percent Reaction A relation that approximates the relative loading of the rotor and stator based on the enthalpy rise: October 27, 2012 Aircraft Propulsion 19

Incompressible Flow For comparison, the pressure rise and percent reaction of a turbomachine with an incompressible fluid can be found from the following equations: • Power Input to the Shaft • Total Pressure Rise of the Stage • Percent Reaction October 27, 2012 Aircraft Propulsion 20

Relationships of Velocity Polygons to Percent Reaction and Pressure Ratio October 27, 2012 Aircraft Propulsion 21

Relationships of Velocity Polygons to Percent Reaction and Pressure Ratio October 27, 2012 Aircraft Propulsion 22

Limit on Stage Pressure Ratio • The rotor is moving, the relative velocity must be used: • For the stator, which is stationary the relative velocity must be used: 1 and 2 refer to the stage inlet and midstage properties. October 27, 2012 Aircraft Propulsion 23

Limit on Stage Pressure Ratio Rotor October 27, 2012 Stator Aircraft Propulsion 24

Ex 1 : Velocity Polygon A stage approximating the size one of the last stages (rotor and stator) of a high-pressure compressor is to be analyzed. It rotates at 8000 rpm and compresses 127 kg s-1 of air. The inlet pressure and temperature are 1. 875 MPa and 727. 6 K, respectively. The average radius of the blades is 335. 28 mm and the inlet blade height is 31. 496 mm. The absolute inlet flow angle to the rotor is the same as the stator exit flow angle 15°, and the rotor flow turning angle is 25°. The stage has been designed so that the blade height varies and the axial velocity remains constant through it. The efficiency of stage is 90%. The value of cp and γ are 1. 08311 k. J kg-1 K-1 and 1. 361, respectively, which are based on the resulting value of T 2 or average static temperature of the stage. The following details are to be found: blade heights at the rotor and stator exits, the static and total pressures at the rotor and stator exits, the required power for the stage, and the percent reaction for the stage. October 27, 2012 Aircraft Propulsion 25