The fan, booster, and high-pressure compressor form the compression system. The basic task of the compressor is to add energy to the air stream with minimum loss, and thus raise the temperature and pressure of a specified amount of mass flow rate from State 2 to State 3 in the thermodynamic cycle determined during preliminary de-sign, shown schematically in Figure 3.6.
Generally in a design process the overall pressure ratio and the mass flow rate are considered to be
non-A 4
Figure 3.6 H-S Diagram Showing the Compression Phase of the thermodynamic Cycle
negotiable items because they are principal factors af-fecting the cycle and thrust of die engine.
Designers have selected axial-flow or centrifugal com-pressors to achieve the high compression ratios required for turbojet and turbofan engines rather than piston-type compressors because of weight and size constraints.
This overview will concentrate on axial-flow compres-sors.
The expanded view of an axial-flow compressor system in Figure 3.7 reveals mat the compressor consists of an alternating sequence of rotating and fixed sets of airfoils.
The sets of rotating airfoils are attached to a rotating spool, and the combination is called the rotor. The sets of fixed airfoils are spaced around the periphery of an outer stationary casing, and this combination is called the stator. One set of rotor airfoils (blades) and one set of stator airfoils (vanes) constitutes a stage. The number of stages varies, depending upon the pressure ratio re-quired. Inlet guide vanes (IGVs) are a set of airfoils placed ahead of the first stage rotor to direct the air at the correct angle to this rotor.
A cross-sectional drawing of a typical axial-flow com-pressor including station location numbers along die flowpath is shown in Figure 3.8. A sketch showing the projection of the blading on the casing (outside looking radially inward toward the centerline) is also presented.
Rotation is clockwise, aft, looking forward. Several im-portant terms are used to identify various compressor components.
Casing =the outer wall boundary of a com-pressor or fan, Re
Hub - t h e inner wall boundary of a com-pressor or fan, R H
Radius Ratio =the hub radius divided by the casing radius, R H ' R C
Blades —the rotating airfoils in the fan or compressor
Vanes =the non-rotating airfoils in the fan or compressor
LE = leading edge of blading TE = trailing edge of blading
1.0 = Location of die first stage rotor at LEhub
1.5 = Location of first stage rotor at TE hub
3-6 FANS AND COMPRESSORS
4
Legend
1. Spinner Cone 2. Fan Blades (7 of 38) 3. Fan Forward Case
4. Low Pressure Compressor (LPC) Booster Stator 5. LPC Booster Rotor
6. Fan Shaft
7. Fan Frame, Aft Case and Outlet Guide Vanes 8. High Pressure Compressor (HPC) Stator 9. HPC Rotor
Figure 3.7 Exploded View of the Fan Compressor Assembly
Compressor
tnlet / Compressoi
Discharge
Figure 3.8 Cross-Sectional Drawing of a Typical Axial-Flow Compressor
1.6 1.9
Tip clearance
=Location of first stage stator at LE casing
—Location of first stage stator at TE casing
Stations for the second stage would be: 2.0, 2.5, 2.6, 2.9, where the ' number preceding the decimal
indi-cates the stage number and the fol-- lowing number indicates the
location within the stage
=the radial distance between the air-foil tip and the adjacent flowpath, e The basic principle of operation of a compressor is that kinetic energy (increase in tangential momentum) is im-parted to the air by means of the rotating blades. This en-ergy is then converted to a pressure rise by the diffusion process, which occurs in both the rotor blades and the stator vanes. The process occurs in a singie stage and is repeated by succeeding stages until the desired pressure ratio is attained. The power to drive the compressor is derived from the turbine. An axial-flow compressor can
achieve high pressure ratios (e.g. 23:1) on a single spf at high levels of efficiency.
Vector Diagrams - Vector diagrams are extremely ur ;
fill tools for describing the magnitude and direction : the air velocity at various locations in the compressiu system. They are used extensively by designers as the
"set" the blading.
A velocity vector diagram, shown in Figure 3.9, relatt the absolute velocity, C, the velocity relative to a fra~
of reference fixed to the moving rotor blades, W, and velocity of the reference frame (i.e., wheel speed), «.
Several quantities and relationships are also shown, ii eluding:
W & 0 C&a
= relative velocity and air angle
^absolute velocity and air angle
= relative tangential velocity
=absolute tangential velocity
3-8 FANS AND COMPRESSO
Rotor
Stator
Figure 3.9 Vector Diagrams
u
—wheel speed Trailing edgec
z =axial velocityChord, C A0 = 0,
- f t
=flow turning in the relative frameAor = orj - tt2 =flow turning in the absolute frame
Meanline a or (3 =air angle, the angle between the
air-Flow Turning
stream and the axial direction, a, (absolute) for stators and 0, (rela-tive) for rotors
=the difference between the inlet and exit air angle
Blading - The orientation of the airfoil relative to the vector diagrams and the shape of the airfoil have a domi-nate effect upon compressor efficiency. Airfoils and their orientation in the compressor are shown in Figure 3.10. Important definitions of terms associated with blades in compressors include:
Leading edge =the forward most portion of the air-foil
Thickness envelope
Metal (meanline) angles
Camber, <t>
=the rearward most portion of the airfoil
=the length of the straight line con-necting the leading edge and the trailing edge of a blade or vane
=a line bisecting the suction (convex) and pressure (concave) surfaces of an airfoil
=a distribution specifying the thick-ness at each point on the meanline.
The thickness is applied perpendic-ular to the meanline,
-the angle between the meanline of the blade and the axial direction, or*
for stators and 0* for rotors
= the curvature of the meanline of an airfoil (often as expressed as 0t* -02* for rotors or a j * - aj* for sta-tors), where 1 =LE and 2=TE)
FANS AND COMPRESSORS 3-9
Solidity
Figure 3.10 Blading
Incidence Staggered
angle, i =the difference between the inlet air angle and the leading edge
spacing meanline angle i = /3j - /3j* or a\ - Passage aspect
a,* ratio
Deviation
angle, 6 =the difference between the exit air angle and the trailing edge meanline angle = 02 " &2* o r a2 "
Q2*
Stacking point
Spacing, S =the circumferential distance be-tween two blades or two vanes. S = 2 r/N where N = number of blades or vanes
Solidity, 0 = o = chord/spacing
Aspect ratio = blade height, h, divided by pit-chline chord, c
Stagger
angle, £ = the angle between the chord of a blade or vane and the axial direc-tion
~g = S cos £ (see sketch)
= H/g, blade length (height) divided by staggered spacing (see sketch)
=the point on which the airfoil sec:
tions are stacked to form a blade oi vane
Basic Equations - The basic equations for computin, the temperature rise obtained in a compressor for a giver pressure ratio and efficiency are given in Table 3.1 t equations 3-1 and 3-2 where:
Adiabatic efficiency is the ratio of the ideal enthalpy change to the actual enthalpy change between th same two pressures. This definition has die disadvai tage of varying with pressure ratio, 7jad is usuall;
shown on a compressor performance map.
Polytropic efficiency is the ratio of the energy whic contributes to the pressure rise to the total energy de livered to the compressor. T/p is used to compare oi compressor to another.
3-10 FANS AND COMPRESSOR
BASIC EQUATIONS FOR COMPRESSORS
* Pressure Ratio of Compressor
^ 3 T„ (
Tr3 A A yh -
1TT3 \ T ^ P PT2
PT3
PT2 TT 2 / 7 - 1
3-1
3-2
• Loss Coefficient
pTi n " pSi n
• Reaction
3-5
R = Where 7jad — Adiabatic Efficiency
t]p — Polytropic Efficiency
Change in Tangential Momentum Across Rotor
(APQ) rotor
(APS) stage 3-6
(^rCu)Pitch =
gJCptrATT)p Up
gJCp(ATT)p
up 3-3
• Pressure Ratio In Terms of Flow Turning
P72 f. r w ^Z (t a n a2 ~ t a n ai ) 1 7 / 7 - 1 PT1
- [ ' -
HT . ± ] 3-4Table 3.1 Basic Equations for Compressors
The change in tangential momentum across each rotor can be computed from the temperature rise across the ro-tor using equation 3-3. The pressure ratio achieved can be related to the flow turning across a blade row (c*2 -a]), equation 3-4.
The loss in a compressor is usually expressed in terms of a total pressure loss coefficient, OJ, as defined in equation 3-5. For a rotor blade row, the relative total pressure is used, and for a stator the absolute total pressure is used.
The reaction of a compressor is defined as the static pressure rise occurring across the rotor blade divided by the static pressure rise occurring across the stage, equa-tion 3-6.