INCAS BULLETIN, Volume 10, Issue 3 2018, pp . 15 26 (P) ISSN 2066 -8201, (E) ISSN 2247 -4528 [623870]

INCAS BULLETIN, Volume 10, Issue 3/ 2018, pp . 15 – 26 (P) ISSN 2066 -8201, (E) ISSN 2247 -4528
Numerical Simulation s of Flow in Axial Compresso r
System , Preparatory Steps for Active Control
Irina -Carmen ANDREI1, Gabriela STROE *,2
*Corresponding author
1INCAS ‒ National Institute for Aerospace Research “Elie Carafoli” ,
B-dul Iuliu Maniu 220, Bucharest 061126, Romania ,
[anonimizat], [anonimizat]
2“POLITEHNICA ” University of Bucharest, Faculty of Aerospace Engineering,
Gh. Polizu Street 1 -7, Sector 1, Bucharest , 011061 , Romania ,
[anonimizat] *
DOI: 10.13111/2066 -8201. 2018.10.3. 2
Received: 29 May 2018/ Accepted: 21 June 2018/ Published: September 2018
Copyright © 2018. Published by INCAS. This is an “open access” article under the CC BY -NC-ND
license (http://creativecommons.org/licenses/by -nc-nd/4.0/)
6th International Workshop on Numerical Modelling in Aerospace Sciences, NMAS 2018 ,
16 – 17 May 2018, Bucharest, Romania, (held at INCAS, B -dul Iuliu Maniu 220, sector 6)
Section 1 – Launchers propulsion technologies and simulations of rocket engines
Abstract: For most compressor flows, the existence of turbulent shear stress es is essential to
compensate and overcome the adverse pressure gradients without separation. The performance of
compressor improves as the turbulent stresses get stronger relative to the laminar viscous stresses ,
which become significant as the Reynolds number increases. The methods used to desig n subsonic
axial flow compressors and fans are based on correlations such as deviation angle vs. incidence angle
to obtain the expected performance for a particular blade profile. In this paper, c ompressor models
that are used frequently in controller desi gns are described and the effect of the assumptions on the
prediction capability of these models is explained. The control concepts used to maintain and
therefore grant the stable operation are presented and monitored, as well as an outline of sensors,
actuators and automatic controllers is given . The survey concludes that the behavior of a compressor
system subsequent to instability onset is reasonably understood but the considered models can not
describe all types of instabilities encountered. There comes the recommendation that for a better
understanding of the mechanisms behind the onset of such instabilities , refinements as new
improvements of the existing models are require, which may give insights into new automatic control
methods.
Key Words: Automat ic control, axial compressor sy stem, controller, instabilities
1. INTRODUCTION
The axial compressors have a large variety of applications in aircraft industry; they represent
a basic component of the jet engines (turbojet, turbofan, turboprop and turboshaft engines ),
which are supposed to perform for aerospace propulsion, industrial gas turbines that generate
high power output , and processors targeted to pressurize gas or fluids , with applications to
chemical phenomena . [9] The axial compressors can maneuver a higher mass flow rate for
the same frontal cross section . This particular reason, in conjunction with the fact that the

Irina -Carmen ANDREI, Gabriela STROE 16

INCAS BULLETIN, Volume 10, Issue 3/ 2018 thrust of the engine is highly increased due to the signifi cant enlargement of the compressor
pressure ratio and air mass flow rate, are the important justifications for the use large scale
use of the axial compressors in aircraft jet engines [9].
For most type of the jet engines, the compression system consists in fan, low -pressure
compressor LPC and high -pressure compressor HPC. From the flow direction’s standpoint,
the fan is of axial flow -type, while the LPC and HPC can be of axial flow -type, centrifugal
flow-type or both [9]. An axial compressor system consis ts in sequence of blade row
cascades; usually, a compressor stage is made up from a r otating cascade and a fixed one [9].
The working fluid is passed through the fixed and rotating blades; the energy transfer is
performed within the rotating blade row; the fluid flow is guided inside the fixed blade rows,
and in certain cases, according to the cascade design, the fluid flow can be directed through
the rotating blade rows [9]. The high -performance compression stages are transonic (most of
them high subsonic to transonic) , where regions of subsonic and supersonic flow both exist
in the blade passages. The steady state performance of an axial compressor system is
described by the universal map of the axial compre ssor, which plots the averaged mass flow
rate versus the total pressure ratio ; such diagram is also referred as the characteristic or
performance map of the compressor.
2. COMPUTATIONAL STUDIES OF AXIAL COMPRESSOR SYSTEM
It is strongly recommended to avoid surge and rotating stall , because such phenomena can
lead to mechanical and thermal loads and can cause significant structural damage of the
aircraft engine parts . These aerodynamic instabilities are oscillating phenomena and can
reduce the possibility to increase the pressure and may induce a severe decay of the
compression system efficiency . There are three different methods to compensate for these
problems, namely: surge or stall avoidance, surge detection and avoidance, and increasing
the stall margi n approach [3-7].
For the first method , the automatic control systems , does not allow the compressor to
operate on the left side of the surge line. In order to locate the surge line on the compressor
map, a safety margin is specified. This safety margin may be defined as a function of the
pressure ratio, corrected mass flow or a numerical combination of pressure ratio and
corrected mass flow. The safety margin noted SM, is defined as a function of total pressure
ratio as [5-7]
𝑆𝑀=(𝑃02
𝑃01)𝑆𝑢𝑟𝑔𝑒−(𝑃02
𝑃01)𝑆𝑢𝑟𝑔𝑒 𝐴𝑣𝑜𝑖𝑑𝑎𝑛𝑐𝑒
(𝑃02
𝑃01)𝑆𝑢𝑟𝑔𝑒 𝐴𝑣𝑜𝑖𝑑𝑎𝑛𝑐𝑒 (1)
The total pressure at the compressor outlet and inlet is noted by P01 and P02,
respectively. In both case of control, active and passive , the characteristic performance maps
of the axial compressor are modified and the surge line is shifted to a lower mass flow. An
advantage of this methodology consists in the fact that the axial compressor can operate near
the peak efficiency and high-pressure ratios at lower mass flow rates. In the case of the
passive surge or stall, the geometry of the axial compressor is changed such that to modify
the stall margin [9-10]. The use of the variable inlet guide vanes represents a good option for
increasing the stall margin ; it has been successfully used to control the stall in axial
compressors. With this method ology , the inciden ce angle in axial compressors at lower mass
flow rates is reduced and the leading -edge LE separation is avoided . With in the inlet g uide

17 Numerical Simulations of Flow in Axial Compressor System, Preparatory Steps for Active Control

INCAS BULLETIN, Volume 10, Issue 3/ 2018 vanes, the direction of the flow at the leading edge is modified such that the angle of attack
decreases. The v ariable inlet guide vanes are very often used wh ile starting and accelerating
the aircraft engines , in order to avoid the crossing of the surge line. In active stall or surge
control, the axial compressor is equipped with devices , such as a bleed valve that can be
switched in the positions on or off. The active surge or stall control may be classified into
two classes: open -loop and closed -loop. In the case of closed -loop control, a feedback law is
applied to activate the controller, while in the open -loop control there are no feedback
signals [11-13].
Other methods used to avoid active surge or stall control , are represented by a ir
injection, air bleeding, and recirculation (a mix of injection and bleeding) [11-13].
Air injection is a method in which small amount of high pressure and high velocity air is
injected into the axial compressor upstream of the compressor inlet. This way, the flow and
the axial velocity component are increased; also, the local angles of attack are reduced, and
therefore, the leading -edge LE separation is compensated. The injected air may be obtained
from the diffuser downstream of the axial compressor or from a n auxili ary device [14-17].
Bleeding technique has been used to enable the efficient operation of the compressor
system, for a wide range of operating conditions.
The bleed valve is usually located either in the plenum exit or downstream of rotor on
the shroud. The bleeding system operation consists in introducing or expelling of a lower
fluid mass flow rate, taken from the working fluid , which does not have enough momentum
to overcome the vi scous and adverse pressure gradient forces in the plenum. By removing a
part of the highly pressurized flow downstream of the compressor, flow acceleration can
increase and surge -free operation is obtain ed [14-17].
Closed -loop active control represents an important integral part of the aircraft engines,
called the smart or intelligent aircraft engines. The closed -loop active control systems use a
sensor for detecting the growth of instabilities when a n axial compressor reaches the stall
conditions. By apply ing t his method, a control unit processes measured flow field data, the
temperature, pressure or axial velocity, from a stall -detection system .
A feedback control law connecting the sensed fluctuations to the rate of bleed is used to
stabilize the axial compressor. The control unit activates a set of actuator systems . Various
types of actuators are used for stabilizing the axial compressor system, for example: bleed
valve actuators, variable inlet guide vanes, recirculation, loudspeakers, movable plenum
walls, and air injections [14 -17].
One of the most important methods for investigating the complex flow phenomena in
aircraft engines system is the Computational Fluid D ynamics – CFD .
For many applications, the flow regime can be considered incompressible, and the
Laplace and Poisson equations can be applied to model the flow within the inl et and exit
ducts, respectively [14-17]. The pressure field calculations are independent of viscous effects
and can be calculated with mathematical classical models [18-20].
It is convenient to work with the full Navier -Stokes equations for mathematical
modeling and to perform numerical simulations, purposed to study the off-design conditions
by Computational Fluid Dynamics methods.
The Navier -Stokes equations describe the physical 3D model for the unsteady
compressible viscous flow, with the assumptions of the perfect Newtonian fluids and the
Stokes linear st ress-strain rate low considered [4].
In 3D Cartesian coordinates, the conservative equations in vector form are written, and then
for 3D unsteady compressible Reynolds -Averaged Navier -Stockes RANS equations are
presented in integral form and calculated w ith a finite volume scheme [4] .

Irina -Carmen ANDREI, Gabriela STROE 18

INCAS BULLETIN, Volume 10, Issue 3/ 2018 𝜕𝑞
𝜕𝑡+𝜕𝐸
𝜕𝑥+𝜕𝐹
𝜕𝑦+𝜕𝐺
𝜕𝑧=𝜕𝑅
𝜕𝑥+𝜕𝑆
𝜕𝑦+𝜕𝑇
𝜕𝑧 (2)
𝑞=
{ 𝜌
𝜌𝑢
𝜌𝑣
𝜌𝑤
𝐸𝑡}
,𝐸=
{ 𝜌𝑢
𝜌𝑢2+𝑝
𝜌𝑢𝑣
𝜌𝑢𝑤
(𝐸𝑡+𝑝)𝑢}
,𝐹=
{ 𝜌𝑣
𝜌𝑢𝑣
𝜌𝑣2+𝑝
𝜌𝑣𝑤
(𝐸𝑡+𝑝)𝑣}
,𝐺=
{ 𝜌𝑤
𝜌𝑢𝑤
𝜌𝑣𝑤
𝜌𝑤2+𝑝
(𝐸𝑡+𝑝)𝑤}
(3)
𝐸𝑡=𝜌[𝐶𝑣𝑇+1
2(𝑢2+𝑣2+𝑤2)] (4)
𝑝=𝜌𝑅𝑇 (5)
𝑝=(𝛾−1)[𝐸𝑡−1
2𝜌(𝑢2+𝑣2+𝑤2)] (6)
𝑅=
{ 0
𝜏𝑥𝑥𝜏𝑥𝑦
𝜏𝑥𝑧
𝑢𝜏𝑥𝑥+𝑣𝜏𝑥𝑦+𝑤𝜏𝑥𝑧+𝑞𝑥}
,𝑆=
{ 0
𝜏𝑦𝑥
𝜏𝑦𝑦
𝜏𝑦𝑧
𝑢𝜏𝑦𝑥+𝑣𝜏𝑦𝑦+𝑤𝜏𝑦𝑧+𝑞𝑦}
, (7)
𝑇=
{ 0
𝜏𝑧𝑥𝜏𝑧𝑦
𝜏𝑧𝑧
𝑢𝜏𝑧𝑥+𝑣𝜏𝑧𝑦+𝑤𝜏𝑧𝑧+𝑞𝑧}
(8)
𝜏𝑥𝑥=𝜆(𝑢𝑥+𝑣𝑦+𝑤𝑧)+2𝜇𝑢𝑥 (9)
𝜏𝑥𝑦=𝜏𝑦𝑥=𝜇(𝑢𝑦+𝑣𝑥) (10)
𝜏𝑥𝑧=𝜏𝑧𝑥=𝜇(𝑢𝑧+𝑤𝑥) (11)
𝜏𝑦𝑦=𝜆(𝑢𝑥+𝑣𝑦+𝑤𝑧)+2𝜇𝑢𝑦 (12)
𝜏𝑦𝑧=𝜏𝑧𝑦=𝜇(𝑣𝑧+𝑤𝑦) (13)
𝜏𝑧𝑧=𝜆(𝑢𝑥+𝑣𝑦+𝑤𝑧)+2𝜇𝑢𝑧 (14)
𝑞𝑥=−𝑘𝜕𝑇
𝜕𝑥 (15)
𝑞𝑦=−𝑘𝜕𝑇
𝜕𝑦 (16)
𝑞𝑧=−𝑘𝜕𝑇
𝜕𝑧 (17)
𝑝𝑟𝑒𝑓=𝜌𝑟𝑒𝑓𝑉𝑟𝑒𝑓2
𝛾 (18)
∰𝜕𝑞
𝜕𝑡𝑑𝑉+∬(𝐸𝑖⃗+𝐹𝑗⃗+𝐺𝑘⃗⃗)𝑛⃗⃗𝑑𝑆
𝑆−∬𝑞𝑉𝐺⃗⃗⃗⃗⃗
𝑆𝑛⃗⃗𝑑𝑆=∬(𝑅𝑖⃗+𝑆𝑗⃗+𝑇𝑘⃗⃗)𝑛⃗⃗
𝑆𝑑𝑆
𝑉 (19)

19 Numerical Simulations of Flow in Axial Compressor System, Preparatory Steps for Active Control

INCAS BULLETIN, Volume 10, Issue 3/ 2018 ∬(𝐸𝑖⃗+𝐹𝑗⃗+𝐺𝑘⃗⃗)𝑛⃗⃗𝑑𝑆
𝑆−∬𝑞𝑉𝐺⃗⃗⃗⃗⃗
𝑆𝑛⃗⃗𝑑𝑆=
=∑(𝐸𝑛𝑥+𝐹𝑛𝑦+𝐺𝑛𝑧)Δ𝑆−[𝑞𝑉𝐺⃗⃗⃗⃗⃗𝑛⃗⃗Δ𝑆]=
𝐴𝑙𝑙 𝐹𝑎𝑐𝑒𝑠
=[𝐸̂
𝑖+1
2,𝑗,𝑘+𝐸̂
𝑖−1
2,𝑗,𝑘]+[𝐹̂
𝑖,𝑗+1
2,𝑘+𝐹̂
𝑖,𝑗−1
2,𝑘]+[𝐺̂
𝑖,𝑗,𝑘+1
2+𝐺̂
𝑖,𝑗,𝑘−1
2] (20)
𝐸̂| 𝑖±1
2,𝑗,𝑘=[𝐸𝑛𝑥+𝐹𝑛𝑦+𝐺𝑛𝑧−𝑞(𝑉𝐺⃗⃗⃗⃗⃗𝑛⃗⃗)]∆𝑆] | 𝑖±1
2,𝑗,𝑘 (21)
𝐹̂| 𝑖,𝑗±1
2,𝑘=[𝐸𝑛𝑥+𝐹𝑛𝑦+𝐺𝑛𝑧−𝑞(𝑉𝐺⃗⃗⃗⃗⃗𝑛⃗⃗)]∆𝑆] | 𝑖,𝑗±1
2,𝑘 (22)
𝐺̂| 𝑖,𝑗,𝑘±1
2=[𝐸𝑛𝑥+𝐹𝑛𝑦+𝐺𝑛𝑧−𝑞(𝑉𝐺⃗⃗⃗⃗⃗𝑛⃗⃗)]∆𝑆] | 𝑖,𝑗,𝑘±1
2 (23)
∬(𝑅𝑖⃗+𝑆𝑗⃗+𝑇𝑘⃗⃗)𝑛⃗⃗𝑑𝑆
𝑆=∑(𝑅𝑛𝑥+𝑆𝑛𝑦+𝑇𝑛𝑧)∆𝑆
𝐴𝑙𝑙 𝐹𝑎𝑐𝑒𝑠 (24)
𝑓𝑁𝑢𝑚𝑒𝑟𝑖𝑐𝑎𝑙 =
[ 1
2[𝑓(𝑞𝐿)+𝑓(𝑞𝑅)]⏞ 𝑃ℎ𝑦𝑠𝑖𝑐𝑎𝑙 𝐹𝑙𝑢𝑥 𝑇𝑒𝑟𝑚
−1
2[|𝐴̃(𝑞𝐿,𝑞𝑅)|(𝑞𝐿−𝑞𝑅)]⏞ 𝐴𝑟𝑡𝑖𝑓𝑖𝑐𝑖𝑎𝑙 𝑉𝑖𝑠𝑐𝑜𝑠𝑖𝑡𝑦 𝑇𝑒𝑎𝑚
]
∆𝑆 (25)
𝑓(𝑞𝐿)=
[ 𝜌𝐿𝑈𝐿
𝜌𝐿𝑈𝐿𝑢𝐿+𝑝𝐿𝑛𝑥
𝜌𝐿𝑈𝐿𝑣𝐿+𝑝𝐿𝑛𝑦
𝜌𝐿𝑈𝐿𝑤𝐿+𝑝𝐿𝑛𝑧
𝑈𝐿𝐻0𝐿−𝑝𝐿𝑛𝑡]
,𝑓(𝑞𝑅)=
[ 𝜌𝑅𝑈𝑅
𝜌𝑅𝑈𝑅𝑢𝑅+𝑝𝑅𝑛𝑥
𝜌𝑅𝑈𝑅𝑣𝑅+𝑝𝑅𝑛𝑦
𝜌𝑅𝑈𝑅𝑤𝑅+𝑝𝑅𝑛𝑧
𝑈𝑅𝐻0𝑅−𝑝𝑅𝑛𝑡]
(26)
𝑞𝐿=
[ 𝜌𝐿𝑢𝐿
𝑣𝐿
𝑤𝐿𝑝𝐿]
,𝑞𝑅=
[ 𝜌𝑅𝑢𝑅
𝑣𝑅
𝑤𝑅𝑝𝑅]
(27)
𝑈=(𝑉⃗⃗−𝑉𝐺⃗⃗⃗⃗⃗)𝑛⃗⃗ (28)
𝐻0=𝐸𝑡+𝑝 (29)
𝑛𝑡=−𝑉𝐺⃗⃗⃗⃗⃗𝑛⃗⃗ (30)
𝑞𝐿=𝑞𝑖+1
6Φ
𝑖−1
2+(𝑞𝑖−𝑞𝑖−1)+1
3Φ
𝑖+1
2−(𝑞𝑖+1−𝑞𝑖) (31)
𝑞𝑅=𝑞𝑖+1−1
3Φ
𝑖+1
2+(𝑞𝑖+1−𝑞𝑖)−1
6Φ
𝑖+3
2−(𝑞𝑖+2−𝑞𝑖+1) (32)
Φ
𝑖−1
2+=Φ(𝑟
𝑖−1
2+),Φ
𝑖+1
2−=Φ(𝑟
𝑖+1
2−) (33)

Irina -Carmen ANDREI, Gabriela STROE 20

INCAS BULLETIN, Volume 10, Issue 3/ 2018 𝑟
𝑖−1
2+=𝑞𝑖+1−𝑞𝑖
𝑞𝑖−𝑞𝑖−1,𝑟
𝑖+1
2−=𝑞𝑖−𝑞𝑖−1
𝑞𝑖+1−𝑞𝑖 (34)
Φ(𝑟)=𝑚𝑎𝑥[0,𝑚𝑖𝑛(2𝑟,1),𝑚𝑖𝑛(𝑟,2)] (35)
|𝐴̃(𝑞𝐿,𝑞𝑅)|(𝑞𝑅−𝑞𝐿)=|𝜆1̃|Δ𝑞+𝛿1𝑞∗̃+𝛿2𝑁𝑛 (36)
𝛿1=(−|𝜆1̃|+|𝜆2̃|+|𝜆3̃|
2)Δ𝑝
𝑎2̃+|𝜆2̃|−|𝜆3̃|
2𝑞̃Δ𝑈
𝑎̃ (37)
𝛿2=(−|𝜆1̃|+|𝜆2̃|+|𝜆3̃|
2)𝑞̃Δ𝑈+|𝜆2̃|−|𝜆3̃|
2Δ𝑝
𝑎̃ (38)
𝑞∗̃=
[ 1
𝑢̃
𝑣̃
𝑤̃
𝐻0̃
𝜌̃]
,𝑁𝑛=
[ 0
𝑛𝑥𝑛𝑦
𝑛𝑧
𝑈̃]
Δ𝑆 (39)
𝜆1̃=𝑈̃ (40)
𝜆2̃=𝑈̃+𝑎̃ (41)
𝜆3̃=𝑈̃−𝑎̃ (42)
𝑞̃=√𝜌𝐿𝜌𝑅 (43)
𝑢̃=√𝜌𝐿𝑢𝐿+√𝜌𝑅𝑢𝑅
√𝜌𝐿+√𝜌𝑅 (44)
𝑣̃=√𝜌𝐿𝑣𝐿+√𝜌𝑅𝑣𝑅
√𝜌𝐿+√𝜌𝑅 (45)
𝑤̃=√𝜌𝐿𝑤𝐿+√𝜌𝑅𝑤𝑅
√𝜌𝐿+√𝜌𝑅 (46)
𝐻̃=√𝜌𝐿𝐻𝐿+√𝜌𝑅𝐻𝑅
√𝜌𝐿+√𝜌𝑅 (47)
𝑈̃=(𝑢𝑖⃗̃+𝑣̃𝑗⃗+𝑤̃𝑘⃗⃗)𝑛⃗⃗−𝑉𝐺⃗⃗⃗⃗⃗𝑛⃗⃗ (48)
∆𝑝=𝑝𝑅−𝑝𝐿 (49)
In the case of the experimental analysis carried on for studying the engine core flow, for
many applications, the Reynolds number has large values and the flow regime is turbulent.
Time and length scales of the turbulent flow regime s are very small. In orde r to apply the
turbulen t regime conditions directly from the Navier -Stokes equations, the grid resolution
must be as high as possible [4].

21 Numerical Simulations of Flow in Axial Compressor System, Preparatory Steps for Active Control

INCAS BULLETIN, Volume 10, Issue 3/ 2018
Fig. 1 – Schematic diagram of the turbofan engine, illustrating the Axial Compressor and Turbine Operation ,
Rolls -Royce [1]

Fig. 2 – Schematic diagram of the compressor and turbine stages, turbofan engine, Rolls -Royce [1]

Fig. 3 – Schematic diagram of multiple -staged a xial flow compressor [ 2]

Irina -Carmen ANDREI, Gabriela STROE 22

INCAS BULLETIN, Volume 10, Issue 3/ 2018
Fig. 4 – Temperature field distribution in Turboprop [2]

Fig. 5 – Pressure field distribution in Turboprop [2 ]

Fig. 6 – Flow field distribution in Turboprop [2 ]
The flow through modern axial compressors systems is high ly complex. The flow is 3D
and unsteady due to the relative motion between the successive blade rows and the

23 Numerical Simulations of Flow in Axial Compressor System, Preparatory Steps for Active Control

INCAS BULLETIN, Volume 10, Issue 3/ 2018 occurence of the viscous effects within each row. The flow is also transonic where there
regions of subsonic and supersonic flow coexist. Supersonic flow generally appears near the
rotor tip leading edge , usually for large diameter blades, where the highest rotational
velocit ies are combined with the axial flow velocit ies, and the relative Ma ch number
frequently exceeds the unity.
3. NUMERICAL SIMULAT ION AND CONCLUSION S
The fluid flow path of an axial compressor system is convergent, with the meaning that the
cross -sectional area in the direction of flow decreases . The area is diminished proportion ally
with the increased density of the air , as the compression progresses from stage to stage. Each
stage of an axial compressor produces a small compression pressure ratio at a high
efficiency , with the concern to preserve the subsonic flow regi me within the compressor
system . For this reason, for the hi gh pressure ratios, the multiple stage d axial flow
compressors are used.
Axial compressors are also more compact , have a smaller cross section and provide
higher values for the pressure ratio, in comparison with the centrifugal compressor, which
are the most important advantages.
For running with the best efficiency, the compressor system must operate at constant
axial velocity.
At high pressure ratios, a single staged axial compressor does not work efficient ly, in
comparison with multiple staged axial compressor.
When running the engine from lower speeds up to higher speeds, in case of multiple –
staged axial flow compressors, oper ating with high values of the pressure ratio, the design of
the engine must be of twin -spool or triple -spool type .
In order to achieve more flexibility and a more uniform loading of each axial
compressor stage, the twin spool or triple spool construction with two or three different
rotational speeds is generally used in high pressure ratio axial compressor systems .

Fig. 7 – Three stage d axial flow compressor [ 8]

Irina -Carmen ANDREI, Gabriela STROE 24

INCAS BULLETIN, Volume 10, Issue 3/ 2018
Fig. 8 – Powerplant engine control [8]

Fig. 9 – Engine instruments [8]
In order to im prove the performances of the axial compressor system, with the
referrence to significantly reduce the pressure losses due to shock waves and boundary layer
separation, the design of the blade row cascades can be modified by the use of the sweep
effect, whi ch eventually lead to diminish of the flow velocity at tip blad, from supersonic to
subsonic, such that to prevent the occurence o f the shock waves at tip blade.
The numerical simulations of the flow through the transonic first rotor blade test case ,
have been performed for the basic reference blade, shown in Fig. 10, and the sweep effect
has been in vestigated for the swept blades, shown in Fig. 11.
The mathematical model is based on the RANS equations; following the numerical
simulations using the AN SYS FLUENT CFD code, the flow parameters have been
determined .

25 Numerical Simulations of Flow in Axial Compressor System, Preparatory Steps for Active Control

INCAS BULLETIN, Volume 10, Issue 3/ 2018
Blade view from hub to tip Airfoil overlap – cross section in rotor blade
Fig. 10 – Test Case – NASA Blade from the rotor first stage of a mu ltistaged axial flow compressor

Sweep angle [deg], from hub to tip: (0,5,7,5,0) Sweep angle [deg], from hub to tip: (0, -5,-7,-5,0)
Fig. 1 1 – Swept blade versus basic blade

a) Basic blade

b) Forward swept blade (sweep angle =7 [deg]) c) Backward swept blade (sweep angle = – 7 [deg])
Fig. 1 2 – Contours of the relative Mach number , at mid section

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INCAS BULLETIN, Volume 10, Issue 3/ 2018 The airfoils at mid section have been considered for the 2D flow numerical analysis.
The effect of sweep has been highlighted by the cases: forward sweep, with the sweep angle
being considered 7 [deg], and backward sweep, wi th the sweep angle = – 7 [deg].
Fig. 12 illustrates the co ntours of relative mach number.
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