Proceedings of TURBO EXPO 2005:
ASME Turbo Expo: Land, Sea & Air 2005
June 6-9, 2005, Reno-Tahoe,Nevada, USA
GT-2005-68913
DRAFT
DEVELOPMENT OF NEW HIGH AN2 LAST LP STAGE TURBINE & EXHAUST
SYSTEMS – A COST EFFECTIVE SOLUTION FOR THE 21ST CENTURY
Brian Haller
Siemens Industrial Turbomachinery , Ruston House,
PO Box 1,Waterside South, Lincoln LN5 7FD, UK
brian.haller@industrial-turbines.siemens.com
ABSTRACT
A new high AN2 last LP stage turbine has been developed to
provide leading performance for turbomachines in the 21st
Century. It required a multi-disciplinary design approach
involving aerodynamics, new materials (forged γ- Titanium
Aluminide alloy), mechanics (stress and vibration) and new
manufacturing technologies .
The objective of the design was to achieve around 5% gain in
last stage total-to-static efficiency ,relative to the current
competitive datum, for a very compact machine with few parts
count ie reduced cost. It will be shown that the new material
allows the optimum aerodynamic design for the stage to be
achieved which cannot be done with conventional nickel-based
alloys .
The paper will present details of the novel approaches used
for the design including preliminary optimization, blading
design and results from multistage 3D viscous predictions. The
new stage has been tested in a Warm Air Turbine test rig at full
scale engine representative conditions. The “loop was closed”
by comparing the detailed test measurements with 3D viscous
analyses. This gives high confidence in the new approaches
used.
Additionally, the development of new compact high
performance axial-radial and axial exhaust systems ,which
were designed to operate downstream of the new last LP
stage, are described.
The new LP technologies have already been scaled and cloned
to a conventional material design and sold by the Company.
Nomenclature
A exhaust annulus area
D diameter (h=hub, t=tip)
LE leaving energy
m mass flow rate
N shaft rotational speed
P01 inlet stagnation pressure
T01 inlet stagnation temperature
U mean blade speed
Va axial velocity
x exhaust loss coefficient
∆H stage total-to-total heat drop
Introduction
The purpose of this Technology project was to exploit the
material properties of new forged γ-TiAl alloy (PX5-300 ie
TNB V5) to produce a new high efficiency/cost effective LP
stage design.
Forged TiAl has around ½ the density of conventional nickelbased alloys with higher specific properties. Therefore much
longer LP stage designs can be used with higher efficiency. The
material also allows the optimum aerodynamic designs to be
used eg pitch/width ratio and number of blades etc. Its
elongation at room temperature is satisfactory (around 1%) and
it can be ground/5-axis milled by specialist techniques. This is
the same material which has revolutionized the performance of
F1 racing cars and is also finding applications in turbochargers,
other IC engine components etc. Initially cast TiAl material
was investigated but this had problems in the casting (cracks,
cavities ..) and had around half the specific strength of the
forged material. Also casting requires a thicker blade trailing
edge and other thickening compromises which increases
weight.
Specifically the objectives were :
- develop new high AN2/high load TiAl LP stage to provide
stage total-to-static efficiency gain of 5% at a total-to-total
pressure ratio of around 4:1 cf. current production design
- achieve better off-design performance characteristic and
explore extending to even higher load conditions.
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- achieve life and reliability first time (34,000 hour, 99%
availability, 2400 normal starts)
- halve the cost for LP rotor blades
- (also the new material can be used to reduce weight in aero
engines and to reduce disc stresses for the same blade height).
Note that a) the design is fully shrouded for improved vibration
control and to reduce overtip leakage losses and b) the design
inlet relative stagnation temperature to the rotor is around 650
deg C at mid-height. Note that the TiAl material is currently
limited to around 700 deg C at the tip.
This new LP stage was designed to operate downstream of the
new HP stage described in Ref. 1 in a two-stage overhung
turbine configuration. A pressure ratio of around 16:1 and
above can be achieved with only 2 turbine stages. This gives
large advantages in terms of :
- cost/parts count
- performance
- length and weight
- being able to achieve the optimum downstream axial exhaust
performance
- assembly
- on-site maintainability etc…...
New stage design
The new LP stage and exhaust systems detailed in this paper is
patent applied for and pending (Ref. 2 and 3) and has the
following overall design characteristics:
Dt/Dh=1.72
AN2=60*1012 mm2 rpm2 (with solid LP rotor )
Va/U=0.47
Loading ∆H/U2=1.55
Inlet swallowing capacity
m& T01
P01
of the LP stage and exhaust system This involves maximising
the LP stage total-to-total efficiency and minimising the
exhaust loss parameter (x.LE/ isentropic stage total-to-total
heat drop).
A Controlled Flow design philosophy (see Ref. 1) was used
for the LP NGV to :
- reduce secondary loss and give the highest total-to-total
efficiency for the stage
- match the flow at outlet from the HP stage , ensure a
consistent design philosophy throughout the turbine and
maintain a smooth streamline pattern.
Three alternative types of LP rotor design philosophy were
studied via a streamline curvature throughflow code :
- constant relative outlet angle, beta2
- beta2 increasing towards the tip (relative to the axial
direction ) to reduce the leaving energy from the LP stage.
Differing amounts of twist were studied.
- Controlled Flow vortex distribution (this is the first time it
has been tried for a last LP stage ).
The Controlled Flow rotor was selected as this gave :
- best flow structure and efficiency
- consistent design philosophy and streamline pattern
throughout the whole turbine
- ideal stagnation pressure distribution into the exhaust system
to minimise the losses (see later section).
The layout of the new stage installed in the Warm Air Turbine
test Rig is shown on Fig.1 . In this case for the first test there is
an Inlet Guide Vane which simulates the flow (swirl) delivered
from the upstream HP stage.
= 164.6 Kg/s K 1/2 bar -1
This is an increase in exhaust area of around 25% cf
conventional design ie 50% lower leaving energy. The amount
of disc/root blade cooling is very small (this affects AN2 which
can be achieved) .The design pressure ratio of the stage is
around 4:1 but it has the capability to operate efficiently to
higher load conditions.
It was originally designed for a 5 MW class single-shaft
industrial gas turbine but the new methods and technologies are
generic and it can be scaled/cloned to any application. The
methods and philosophies used to design the stage and exhaust
systems are detailed in the following sections.
Stage layout and throughflow design
Optimising the stage design involved aerodynamic and
mechanical disciplines with a number of iterations to establish
the LP rotor base radius and exhaust area. The next step was to
optimise the stage layout. For the last stage the important
performance parameter is the overall total-to-static efficiency
Fig.1 Layout of the new LP stage and axial-radial exhaust in
the Warm Air Turbine Test Facility
U3 optimisation
A U3 optimisation study (for method see Ref. 4 and 5 ) was
carried out for the LP rotor and LP NGV using the velocity
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Copyright © 2005 by ASME
triangles from the throughflow calculations. This a physicallybased method. Space precludes the publication of all the U3
results as around 1000 alternative sections were calculated for
the NGV and rotor. This gives the optimum pitch/width for the
blades and optimum profile designs at root , mean and tip ie
ideal velocity and turning distributions through the passages. It
is not possible to do this optimisation with conventional
correlations or Navier Stokes codes.Fig.2 shows the results for
the rotor where the stage efficiency debit due to profile and
secondary losses for the rotor is plotted against pitch/axial
width at the root.
0.65
0.6
(pitch/axial width)root
0.55
Variation of stage efficiency debit with (s/w)root
Stage Efficiency debit(TiAl rotor)
0.0245
0.5
0.45
TiAl
GT 1(unshrouded)
GT 2
GT3
ST 1
ST 2
ST 3
0.4
0.024
0.0235
0.023
0.35
0.0225
0.022
TiAl rotor
0.3
0.0215
1.3
0.021
1.4
1.5
1.6
1.7
1.8
Diameter ratio=Dt/Dh
0.0205
0.02
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
(s/w)root
Fig.2 Summary of results from U3 optimisation study on LP
rotor
Conventional shrouded nickel-based designs have to use much
lower values of pitch/width at the root which increases the
profile losses and lowers the stage total-to-total efficiency (see
Fig.3 ). There were only 36 LP NGV’s and 53 LP rotor blades.
This lowers the transonic trailing edge blockage losses.
Previously 52 LP NGV’s were used and a nickel-based LP
rotor design typically has around 80 blades . Thus the changes
are very substantial.
Fig.3 Comparison of root (pitch/axial width) for shrouded LP
blade designs (GT= gas turbine , ST=steam turbine). All the
designs are shrouded apart from GT1.
3D design and analysis using multistage 3D RANS code
Having decided on the skeletal parameters for the stage , the
blading was designed using the in-house “XDESIGN” system.
This is a full interactive Q3D transonic design system with
cloning, inverse, stacking options etc. It includes a finite
element potential flow solver , MISES (with base drag losses)
and integral/differential boundary layer methods. It is very fast
and well validated over 10 years. The blade profiles are defined
by Nutbourne parametric cubic splines with continuous smooth
curvature Fig. 4 and 5 show the 3D views of LP NGV and LP
rotor. Both blades are Controlled Flow. The LP rotor was
stacked on the centroid . The LP NGV has :
Tangential lean of +10 degrees (pressure surface pointing
radially inwards)
Taper
Tilt
All the above were to optimize the flow structure and maximize
the efficiency.
The performance of the stage was then analysed using a
number of 3D multistage viscous methods - Stage3d (adapted
from Dawes), MULTALL (Denton), TF3D (Li He) by 3
independent specialists. All results showed very good flow
structure and performance.
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Fig.5 Isometric view of high AN2 LP rotor
The grid for the Stage3d calculation engine configuration is
shown on Fig.6 .This case used a standard H-grid but a
multiblock curvilinear grid is also available. The BaldwinLomax turbulence model was used. The relative Mach numbers
contours for the 2 stages are shown on Fig.7 . It is clear that all
the blades are transonic /high load/high lift designs. As an
example of the level of checking, the relative velocity vectors
onto the LP rotor at mid-height are shown on Fig.8 It can be
seen that the stagnation point is correctly aligned with no
incidence onto the blade. All the sections were checked in this
manner.
Fig.4 Isometric view of LP NGV
Fig.6 Meridional view of grid for HP and LP stages for
multistage 3D RANS calculation
Fig.7 Relative Mach number contours at mid-height
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New compact high performance axial-radial and axial
exhaust systems
In the axial-radial exhaust the flow is diffused and turned
through 90 degrees. The flow on the lip governs the exhaust
performance since this is where the velocity is lowest by simple
streamline curvature considerations. The dynamic head is lost
in the large exhaust plenum and so the stagnation pressure at
outlet pipe discharge is simply equal to the static pressure at the
separation point on the lip . Therefore for the new axial-radial
exhaust diffuser, the peak velocity on the lip was minimized
in order to provide higher recovery . A conventional circular
arc lip design will accelerate the inlet flow due to its curvature ,
give a higher peak velocity on the lip and therefore always
give separation at a higher velocity (see Fig. 10).
Fig.8 Relative velocity vectors onto LP rotor leading edge at
mid-height
The root reaction for the stage was optimum at around 20%
(healthy and positive otherwise large root separation losses can
occur). In terms of reaction, the shocks losses were balanced in
the NGV and rotor to obtain the highest stage efficiency . The
radial distribution of stagnation pressure delivered by the LP
stage into the exhaust is shown on Fig. 9. It is ideal for high
exhaust performance being very flat with a slight increase at
the tip.
Fig.10 Static pressure distributions on hub and casing of
original circular arc axial-radial diffuser and new design (lip
are lower 2 curves and hub are higher 2 curves)
Fig.9 Radial distribution of stagnation pressure delivered into
exhaust
In the new axial exhaust a higher performance can be achieved
since the flow is not turned 90 degree. Again , improvements
can be achieved via an understanding of the boundary layer
physics eg see Ref..6. Initially when the starting turbulent
boundary layer is thin and has small momentum thickness it
can be diffused more rapidly without danger of separation .
Gradually as the surface boundary layer thickens the diffusion
is reduced . This enables diffusion down to a lower exit
velocity (ie higher pressure recovery ) and the ability to achieve
the recovery in a smaller axial distance. Initially the turbulent
boundary layer shape factor is around 1.4, is diffused up to
around 2.0 and held constant at this value along the surface
length of the diffuser on both hub and tip . This provides
margin against any danger of separation.. The diffuser was
designed using an
axisymmetric streamline curvature
throughflow method coupled to an axisymmetric integral
boundary layer method .It was then finally checked in a
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Copyright © 2005 by ASME
multistage viscous calculation with the upstream turbine
stages. This also enabled the strut angles to be optimized.
The principles are similar to those used for the design of
“Controlled Diffusion” compressor blades .
On Fig. 11 the distinctive shape of the new axial exhaust is
shown together with the streamline pattern. The hub is a
simple cone and the outer wall shape is designed to be concave
to give the “Optimised Diffusion” velocity distribution. The
inner cone has to flare outwards to avoid the swirl from the
turbine being magnified. It can be seen that the area increases
more rapidly initially. Fig. 12
shows the the velocity
distributions on the hub and tip of the “Optimised Diffusion”
diffuser. It is also an ideal distribution along the surface of the
hub of the diffuser because it is mainly controlled by the 1D
area distribution. The static pressure distributions on the hub
and tip are shown on Fig.13 and the recovery can be seen. The
outer wall shape is easily made by spinning. The new axial
exhaust gave very high predicted recovery and is currently
being tested on 2 engines.
Fig.13 Static pressure distributions on hub and tip of new
“Optimised Diffusion” axial exhaust
Fig.11 Streamlines for new “Optimised Diffusion” axial
exhaust system
160
ABS.CON. velocity (m/s)
140
120
100
Tip
80
Hub
60
40
20
Test results
The new LP stage and axial-radial exhaust has been tested in a
full scale Warm Air Turbine Test Rig at full scale and correct
Reynolds numbers /Mach numbers (see Fig.1).In the rig the
rotor blade was made from Titanium. The power produced by
the turbine is absorbed by a hydraulic dynamometer and full
area traversing carried out at LP stage exit and IGV exit using a
5-hole probe. The rig is extensively instrumented. Tests were
made over a range of speed and pressure ratio with two tip
clearance levels . Typical conditions are around :
stagnation pressure at inlet = 5 bar
stagnation temperature at inlet = 650K
mass flow rate = 35 Kg/s
rotational speed = 12000 rpm
shaft power = 4.4 MW
The picture of the balanced LP rotor assembly is shown on
Fig.14 The quality of the test hardware is evident. The rig
measurements confirmed the original design predictions and
that the targets had been surpassed. Presently “the loop is
being closed” by comparing all the detailed rig measurements
(swallowing capacity, efficiency, radial distributions of flow
parameters, off-design performance, exhaust recovery) with
3D analysis predictions.
0
0
200
400
600
800
1000
1200
x (m m )
Fig.12 “Optimised diffusion“ velocity distributions on hub
and tip of new axial exhaust
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Copyright © 2005 by ASME
The detailed material test results
specialized materials paper.
will be published in a
New manufacturing technologies and processes
Fig.14 Picture of LP rotor assembly for Warm Air Turbine Rig
Test
Material test results
Extensive material tests have been carried out over a number of
years including :
Tensile
Creep
Low cycle fatigue
High cycle fatigue
Pre-exposure tests at 400, 550, 700 deg C
Coating tests
Advanced heat treatment
Wear and fretting tests for shroud for shroud interlock face
Residual stress measurements caused by the manufacturing
processes
1400
1200
Strength (MPa)
1000
800
600
400
200
0.2% PS (MPa)
TS (MPa)
0.2% PS (MPa) Cast 4722XD (Mean)
TS (MPa) Cast 4722XD (Mean)
Poly. (TS (MPa))
Poly. (0.2% PS (MPa))
Poly. (0.2% PS (MPa) Min)
0
0
100
200
Poly. (TS (MPa) Min)
300
400
500
600
700
800
Temperature (°C)
Fig.15 Tensile strength of PX5-300, extruded material 14:1
As an example, Fig.15 shows the tensile strength of the forged
material compared to 4722XD cast TiAl material plotted
against temperature up to 800 deg C. The much higher strength
of the forged material is obvious.
Fig.16 Picture of forged/ground/machined TiAl last LP rotor
blade (this is conventional blade /”material substitution” design
for spin-rig and engine mechanical strain-gauge tests)
Fig.16 shows a picture of a finished blade. The manufacturing
processes are :
- Raw billet production at foundry
- Forging and heat treatment
- Grinding of blade roots
- 5-axis NC machining of aerofoil (possibly ECM could also be
used)
- Grinding of shroud
On Fig. 16 some waviness on the profile caused by the milling
is evident. This was within the manufacturing tolerances for the
blade but the waviness could be halved at a small additional
cost. The trailing edge thickess is 0.7 mm and no additional
thickening relative to the conventional cast nickel-based design
was required. No problems were experienced with the blade
manufacture . However forged TiAl is a different material and
therefore requires different “sealed processes” for all aspects in
its life (manufacture, handling, assembly, testing etc).
Conclusions
A new design philosophy for low pressure turbines, utilising
“Controlled Flow”/“U3” optimisation /high AN2 has been
demonstrated. New compact high performance axial-radial and
“Optimised Diffusion” axial exhaust systems have been
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Copyright © 2005 by ASME
presented. The aerodynamics of the new LP stage and axialradial exhaust was validated via a Warm Air Turbine Rig Test.
The new axial-radial and axial exhaust systems are also
presently being validated on engines. Very significant step
changes have been obtained and the original aggressive design
targets exceeded. The “loop is being fully closed” by
comparing all the detailed rig and engine measurements with
the detailed test information..
Acknowledgements
The author wishes to thank UK DTI New and Renewables
Energy Programme for funding this work. Gordon McColvin ,
Materials Technology Manager provided Fig.15.
References
1. Haller B.R. Anderson J. “Development of new high
load/high lift transonic shrouded HP gas turbine stage
design – a new approach for turbomachinery “, ASME
paper GT-2002-30363.
2. Haller B.R. “Gas turbine low pressure stage”, UK Patent
Application GB 2384276 A, 23/7/2003.
3. Haller B.R. “A diffuser for diffusing the exhaust gas
produced by an engine”, UK Patent Application
2004P01249GB, July 2004.
4. Denton J.D. “Loss mechanisms in turbomachines”, ASME
93-GT-435.
5. Haller B.R. VKI Lecture Series on “Secondary and Tip
Clearance Flows in Axial Turbines”, Full 3D turbine blade
design, 10-13 February 1997.
6. Stratford B.S. “An experimental flow with zero skin
friction throughout its region of pressure rise”, J. Fluid
Mech, Vol. 5, 17-35, 1959.
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