The effect of forward sweep in a transonic - MTU Aero Engines

THE EFFECT OF FORWARD SWEEP
IN A TRANSONIC COMPRESSOR ROTOR
Harald PASSRUCKER Martin ENGBER
MTU Aero Engines, Dachauer Straße 665, 80995 München / Germany
Harald.Passrucker@muc.mtu.de Martin.Engber@muc.mtu.de
Stephan KABLITZ Dietmar K. HENNECKE
Darmstadt University of Technology, Gas Turbines and Flight Propulsion,
Darmstadt / Germany
kablitz@gfa.tu-darmstadt.de hennecke@gfa.tu-darmstadt.de
ABSTRACT
This paper presents design and testing of a transonic compressor rotor with
forward sweep. The rotor was used to investigate the influence of forward sweep on
performance and stability of a single stage transonic compressor compared with a
baseline design with radially stacked blade sections. The comparison was done
numerically with the 3D Navier-Stokes code TRACE_S and experimentally in the
Darmstadt Transonic Compressor test rig. It was found that the new rotor with forward
sweep has an increased efficiency and also a much better stall margin (much more in the
rig test than predicted by the 3D Navier-Stokes calculation). Particularly close to stall
the forward sweep diverts the flow towards the blade tip region which helps to stabilise
this region. For that reason it is possible to throttle the forward swept rotor much
further as the radially stacked rotor although the forward-swept rotor does already
suffer from separated flow in the hub.
INTRODUCTION
A major part of the losses in transonic compressor rotors is created near the blade tip.
Shock losses and the interaction of the shock with other flow phenomena, like tip clearance
flow or boundary layers, also contribute to these losses. The tendency of the shock to cause
boundary layer separation can account for an amount of loss which is significantly higher than
the actual shock loss. Therefore, sweep has been considered as a method to reduce shock
strength and to improve efficiency and surge margin. Denton et al. [1] analysed by CFD that
the effect of sweep and lean on transonic fan efficiency and pressure ratio is remarkably
small, but have a significant influence on the stall point of the fan. Ulrich [2] numerically
investigated the influence of sweep and lean on a transonic rotor with the same result. There
are mainly 3 physical effects how sweep does influence the flow in a blade row.
Basic effects of swept blades on compressor flow
a) Influence of the blade loading (figure 1): The pressure gradient perpendicular to a
plane end wall must be zero, since there can be no acceleration perpendicular to the wall. In
the case of figure 1 (aft sweep), the blade loading near the lower wall must be reduced near
the leading edge where the loading rapidly falls to zero (no blade) as one moves
perpendicularly away from the wall. Conversely the loading on the lower wall will tend to be
increased near the trailing edge since there can be little pressure difference between it and the
more highly loaded region above it. The opposite effect occurs near the upper wall. Generally
the loading in the tip region is reduced in the front area with the forward-sweep which results

in a leading edge which is more tolerant to changes in incidence. Furthermore the tip leakage
is reduced in this area (lower loading).
b) Influence on the shock position (figure 2): In the spanwise direction, the shock cannot
intersect the outer casing obliquely. It must either turn normal to the casing or possibly
bifurcate in a shock/boundary-layer intersection. This requirement on the spanwise shock
shape near the casing is an inviscid phenomenon. In the absence of an endwall, the shock
shapes for the forward- and aft-swept rotors would be bent forward or backward in similar
fashion. In the presence of the endwall, however, the shock must turn normal to the casing,
moving upstream for an aft-swept rotor and downstream for the forward-swept rotor.
Generally, a shock position which is further downstream in the tip region, leads to a better
stall margin because the rotor can be throttled further until the bow shock detaches from the
leading edge.
c) Influence on the accumulation of low momentum fluid near the endwall (figure 3): In
a conventional rotor, fluid particles inside the blade boundary layer move radially outward
due to the imbalance between the centrifugal forces and the pressure gradient. The
accumulation of low momentum fluid near the endwall is considered to be a major cause of
increased aerodynamic loss and reduced operation range. In the case of a forward-swept rotor,
two mechanisms lessen the accumulation of low momentum fluid near the endwall. First, the
radially migrating boundary layer flow cannot reach the endwall region due to the forward
sweep of the blade. Second, the region of high pressure on the suction surface after the peak
in the pressure distribution is located further upstream at the tip region than at the hub region.
Therefore, radial migration of the low momentum fluid is suppressed and accumulation of
low momentum fluid near the tip is reduced.
Figure 1: Effect of sweep on blade loading
(Denton [3])
Figure 2: Endwall effect on shock structure
near the casing (Hah et al. [4])
Figure 3: Secondary flow in a forward swept rotor (Yamaguchi et al. [5])
History of different swept Rotors tested at TU-Darmstadt
The Darmstadt Transonic Compressor test rig was brought into operation in close cooperation
with MTU Aero Engines Munich in 1993. A series of 3 rotors was used to
investigate the influence of sweep and lean on performance and stability of a single stage
transonic compressor. The baseline Rotor No.1 (Rotor 1) with radially stacked blade sections

was designed by Schulze et al. [6]. To investigate the influence of blade sweep, especially in
the tip region, a Rotor No.2 (Rotor 2) was designed with considerable aft-sweep. It was
investigated in 2000 (Blaha et al. [7]). Rotor No.3 (Rotor 3) features forward sweep and was
tested in 2001.
TEST FACILITY AND MEASUREMENT EQUIPMENT
This section briefly describes the test rig (figure 4, 5) and conventional instrumentation
used to determine speedlines and efficiency.
Figure 4: Sketch of the test facility Figure 5: Cross section of the test compressor
Inlet total pressure and temperature are taken in the settling chamber in front of a
bellmouth. At the inlet, wall static pressure is measured to determine the mass flow by using a
calibrated nozzle. Pressure losses in the inlet duct are taken into account by an experimentally
determined loss coefficient. The downstream flow conditions are taken from fixed total
pressure and total temperature probe rakes located on the bearing support struts behind the
stator (figure 6), while the stator is traversed circumferentially. Shaft speed, power and torque
are measured by a Torquemeter measuring device between the 800kW DC-drive and the
compressor.
Figure 6: Total pressure and total temperature probe rakes
Measurements of speedlines were performed by recently upgraded Pitot type total
pressure and total temperature rakes mounted on the five struts downstream of the stator.
Complementing the eleven pitot type total pressure probes two static pressure taps are located
at the same axial position to gather static pressure information at hub and casing. By
traversing the stator upstream of the rakes in increments of 5% stator pitch, the stator exit

plane is resolved with 20 positions pitchwise and 13 probe locations from hub to tip, yielding
260 single pressure values. For determination of the total pressure ratio the data is at first
averaged circumferentially, using an arithmetic average (since measuring the whole
temperature distribution required for the massflow weighed average would take way too much
time). For averaging in the radial direction, the pressures are weighted according to local
massflow using the measured radial distribution of total temperature between two stator
wakes. Isentropic efficiency is calculated by comparing compressor work input to the flow
taken from pressure measurements with work input at the shaft which is measured with the
Torquemeter device of the test facility. Total temperature measurements at stator exit give
good general information about radial distributions of efficiency but quantitatively precise
averaged results are too difficult to obtain due to the rather long duration of temperature
measurements and become even less reliable at part speed conditions. The data aquisition
takes 4 minutes for each operating point. For a 95% confidence level (U95) at 100% speed
and peak efficiency operating point this yields:
• mass flow rate +/- 1.1%
• pressure ratio +/- 0.5 %
• isentropic efficiency +/- 1.4 %
GEOMETRIC DESIGN OF ROTOR 3
Figure 7 shows both rotors. Rotor 1 was designed conventionally with blade profiles
stacked radially along their centres of gravity. Rotor 3’s design features are lower blade
number, pronounced forward-sweep and higher blade chord length in the tip region to reach a
better stability and efficiency.
Rotor 1 (radially stacked) Rotor 3 (forward-swept)
Figure 7
The blade number of Rotor 3 was reduced compared to Rotor 1 from 16 blades to 14
blades. For this reason the solidity of Rotor 3 is generally lower than of Rotor 1 (figure 9).
The blade chord length of Rotor 3 was increased in the tip region (figure 9) which increases
the solidity there and improves stability and efficiency. The inlet and exit metal angles (figure
10) are more or less the same for both rotors to deliver the same work. Figure 11 shows the
stagger line of Rotor 3 profiles (centre of gravity) in axial direction. The light-grey curves
illustrate the displacement to indroduce the forward sweep. A small lean (grey curve) was
required in the tip region to balance the mechanical stress in this area. The black curve is the

actual stagging line, the sum of both displacements. The sweep displacement is counted in
direction of the blade chord, the lean displacement perpendicular to the chord. The sweep
stagging line with the backward displacement in the middle was designed in a manner to
lower the stress in the leading edge. This is important for FOD (foreign object damage) cases.
chord length L [m]
0.005
0.0 0.2 0.4 0.6 0.8 1.0
relativ height X/H [-]
ROTOR 1
ROTOR 3
Figure 8: Chord length
metal angle [°]
exit angle
L/T [-]
ROTOR 1
ROTOR 3
0.0 0.2 0.4 0.6 0.8 1.0
relativ height X/H [-]
Figure 9: Solidity
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0
radius [m]
0.200
0.190
0.180
0.170
0.160
0.150
0.140
0.130
0.120
0.110
0.100
inlet angle
relativ height X/H [-]
blade chord angle
Figure 10: Blade metal angles
sweep
lean
both
axial displacement
Figure 11: Profiles stagger line Rotor 3
5
ROTOR 1
ROTOR 3
0.1

COMPUTATIONAL GRID AND BOUNDARY CONDITION S
The simulation was performed with the steady 3D Navier-Stokes-Solver TRACE_S
(details are found in Fritsch et al. [8]) on a relative fine H/O-grid (figure 12) with a H-grid in
the rotor tip gap (0.9mm) (Rotor H(105×33×65), O(129×9×65), Hgap(129×5×9); Stator
H(105×33×65), O(129×9×65)) for both rotors. The k-ε model was uses for turbulence
modelling with wall function. Design speed is 20,000 rpm. At the inlet total temperature, total
pressure and flow angles are forced. The strong gradient of the total pressure at the casing
boundary layer was accounted for by an appropriate boundary condition for that region.
Information between rotor and stator domain is transferred with a mixing plane interface. At
the stator outlet the average pressure with radial equilibrium was set.
Figure 12: Computational grid (Rotor 3)
3D-ANALYSIS AND TEST RESULTS
There is good overall agreement between calculation and measurement (see figure 13,
15). Surge margin of Rotor 3 is fortunately much better in the measurement. The reason is
that the numerical stability is determined by the stator tip region of where separation occurs.
In the experiment this is no trigger for rotating stall. Therefore in reality there is separation in
the stator but no rotating stall at this operation point and the compressor can be throttled
further. This is also shown in the compressor maps (figure 13, 14). The speed line of the rotor
alone is still rising and delivers more pressure rise, whereas the stage’s speed line turns
horizontally with no further pressure rise.
With Rotor 3 it was possible to increase the peak efficiency (Navier-Stokes 1.5%,
measurement 1.5%) with the above described features (reduced blade number, increased blade
chord length in tip region, forward sweep). The blade number reduction reduces the blockage
and therefore the choke margin moves to higher massflow with a wider operating range.
Furthermore there is a much higher stability at the throttled condition. Unfortunately it is not
possible to allocate the efficiency profit and the higher stability to each design change, but
numerical experiments show that the efficiency profit can be accounted to equal parts to
increased blade chord length, forward sweep and slight profile adoptions. The higher stability
is attributed mainly to the forward sweep.
Figure 16 shows the measured compressor map for both rotors. Rotor 3 was measured
from 100% down to 30% speed, Rotor 1 only from 100% to 80% speed. Rotor 3 has nearly
constant peak efficiency from 80% to 100% speed, which covers the most operating points in
an engine application. Rotor 1 has its peak efficiency at 80% speed and is about 0.4% better
than Rotor 3. The difference to 100% speed in efficiency is 2.1%. The stability at all speed
lines of Rotor 1 is significantly lower than Rotor 3. The last stable point at 90 % speed of
Rotor 3 has been taken in the experiment, although there is a strong “negative” pressure rise.
Rotor 3 obviously suffers from strong separation in the hub region and produces a lot of
losses. But in the tip region the rotor is still stable and rotating stall can not yet be detected.

total pressure ratio [-]
total pressure ratio [-]
1.85
1.80
1.75
1.70
1.65
1.60
1.55
1.50
1.45
1.40
1.35
Figure 13: Stage compressor map
1.85
1.80
1.75
1.70
1.65
1.60
1.55
1.50
1.45
1.40
1.35
100%
100%
STAGE (Navier-Stokes)
ROTOR 1
ROTOR 3
1.30
14 14.5 15 15.5 16 16.5 17
STAGE (measurement)
ROTOR 1
ROTOR 3
massflow [kg/sec]
1.30
12.5 13 13.5 14 14.5 15 15.5 16 16.5 17
Δη=2.5
massflow [kg/sec]
Δη=2.5
Figure 15: Stage compressor map
isentrop efficiency[%]
isentrop efficiency[%]
total pressure ratio [-]
total pressure ratio [-]
1.85
1.80
1.75
1.70
1.65
1.60
1.55
1.50
1.45
1.40
1.35
100%
ROTOR (Navier-Stokes)
ROTOR 1
ROTOR 3
1.30
14 14.5 15 15.5 16 16.5 17
2.00
1.90
1.80
1.70
1.60
1.50
1.40
1.30
1.20
1.10
massflow [kg/sec]
Figure 14: Rotor compressor map
30%
40%
=5
STAGE (measurement)
50%
65%
80%
90%
100%
1.00
2 4 6 8 10 12 14 16 18
massflow [kg/sec]
Δη=2.5
ROTOR 1
ROTOR 3
Figure 16: Stage compressor map
isentrop efficiency[%]
isentrop efficiency[%]

Figure 17 illustrates radial distributions of stage total pressure ratio and isentropic
efficiency near choke (m=16.7), peak efficiency (m=16.4), near stall (m=12.6) and one point
in between (m=15.2). Remarkable is the near stall radial distribution, where the rotor still
delivers high pressure rise in the tip region which indicates high stability. In the mid there is a
little penetration in pressure rise. The other curves seem to be consistent with a continuous
pressure rise at different throttle conditions. The highest efficiency is as expected in the mid
region. Towards the tip there is a strong decrease of efficiency and pressure rise as a result of
the high mach number with the resulting shock losses. Radial distributions of stage total
pressure ratio and isentropic efficiency of Rotor 1 are displayed in figure 18 near choke
(m=16.4), peak efficiency (m=16.2) and near stall (m=15.0). Here the character of the radial
pressure rise distribution is not changing at throttled conditions.
total pressure ratio [-]
total pressure ratio [-]
100%
0.1
m=16.7, PIT=1.31
m=16.4, PIT=1.48
m=15.2, PIT=1.54
m=12.6, PIT=1.55
0.0 0.2 0.4 0.6 0.8 1.0
relative height X/H [-]
100%
0.1
m=16.4, PIT=1.37
m=16.2, PIT=1.45
m=15.0, PIT=1.53
0.0 0.2 0.4 0.6 0.8 1.0
relative height X/H [-]
isentropic efficiency [%]
10
Figure 17: Rotor 3 (measurement)
m=16.7, PIT=1.31
m=16.4, PIT=1.48
m=15.2, PIT=1.54
m=12.6, PIT=1.55
0.0 0.2 0.4 0.6 0.8 1.0
relative height X/H [-]
100%
0.0 0.2 0.4 0.6 0.8 1.0
relative height X/H [-]
Figure 18: Rotor 1 (measurement)
Figure 19 clearly shows the difference in leading edge and trailing edge contour of Rotor
3 with forward sweep and the baseline Rotor 1. The streamlines of Rotor 3 are generally
diverted towards the tip region as a result of the forward sweep compared to Rotor 1. This
feature is more prominent towards the stall condition. The forward sweep sucks particularly at
the near stall condition the flow in the tip region which stabilise this region.
isentropic efficiency [%]
10
m=16.4, PIT=1.37
m=16.2, PIT=1.45
m=15.0, PIT=1.53
100%

peak efficiency near stall
- - - - Rotor 1 —— Rotor 3
Figure 19: Meridian stream lines (Navier-Stokes)
Figure 20 illustrates the isentropic mach number distribution of both rotor’ blade sections
at peak efficiency operation conditions. The biggest difference can be detected in the tip
region, where the most losses are located. The loading near the leading edge of Rotor 3 is
reduced which causes less tip leakage. The mach number upstream of the shock is also lower
which results in lower shock losses. The shock position is much further downstream on the
profile which is improving stability. In the hub and mid section the loading of Rotor 3 is
increased as a result of the reduced blade number. In the mid section the shock system is split
in two which also helps to reduce losses at a same pressure rise.
0.2
hub mid tip
- - - - Rotor 1 —— Rotor 3
Figure 20: Isentropic mach number ? peak efficiency (Navier-Stokes)
0.2
0.2

CONCLUSIONS
The present study indicates that the numerical prediction of global values like massflow,
pressure rise and efficiency with TRACE_S is very close to the experiment and allows the
designer to optimise the blades with a 3D Navier-Stokes solver. The calculation as well as the
measurement show an increased efficiency and also a much increased stall margin (much
more in the rig test than predicted by the 3D Navier-Stokes calculation) for the forward swept
rotor compared to a radially stacked rotor. The increased efficiency results not only from the
forward sweep but also from the increased chord length in the tip region with simultaneous
reduction of the blade number. Particularly close to stall the forward sweep diverts the flow
into the tip region which improves stability in this region. Even if separation occurs in the hub
region, the forward swept rotor can be operated in a stable condition without developing
rotating stall.
ACKNOWLEDGEMENTS
The work presented here was supported by the German Ministry of Economics Affairs.
We are also grateful to the management of MTU Aero Engines for the permission to publish
the result.
REFERENCES
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auf Wirkungsgrad und Stabilitätsgrenze einer Hochdruckverdichterstufe,
Diplomarbeit, MTU Aero Engines GmbH – Fachhochschule Konstanz
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Paper 98-GT-561, Stockholm - Sweden
[5] Yamaguchi, N., Tominaga, T., Hattori, S., Mitsubishi, T. (1991), Secondary-Loss
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