06 Engine Component Matching PDF

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Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Types of Aero Engine Intakes

Chapter 6: Engine Component Matching Supersonic Flight Intake Subsonic Flight Intake • • • •

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

•

sharp edges, long arrangement, potentially variable High static pressure rise High total pressure loss causes distorsion to fan and compressor at any flight condition (strong aerodynamic interaction between intake and compression components) Can develop intake buzz, i.e. violent fluctuations in pressure and related shock system

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

2019

6-1

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Variable Supersonic Intake Subsonic Flight Net Thrust

Supersonic Flight at Ma = 2

6-0

• • • •

Net Thrust

Variable Arrangement: • Allows adaption to varying flight conditions

Intake Basic Fluidic Process (1) Operating Condition: High Flight Mach Number

Source: Rick, Gasturbinen und Flugantriebe, 2013

2019

rounded edges, relatively short arrangement Low static pressure rise Low total pressure loss Causes distorsion to fan and compressor at high incidence or cross-wind (moderate aerodynamic interaction of intake and compression components)

• • • 2019

Manages poor subsonic inlet flow in this geometry Manages shock configuration at supersonic inlet flow Removes boundary layers to reduce compression system distorsion Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

6-2

2019

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

6-3

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Intake Basic Fluidic Process (2)

Intake Basic Fluidic Process (3)

6-4

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

6-5

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Compressor Basic Work Process

Combustor Basic Heat Input Process Combustion Process 3 > 4

Source: Rick, Gasturbinen und Flugantriebe, 2013

Compression Process 2 > 3

Compressor Duty: • Do compression work and pressure ratio • Can have variable stator vanes to vary its flow capacity and stage matching

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

2019

Combustor Duty: • Low Mach no. f entry flow • Fuel injection • Fuel/air mixing • Multi-zone combustion to achieve low emissions • Increase total temperature via intensive heat input to flow • Energy balance:

CC

e: Rick, Gasturbinen und Flugantriebe, 2013

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

2019

Source: Rick, Gasturbinen und Flugantriebe, 2013

Operating Condition: Static on Ground

Source: Rick, Gasturbinen und Flugantriebe, 2013

Operating Condition: Low Flight Mach Number

2019

m3 ht3  mf hf  CC HBurn    m4 ht 4

Feeds the engine secondary air system, providing cooling as well as sealing air to primarily locations in the hot section of the engine, combustor and turbine(s) Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

6-6

2019

Sourc

•

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

6-7

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Turbine Basic Work Output Process

„Compressor-to-Turbine“ Secondary Air System (SAS)

Turbine Duty: • Do expansion work and pressure ratio • Receive cooling air from compressor • Drive compressor and auxiliary units of the engine; power balance on common shaft:

Source: Rick, Gasturbinen und Flugantriebe, 2013

Expansion Process 4 > 44

LP Compressor Air HP Compressor Interstage Air HP Compressor Exit Air

PT  PC  PAux Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

2019

6-8

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Establishing Mach No. Similitude: Reduced Massflow

 The aerodynamic state of a flow is similar or is equivalent, if the relation of the occuring flow velocity and the local speed of sound is unchanged or is the same => Mach number has to be constant (Mach number similitude)

Ma 

c  const. a

 In turbomachinery the state of flow is given by two components of the velocity vector: 1. the axial velocity cax defining how much mass flow passes the rotor 2. the blade speed defining how fast the rotor goes

u



Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

2019

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Establishing Mach No. Similitude: Basic Logic

 Therefore, Mach number similitude is captured using 2 Mach numbers: 1. axial Mach no. 2. circumferential Mach no.

SAS main purpose: • Lead compressed air from different sources to bearings (sealingair) and to the rear of the engine to cool hot parts of combustor and turbine • Allow most effective cooling with a minimum of compressed air • Cool life-critical turbine components: bladesand vanes in gaspath and disks underneath the gaspath

w cax u

c

Mau Rotor blade rotating at blade speed u

Maax

As long as these 2 Mach numbers are the same at two flow/operating conditions of the

Axial Mach No.:

6-9

turbomachine, which may feature different inlet pressure and/or temperature, the aerodynamic flow behavior and the performance remains similar. 2019

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

m red  Reduced Massflow = f (Ma) 6 - 10

2019

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

6 - 11

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Establishing Mach No. Similitude: Reduced Spool Speed

Choked Flow Condition • • •

Circumferential Mach No.:

•

Reduced mass flow is a measure for the level of Mach number in a flow channel If that channel has a narrowest area (throat), the highest values will occur there When a Mach number of 1 is reached in the throat, this state is termed „choking“, and effectively means, that no increase in reduced mass flow can be achieved - it is fixed The fluidic state can be described by isentropic equations:

κ p 2 κ  1  Ma 1    pt  κ 1

 red m throat area choked: Ma = 1

nred 

p p t

(also termed „aerodynamic speed“) Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

2019

6 - 12

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

κ 1 T t  Ma  1   2  2 κ  1 T  κ  1

κ 1  κ 1 2 κ   2    m red,max R  κ 1 

p pt

Reduced Speed = f (Ma)

T Ma 1  2  Tt  κ  1

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

2019

6 - 13

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Reduced Mass Flow - Some Questions

Describing Component Performance (1) The aerodynamic performance of engine components in terms of massflow and rotational speed is described using the following parameters, which are a function of Mach number. Assumption is: steady, inviscid flow and perfect gas (κ=konst.)

m Tt ,in mred  pt ,in Correct

Reduced Mass Flow

Wrong m red

--- allows to consistently plot turbomachinery performance in a map, --which is widely valid independent of flight altitude ----- using this parameter Reynolds Number similitude is achieved for ---two operating points falling in the same place in the map

------------ when reduced mass flow is constant in a turbomachine ---------

m T t mred  p A t

Reduced Shaft Speed (aerodynamic speed)

 Tt m pt

nD n  red T t n

usually used, when considering components of different size (area)

usually used, when considering components of same size (area)

usually used, when considering components of different size (diameter)

nred  T t

the Mach number related to blade speed is constant 2019

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

6 - 14

2019

usually used, when considering components of same size (diameter)

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

6 - 15

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Describing Component Performance (2)

𝒏

Using reference values

𝑻

𝒎ሶ 𝒓𝒆𝒅 = 𝒎ሶ ∙

𝑻 𝟏 ∙ 𝒑 𝑨

Using reference values

𝒏𝒄𝒐𝒓𝒓 = 𝒏 ∙

𝒎ሶ 𝒄𝒐𝒓𝒓

𝟏

𝑻𝒕 ൘𝑻 𝒓𝒆𝒇

𝑻𝒕 ൘𝑻 𝒓𝒆𝒇 = 𝒎ሶ ∙ 𝒑 𝒕 ൗ𝒑 𝒓𝒆𝒇

𝟏 𝒔

𝒌𝒈 𝒔

corrected speed

corrected massflow

• The advantage of this is, that 𝑚ሶ 𝑐𝑜𝑟𝑟 has the unit of a mass flow. Usually the standard see level conditions are used: 101,3 kPa, 288,15 K

Compressor Total Pressure Ratio

Very often, aerodynamic speed and reduced mass flow are considered using inlet reference values for pressure and temperature >> „corrected“ quantities: 𝒏𝒓𝒆𝒅 =

Compressor Map

p C  t 3 pt 2 2

Compressor

Blue: performance map of the compressor (determined by compressor itself) Red: operating line shown in the compressor map (determined by „the system“ around the compressor: choked turbine or other throttle devices)

3

DP

Inlet Reduced Massflow

• Area A is constant in todays designs, therefore is not taken into account • When Mach no. = 1 is reached in a turbomachine blade row, 𝑚ሶ 𝑟𝑒𝑑 (and 𝑚ሶ 𝑐𝑜𝑟𝑟 ) can not be further increased. The blade row is choked. Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

2019

6 - 16

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

mred,2 

m2 Tt 2 pt 2

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

2019

6 - 17

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

pt 4 pt5

rotor and stator choked

% Reduced Speed

Inlet Reduced Mass Flow

Turbine max. Flow

ht ,T Tt 4

Reduced Specific Work

T 

Turbine Map (2)

throat area choked

Total Turbine Expansion Pressure Ratio

Turbine Map (1)

throat area choked

n2  Tt 2 n red ,2

rotor chokes first

rotor choked

choke boundary

Stator choked

4 t4 mred,4  m t 4T p 2019

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

Capacity

6 - 18

Reduced Massflow x Reduced Speed 2019

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

m4 Tt4 n4  Tt 4 pt 4 6 - 19

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Compressor Operating Line (1)

Compressor Operating Line (2)

Flow capacity and flow variation of the turbine, which acts as a primary throttling device in the „system“, are dictating the level of pressure (pressure ratio) in the compressor. Rotational speed and power of compressor and turbine are directly coupled on the given common shaft. Hence the interaction between these two components is strong.

Compressor

Combustion Chamber

Intake

pt3

Throttle

pt2

. . m4/m 2= const A4/A 2 = const pt4/p t3 = const

3

2 . m2

Tt2

A 2 pt2 . m2

Tt2

A 2 pt2

=

=

. m4

Tt4

A4 pt4

4 . m2 . m4

Tt2

pt4 pt3 A4

Tt4

pt3 pt2 A2

. . m2 A 4 p t4 m4 Tt4 . m4 A 2 pt3 A4 pt4

Tt2 pt3 Tt4 pt2

. m2

Kühl‘s Equation

A 2 pt2

const. (A4 choked) Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

2019

6 - 20

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

2019

Compressor Operating Line (4)

The lowest operating line results from operation with no heat release in the combustor (or combustor removed):

>>

m 2 Tt 2 m 3 Tt 3 Tt 2 pt 3 A3 m       2 3 pt 2  A2 pt 3  A3 Tt 3 pt 2 A2 m nc 1

using the polytropic equation re-writing yields:

   3 A2  pt 3   1 m  m     2 A3  pt 2  2, red m 3,red m       const

6 - 21

Institute of Turbomachinery and Flight Propulsion Faculty of Aerospace and Geodesy Technical University of Munich

Compressor Operating Line (3)

m m2  m3  2 m3

Tt2

Tt3  p t3  nc  Tt2  pt 2 

and

Compressor

Intake

Throttle

2

3

Increase of

Effect on pt3 /pt2

(Position of operating line)

2n c nc1

 2, red)  f (m

The following table shows the effects of a change in performance parameters with the help of the derived operating line equation (no heat release in combustor) :

where

2nc 2 c  1,1887 nc 1  1

using κ = 1,4 and ηc = 0,90 gives a compressor operating line of low positive curvature

 2,corr m red

increase

2 m 3 m (bleeding)

decrease

ηC

A2/A 3

 3,corr m red

(exponent decreases)

increase

decrease

decrease

throttling the compressor … can be used to generate a full compressor map by individually setting a desired reduced speed and then go through a throttle setting variation (from fully open towards closed): pressure ratio and reduced mass

C

mred,3

flow will vary point by point going up the speed line ... until stall occurs and limits the speed line;

… when the throttle area is varied compressor exit reduced mass flow can be modulated, resulting in a lower or higher operating line 2019

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

6 - 22

2019

Flight Propulsion 1 and Gas Turbines Prof. Dr.-Ing. V. Gümmer

mred,2 6 - 23...