Analysis and Optimization of axial flow compressor (Armstrong

Siddeley Viper 12)

A Project Report Submitted By

Rishav Thakur 210303101051
Dhiraj Kumar Sah 210303101050
Shiv Narayan Chaudhary 210303101048
Pranav Kishor Yadav 210303101052
Devendra Singh Lodh 210303101125

In partial fulfilment for the award of the Degree of

BACHELOR OF TECHNOLOGY AERONAUTICAL ENGINEERING
Under the guidance of
Asst. Professor Mansha Kumari
Assistant Professor

PARULUNIVERSITY

VADODARA
October – 2024

1
 A PROJECT REPORT ON

Analysis and Optimization of axial flow compressor (Armstrong

siddeley Viper 12)

AS PROJECT

ACADEMIC YEAR 2024–2025

DEPARTMENT OF AERONAUTICAL
ENGINEERING

PARUL UNIVERSITY

P.O. Limda, Tal: Waghodia

Dist. Vadodara- 391760, Gujarat, INDIA

 CERTIFICATE

DATE:

THIS IS TO CERTIFY THAT THE DISSERTATION ENTITLED “ ANALYSIS AND

OPTIMIZATION OF AXIAL FLOW COMPRESSOR (ARMSTRONG
SIDDELEY VIPER 12)” HAS BEEN CARRIED OUT BY

Rishav Thakur 210303101051

Dhiraj Kumar Sah 210303101050

Shiv Narayan Chaudhary 210303101048

Parnav Kishor Yadav 210303101052

Devendra Singh Lodh 210303101125

UNDER MY GUIDANCE IN FULFILLMENT FOR THE DEGREE OF BACHELOR OF

ENGINEERING IN AERONAUTICAL ENGINEERING (6TH SEMESTER) OF PARUL
UNIVERSITY, DURING THE ACADEMIC YEAR 2023-2024

AKHIL C.K Asst. Professor Mansha Kumari

Project Supervisor Project Coordinator

Dr. BALAJI K
External faculty
Head of Department
and Associate Professor
 ACKNOWLEDGEMENT

We give the whole credit and praise to the omnipresent for his mighty grace
and blessings throughout the course of this research.

Behind any major work undertaken by an individual, there lies the contribution
of the people who helped him to cross all the hurdles to achieve his goal.

It gives us immense pleasure to express our sense of sincere gratitude towards

our respected guide (Asst. Professor Mansha Kumari) for her persistent,
outstanding, invaluable co-operation and guidance. It is our achievement to
have her guidance. She is a constant source of encouragement and momentum
that any intricacy becomes simple. We gained a lot of invaluable guidance and
prompt suggestions from her during the entire project work. We will be
indebted to her forever and we take pride in working under her.

We also express our deep sense of regard and thanks to the Head of
Department (Dr. Balaji k). We feel very privileged to have their precious advice,
guidance, and leadership. As an advisor, she has a great influence on us, both
as a person and as a professional.
 TABLE OF CONTENT

1. INTRODUCTION

2. LITERATURE REVIEW

3. OBJECTIVE

4. FUTURE SCOPE

5. RESEARCH GAP

6. METHODOLOGY

7. RESULTS AND ANALYSIS

8. CONCLUSION

9. REFERENCES
 Nomenclature

1) α - Absolute flow angle

2) Β - Relative flow angle
3) C - Absolute flow velocity
4) w - Relative flow velocity
5) U - Blade velocity
6) Ca - Axial velocity
7) Cθ, Wθ – Swirl or whirl component of velocity
8) T0 - Stagnation temperature
9) P0 - Stagnation Pressure
 INTRODUCTION
Armstrong Siddeley Viper
The Armstrong Siddeley Viper is a British turbojet engine, originally designed
and manufactured by Armstrong Siddeley, and later produced by Rolls-Royce.
The Viper has been widely used in various military and civilian aircraft, known
for its compact size and versatility.
First Flight: The Viper engine first ran in 1951. It was designed as a small, light-
weight turbojet, initially for use in missiles and drones but later found
applications in small jet aircraft.
Applications:
The Viper engine powered various aircraft over its long production life,
particularly in training jets and light attack aircraft. Some notable aircraft
powered by Viper engines include:
• Hunting Percival Jet Provost: A British military training aircraft.
• BAC Strikemaster: A light attack aircraft and trainer, based on the
Jet Provost.
• Aermacchi MB-326: A popular Italian military jet trainer and light
attack aircraft.
• Handley Page Jetstream: A small regional airliner and military
transport.
Specifications
General characteristics
• Type: Turbojet
• Length: 64.0 in (1,625 mm)
• Diameter: 24.55 in (624 mm)
• Dry weight: 549 lb (249 kg)
Components
• Compressor: Seven stage axial
• Combustors: Annular, 24 burners
• Turbine: Single stage
• Fuel type: AVTUR, AVTAG
• Oil system: scavenge, metered
Performance
• Maximum thrust: 2,700 lbf (12 kN) at 13,800 rpm
• Overall pressure ratio: 4.3:1
• Air mass flow: 44 lb/s (20 kg/s)
• Specific fuel consumption: 1.09 lb/ (hr lbf)
Axial Flow Compressor
From an early stage in the history of the gas turbine, it was
recognized that the axial flow compressor had the potential for
both higher pressure ratio and higher efficiency than the centrifugal
compressor. Another major advantage, especially for jet engines,
was the much larger flow rate possible for a given frontal area.
These potential gains have now been fully realized as the result of
intensive research into the aerodynamics of axial compressors: the
axial flow machine dominates the field for large powers and the
centrifugal compressor is restricted to the lower end of the power
spectrum where the flow is too small to be handled efficiently by
axial blading.
Early axial flow units had pressure ratios of around 5: 1 and
required about 10 stages. Over the years the overall pressure ratios
available have risen dramatically, and some turbofan engines have
pressure ratios exceeding 40: 1. Continued aerodynamic
development has resulted in a steady increase in stage pressure
ratio, with the result that the number of stages for a given overall
pressure ratio has been greatly reduced. There has been in
consequence a reduction in engine weight for a specified level of
performance, which is particularly important for aircraft engines. It
should be realized, however, that high stage pressure ratios imply
high Mach numbers and large gas deflections in the blading which
would not generally be justifiable in an industrial gas turbine where
weight is not critical; industrial units, built on a much more
restricted budget than an aircraft engine, will inevitably use more
conservative design techniques resulting in more stages.
 Fig : Blade Row

Elementary theory
The working fluid in an axial flow compressor is normally air, but
for closed-cycle gas turbines other gases such as helium or carbon
dioxide might be used. The analysis which follows is applicable to
any gas, but unless otherwise noted it will be assumed that the
working fluid is air. A sketch of a typical stage is shown in Fig. 5.3.
Applying the steady flow energy equation to the rotor, and
recognizing that the process can be assumed to be adiabatic, it can
readily be seen that the power input is given by:
w=mcp(T 2−T 1)

Repeating with the stator, where the process can again be assumed
adiabatic and there is zero work input, it follows that T02 = T03. All
the power is absorbed in the rotor, and the stator merely transforms
kinetic energy to an increase in static pressure with the stagnation
temperature remaining constant. The increase in stagnation
pressure is accomplished wholly within the rotor and, in practice,
there will be some decrease in stagnation pressure in the stator due
to fluid friction. Losses will also occur in the rotor and the
stagnation pressure rise will be less, than would be obtained with an
isentropic compression and the same power input. A T-s diagram
for the stage, showing the effect of losses in both rotor and stator, is
also shown
Fig: h-s or T-s diagram.

It enters the compressor at point A (low pressure, low temperature).As it

passes through the rotor blades, its kinetic energy increases, and it moves
to point B (higher pressure, still low temperature).The stator blades
further increase the pressure (point C), but the temperature remains
relatively constant. The air exits the compressor at point D (higher
pressure, higher temperature).

Pressure, Temperature and Absolute velocity Rise across Stage (Static

and Stagnation)
 Fig: Pressure, temp., velocity Rise
Pressure Rise:
As air flows through an axial compressor, it passes through multiple stages.
Each stage contributes a small pressure increase.
The pressure ratio per stage varies based on the application:
Industrial Compressors (Subsonic Flow): Achieve pressure ratios of 1.05 to
1.2 per stage with efficiencies of 88% to 92%.
Aerospace Compressors (Transonic Flow): Operate at higher pressure ratios
(1.15 to 1.6 per stage) with efficiencies around 80% to 85%.

Temperature Rise:
The temperature rise across an axial flow compressor is closely related to the
pressure rise.
As air is compressed, its temperature increases due to the work done on it
by the blades.
However, the actual temperature rise might be slightly lower than estimated
from velocity triangles due to factors like losses and real-world behaviour.

Velocity Triangles
 Each stage of an axial flow compressor consists of a set of rotor blades
(attached to a rotating hub) and stator blades (fixed to the casing).
Velocity triangles help us understand the flow within each stage. These
triangles show the relative and absolute velocities of the air as it passes
through the blades. The rotor blades accelerate the air, adding kinetic
energy. The stator blades then convert this kinetic energy back into
pressure energy.
Our air molecule enters the rotor with an absolute velocity, let’s call it V
absolute. It also has a flow angle (β₁) with respect to the axial direction. As
it passes through the rotor, its relative velocity (V relative) forms the first
side of our triangle. The rotor’s tangential velocity (V tangential) is ω times
the rotor radius.

So, Vrelative = Vabsolute - Vtangential

Energy Transfer and Energy Transformation

1
The quantity 2 (C 21-C 22 ¿or its differential value is the change in the kinetic
energy of the fluid through the machine in the absolute frame of
coordinates. This brings about a change in the dynamic head of the fluid
through the machine.
1 2 2
The quantity 2 (u ¿ ¿ 1 −u2)¿ is the change in the centrifugal energy of the 2
fluid in the machine. This arises simply due to the change in the radius of
rotation of the fluid. This term causes a change in the static head of the fluid
through the rotor.
1 2 2
The quantity (W ¿ ¿ 2 −W 1)¿ Is the change in the kinetic energy of the fluid
2
In the relative frame of coordinates. This also causes a change in the static
Head of the fluid across the rotor.

Degree of Reaction.

In turbomachinery, degree of reaction or reaction ratio (denoted R) is

defined as the ratio of the change in static pressure in the rotating blades
of a compressor or turbine, to the static pressure change in the
compressor or turbine stage. Alternatively, it is the ratio of static enthalpy
change in the rotor to the static enthalpy change in the stage. Various
definitions exist in terms of enthalpies, pressures, or flow geometry of the
device. In case of turbines, both impulse and reaction machines, degree of
reaction is defined as the ratio of energy transfer by the change in static
head to the total energy transfer in the rotor:
Isentropic enthalpy change ∈rotor
R=
Isentropic enthalpy change∈ stage
For a gas turbine or compressor, it is defined as the ratio of isentropic heat
drop in the moving blades (the rotor) to the sum of the isentropic heat
drops in both the fixed blades (the stator) and the moving blades:

Isentropic heat drop∈rotor

R=
Isentropic heat drop∈ stage
In pumps, degree of reaction deals in static and dynamic head. Degree of
reaction is defined as the fraction of energy transfer by change in static
head to the total energy transfer in the rotor:
Static pressure rise ∈rotor
R=
Total pressure rise∈ stage

Introduction to CFD and its uses

Computational Fluid Dynamics (CFD) is a powerful branch of fluid

mechanics that employs mathematical models and numerical
algorithms to simulate, analyse, and predict the behaviour of fluid
flows. By solving the fundamental governing equations, such as the
Naviers-Stokes equations and continuity equation, CFD enables
engineers and scientists to study complex fluid flow problems that are
difficult or impossible to analyse through experimental or analytical
means.

Through the use of advanced computer software and high-performance

computing, CFD offers a cost-effective and time-efficient method to
investigate fluid dynamics in a wide range of applications. These include
aerospace, automotive, water and wastewater treatment, chemical
processes, and environmental engineering, among others. By providing
detailed insights into fluid behaviour, CFD assists in optimizing designs,
enhancing performance, and improving safety in various industries.
As a multidisciplinary field, CFD integrates concepts from mathematics,
physics, and computer science. Its continuous advancements in
numerical methods, turbulence modelling, and computational
resources have greatly expanded the scope and accuracy of fluid flow
simulations, making it an indispensable tool in modern engineering and
scientific research

ANSYS Fluent can be used to simulate a wide range of problems,

including:
• The flow of gases and liquids through pipes, valves, and other
components
• Heat transfer in heat exchangers, furnaces, and other equipment
• Fluid-structure interactions, such as the flow around an aircraft
wing or the motion of a floating object in a fluid
• Multiphase flows, such as the flow of a gas and a liquid through a
pipe
• Reacting systems, such as the combustion of fuel in an engine
• Structural analysis, including stress and strain analysis
 LITERATURE SURVEY

1) Aerodynamic Design of a single stage Axial Flow Compressor using CFD

approach
Author: Kiran D Chaudhary, Prof. Dr. N. A. Wankhede.
Observation:
• The prime requirement of Gas turbine engine manufacturers is efficiency
and power to weight ratio. To increase maximum pressure in compressor.
It can be achieved by running the compressor at higher speed.
• It results in the either high subsonic or transonic flow. But the sonic flow
creates high losses in the cascade because of the formation of shock
waves.
• Cascade losses and efficiency: The static pressure rise through a
compressor cascade depends on the deflection of the fluid through it.
Therefore, a maximum value of the fluid deflection is desirable, but on
account of stalling and the associated cascade losses, this is carefully
chosen.
tip speed 167.7 m/s.
pressure rise 1.21
Number of rotations 14800 RPM
Power 276.5 KW

2) Aerodynamic Performance of Low-Speed Axial Flow Compressor Rotors

with Sweep and Tip Clearance

Author: P. V. Ramakrishna and M. Govardhan

Observation:
• The study concludes that the effects of tip clearance are more in swept
rotors compared to unswept rotors, impacting total pressure rise and
efficiency.
 • The swept blade design effectively guides boundary layer fluid towards
the trailing edge, reducing fluid accumulation near the tip and enhancing
performance.
• However, the presence of tip clearance significantly increases losses,
particularly in swept rotors.
• The results indicate that swept rotors prevent the radial migration of low-
energy boundary layer fluid towards the casing, which is favorable for
maintaining a higher stall margin. Overall, while blade sweep offers
performance benefits, managing tip clearance is crucial to maximizing
these advantages.
• Overall, while blade sweep offers performance benefits, managing tip
clearance is crucial to maximizing these advantages.

3) Span wise Mixing in Axial-Flow Turbomachines

Author: G.G. Adkins, Jr. L. H. Smith, Jr.

Observation:
• The study concludes that secondary flows significantly influence the
overall flow process in axial-flow turbomachines, particularly in machines
with low aspect ratios and high aerodynamic loadings.
• The developed method provides a way to account for these effects in
design calculations, helping to better match theoretical predictions with
observed data.
• However, the complexity of mixing and secondary flow phenomena
suggests that further development and refinement of the models are
needed to enhance their accuracy and applicability, particularly in
conditions deviating from the idealized scenarios considered in the study.

4) DESIGN OPTIMIZATION OF AN AXIAL FLOW COMPRESSOR FOR INDUSTRIAL

GAS TURBINE
Author: Nilesh P. Salunke1, S. A. Channiwala2, Juned A. R. A3
Observations:
• The study employs 1D, 2D, and 3D simulations, along with CFD and FEA
tools, to evaluate and optimize design, demonstrating the importance of
simulation in the design process.
• The optimized design provided a wide stall margin and low loss
coefficients, showcasing efficiency under different speed and flow
conditions.
 • Compressor performances validation on preliminary design phase: PD
MAP is the tool to calculate compressor curves necessary for initial
compressor characteristics assessment.
• The performance characteristic curve obtained in AXMAP it indicates the
compressor is matching the performance i.e. delivering the given mass of
air at designed outlet pressure not only this, the compressor having
sufficiently wide range for stall and choke margin.

5) Performance Evaluation of Axial Flow Compressor Using Stages

Characteristics
Author: Avwunuketa, Ayedun Alex, Adamu, Mohamed Lawal²,Ajao,Tofunmi
Ayodele³
Observations:
• The performance characteristics of axial flow compressors due to
variation in size of blade, pressure, and temperature and shaft rotational
speed determined the output variables such as work output and
efficiency.
• The performance depends upon the blade diameter, mass flow rate,
density of the flowing fluid, stage pressure ratio, stage delivering
pressure and temperature. For the axial compressor to deliver
compressed air to the combustor for onward delivery to the turbine,
each stage should be able to attain a better performance that is better
than the preceding stage.

6) Influence of the Blading Huff Surface Roughness on the Aerodynamic

Behavior and Characteristic of an Axial Compressor
Author: K. Bammert G. U. Woelk
Observation:
• The effects of soiling, corrosion and erosion can be reduced and the life
of the blading can be increased by filtering.
 • Surface roughness on axial compressor blades significantly impacts the
compressor’s aerodynamic behavior, causing increased flow losses,
reduced efficiency, and narrowing the operating range.
• Rough blades lead to decreased mass flow rate, pressure ratio, and
efficiency.
• Increased roughness results in formation of the boundary layer, which
can lead to flow separation.

7) Behavior of Tip Leakage Flow behind an Axial Compressor Rotor

Author: M. inoue IVI. Kuroumaru M. Fukuhara
Observations:
• In experimental studies has been concerned with the better
understanding of the complex inviscid and viscous effects in the casing
boundary-layer flow.
• The experiments were done at the flow rate corresponding to the same
incidence angle of the rotor blade in the middle span (not at the same
flow rate) to get rid of the effect of incidence variation caused by
boundary layer blockage.

8) Design and Analysis of Axial Flow Compressor Blade Using Different Aspect
Ratios with Different Materials
Author:[1] K. Sravanmathur, [2] R. Murugan, [3] V.S. Hariharan
Observation:
• The value of velocity is lower for AR1 and higher velocity for AR2.
That means velocity is increasing with aspect ratio.
• More pressure for AR1 and lower for AR2. Pressure is decreasing with
increase in aspect ratio. Mass flow rate is higher for AR2 compared to
AR1.
• In other words, we can say that the mass flow rate is decreasing with
decrease in aspect ratio.
• Deformation values (stress) depend on the material used in the
blades. Titanium alloys have minimum deformation compared to
others materials (nickel alloys, chromium steel).

9) The Effect of Pressure ratio on the Compression work done by an Axial flow
Compressors
Author: Ekong, Godwin I.
Observation:
 • The application of Ideal Final Results in the Establishment and
Management of a Cold storage facility for rural areas, were applied in the
design and the selection of materials for manufacturing of the in-house
axial compressor.
• Polyvinyl chloride (PVC) which is a thermoplastic polymer was used at
the inlet of the compressor

10) Axial flow compressor design

Author: S. J. Gallimore
Abstract: To write a short paper on axial flow compressor design. To simplify
this paper the design process can, somewhat artificially, be split into roughly
four stages: preliminary design, through flow design, blading design (two-
dimensional) and blading design (three-dimensional).
Observations:
• The most important phenomena in compressors, stall, or surge, has not
been touched upon. It is now generally accepted that stall occurs in
compressors when the flow breaks down under increasing adverse
pressure gradients.
• Two types of breakdowns have been identified, one associated with a
short length scale disturbance known as a ‘spike’, and the other
identified with a longer length scale known as a ‘modal oscillation’.
• The ‘modal oscillations’ are related to an instability of the whole
compression system while the ‘spike’ type of instability can be related to
a more local disturbance in the flow, commonly caused by high incidence
at a rotor tip.
• This breakdown occurs near rotor tips in discrete patches which rotate in
the same direction as the rotor but at between 30 and 70 per cent of the
rotor speed.
• This is called part span rotating stall and can have more than one stall cell
initially but as the stall develops, they tend to coalesce into one cell.
PRELIMINARY DESIGN:
• The basic inputs to the design will be requirements for a certain flow
capacity, pressure ratio, efficiency, and surge margin for a range of engine
operating points.
• The basic tools to deal with this part of the design are mean line
performance prediction programs.
THROUGHFLOW DESIGN:
• This means introducing the radial or span wise dimension into the design
using a through flow procedure.
• The simplest of these again does not include the major, significant flow
features for similar reasons to the preliminary design methods.
• In the simplest through flow methods the effect of the end wall
boundary layers is included by specifying a blockage and extra loss in the
calculations.
• BLADING DESIGN (TWO-DIMENSIONAL)

11) Design and CFD analysis of multistage AFC

Author: Lakshya Kumar, Dilipkumar B

Observations:

• the entry to R1, pressure rise is nearly constant over the 50% of the
chord for CP (Choked point), DP (Design point) and NS (Numerical stall)
due to rapid acceleration of the flow.
• Again, at the entry to R2 and R3 there is a sudden dip in the pressure due
to rotor LE acceleration followed by steep diffusion in rest of the rotor
part for DP and NS.
• R1 has huge flow reversal till 60% span for CP and DP which increases at
NS almost
• till 85 % span with spike in total pressure near the blade tip regionR2 exit
profiles have similar distribution over
 • the entire span for all three conditions with flow reversal up to 75% of
span. R3 exit has almost uniform total
• pressure distribution for CP except end-wall region.

12) Parametric study on axial compressor performance in design condition

Author: Sarallah Abbasi, and Ali Joodaki

Observations:

• Blade sweep angle increasing the blade sweep angle causes improving
the flow behavior in axial fan and reducing it, having a completely
contrary result. increasing the sweep angle leads to enhance efficiency
and pressure ratio.
• Velocity distribution is more sensitive to the blades sweep angle relative
to efficiency and the pressure ratio.
• Thickness of blade profile: The effect of the thickness on the hub is
greater than the thickness of the tip, and its increase leads to reduce
both efficiency and pressure ratio.
• Number of blades: the velocity minimum occurs in the stator, and with
reducing the blades number the minimum velocity decreases.

13) Axial-Flow Compressor Performance Prediction in Design and Off-Design

Conditions through 1-D and 3-D Modeling and Experimental Study Journal of
Applied Fluid Mechanics
Author: Peyvan and A. H. Benisi
Observations:
• Comparison between one-dimensional modeling results with
experimental results shows good agreements for different rotational
speeds.
• Maximum performance prediction difference for pressure ratio and
isentropic efficiency are 2.1 and 3.4 percent respectively.
• Since the experimental results are obtained for a compressor, which is
affected by other components of the engine, working inside a gas turbine
engine, it can be concluded that the 1D modeling sufficiently predict the
performance of a compressor while working in an assembly.

14) DESIGN AND OPTIMIZATION OF AXIAL FLOW COMPRESSOR2020 JETIR

Author: NADIPENNAGARI PEDDA VENKATESWARLU
Observations:
• In this theory, a pivotal stream blower is planned and displayed in 3D.
• The current plan has 30 edges, in this proposal it is supplanted with 20
edges and 12 cutting edges.
• Titanium combination and Nickel compound are high quality materials
than Chromium Steel. The thickness of Titanium composite is not as
much as that of Chromium Steel and Nickel amalgam. So, utilizing
Titanium compound for blower edge diminishes the heaviness of the
blower Structural investigation is done on the blower models to confirm
the quality of the blower.
• The pressure esteem is less for titanium composite than Nickel
combination.

15) Aerodynamic Analysis of Multistage Turbomachinery Flows in Support of

Aerodynamic Design
Author: John J. Adamczyk
Observations:
 Historical development of turbomachinery flow models.
 Challenges in simulating complex, unsteady flows in multistage
turbomachines.
 The role of the average-passage flow model for time-average flow field
analysis.
 Impact of unsteady deterministic flows, such as rotor-stator
interactions, on aerodynamic performance.
 Techniques for spanwise and circumferential redistribution of total
enthalpy and momentum.
 Importance of wake recovery processes and their impact on efficiency
and loss generation.

16) A Steady State Analysis on Axial Flow Compressor Subjected to

Stresses and Deformation
Author: Putta Ratna Raju, Shaik Moulali

Observations:
  Analysis is done by using ANSYS and the parameters such as deformation
and the equivalent stresses are obtained for both the materials and are
compared.
 So, using stainless steel for compressor blade decreases the strength of
the compressor Structural analysis is done on the compressor models to
verify strength of the compressor to verify the strength of the
compressor.
 By use of 29 blades stress using 29 blades the stresses are increasing, but
are within the limits So it concluded that using magnesium than stainless
steel for 29 blades is better for compressor

17) The Off-Design Analysis of Axial-Flow Compressors

Author: W. JANSEN, W. C. MOFFATT

Observations:

 The study presents a numerical method for predicting compressor

performance, such as pressure rise and efficiency, for known blade
geometries under various operating conditions. It divides the aerodynamic
analysis into determining the flow field and blade performance.
 The analysis utilizes continuity and momentum conservation equations,
incorporating empirical adjustments for entropy and losses.
 The streamline-curvature method is applied to estimate flow conditions
between blade rows.
 The methodology includes corrections for axial velocity variations, blade
thickness effects, and compressibility, based on two-dimensional cascade
data. It adjusts flow angles and losses to better match real compressor
behaviours.
 The paper compares the predicted results with experimental data from
various compressors, showing good agreement when appropriate
boundary-layer blockage factors are applied.
18) Performance enhancement of an axial flow compressor using
active flow control technique

Author: Prithesh Pinto and Srinivas G

Observations:

 This research article focuses on improving the efficiency of an axial flow

compressor stage, enhancing the stall margin, and optimizing flow
parameters.
 The steady simulation of the NACA 65 rotor blade validates the
improvement in stall margin.
 increase in average static pressure at the exit boundary condition, there
is a decline in mass flow rate and a hike in the pressure rise coefficient
(w).
 Similarly, as the mass flow coefficient (/) decreases there is an increase in
isentropic efficiency (g) reaching its peak at the design point and then
decreases sharply as the system approaches the stall point

19)Optimized Shroud Design for Axial Turbine Aerodynamic

Performance
Author: L. Porreca, A. I. Kalfas, R. S. Abhar
Observations:
 The full shroud configuration showed the highest aerodynamic efficiency
due to minimal tip leakage losses. In contrast, the partial shroud
exhibited increased aerodynamic losses due to flow expansion in the
cavity and formation of strong tip vortices.
 The optimized shroud (EPS), which extends the platform towards the
trailing edge to prevent flow expansion, improved aerodynamic
performance by about 0.6% compared to the partial shroud.
 The EPS configuration balanced efficiency and reduced mechanical
stress on the blades.
 Hub regions, with the partial shroud case showing higher losses due to
mixing of the leakage flow with the main stream.
20) Five-Stage Axial Flow Compressor for Gas Turbine

Author: Khema Theint

Observations:

 The goal of this paper is to calculate the blade design of axial flow
compressor. Main objective of the compressor usage is to compress the
fluid and to deliver higher pressure than its original pressure.
 Number of blades is increasing with pressure in stage.
 Temperature rises constantly with 15K in every stage.

21) Flow studies in a Mixed Flow Compressor Stage for Small Gas
Turbine Applications

Author: Logesh N Dr. R. Rajendran, Emandi Rajesh

Observations:

 Recently mixed flow compressors are being used for the application of
moderate flow and high work coefficient applications.
 In this paper mixed flow stage is designed for 1.75 kg/s and 5 pressure
ratios to fit in a geometric envelope of 220mm diameter and axial length
of 135mm.
 Numerical analysis on the model shows that 4.42 total to total pressure
ratio for 73.8 % isentropic efficiency.
 Increase the blade angle at hub to bring the incidence at negative zone
for better operating range.
 Relaxing the hub curvature [with increased diameter] of diagonal diffuser
will make the flow uniform from hub to tip at diffuser exit. Which also
improve the stall margin.

22) ACTIVE FLOW CONTROL TO IMPROVE THE AERODYNAMIC AND ACOUSTIC

PERFORMANCE OF AXIAL TURBOMACHINES
Author: L. Neuhaus and W. Neise
Observations:
 The study addresses the adverse effects of tip clearance flow on the
aerodynamic and acoustic performance of axial turbomachines. The flow
across the tip of rotor blades, driven by pressure differences, leads to
secondary flows that can result in increased noise (tip clearance noise)
and diminished aerodynamic efficiency.
 Experiments with steady air injection into the tip clearance gap
demonstrated that it is possible to reduce noise and improve efficiency
with small mass flow rates.
 Higher rates further enhance aerodynamic performance but at the cost
of increased noise levels.
 The study used a low-speed, high-pressure axial fan with adjustable tip
clearances and air injection through slit nozzles.
 Injection methods, concluding that unsteady air injection may be more
effective for pressure enhancement and noise reduction with lower mass
flow rates, while steady injection benefits efficiency.
23) Recent developments of axial flow compressors under transonic flow
conditions

Author: Srinivas G1 Raghunandana K*2 and Satish Shenoy B

Observations:

 The objective of this paper is to give a holistic view of the most advanced
technology and procedures that are practiced in the field of
turbomachinery design.
 There are varieties of configurations on non-uniformities possible at the
entry face to rotor, like the one associated with pressure field or velocity
and temperature and so on. Point to note here is they are not
independent but are coupled effects. One discontinuity leads to the
deviation of others.
 Some region of cross section subjecting to discontinuities will have its
implications on remaining blades flow passages and therefore full rotor
analysis is very much inevitable. From the literature it is understood that
they were no attempts made to generate these discontinuities at the
entry face.
 24) Simulation of Rotating Stall in a Whole Stage of an Axial Compressor
Author: Nicolas Gourdain, Stephani Burguburu, Francis Leboeuf Guy Jean Michon4
Observations:
 This paper investigates the ability of an unsteady flow solver to simulate the rotating
stall phenomenon in the full annulus of an axial compressor stage.
 A comparison with experimental data indicates that the simulation correctly estimates
the stability limit.
 However the rotating stall flow patterns are different. While measurements show only
one full span rotating stall cell (40 Hz), the simulation shows first a part span stall
with 10 cells (790 Hz) that evolves then towards a full span stall with 3 cells (170 Hz).
 The effects of downstream volumes and inlet distortions are also discussed, showing
the necessity to consider the whole geometry to correctly predict the rotating stall
frequency

25) Enhancement of aerodynamic performance of axial compressors

utilizing natural aspiration through optimized casing circumferential
slot
Author: Peyman Ghashghaie Nejad & RezaTaghavi Zenouz
Observation:
 This paper presents results of aerodynamic enhancement of axial compressors
utilizing optimum natural aspiration through a circumferential slot made within the
casing wall upstream the rotor blades row.
 Geometries of the slot walls were optimized to find the maximum stall margin and
pressure ratio of the compressor.
 The optimum case is accompanied by more pressure rise coefficients and wider stall
margins for all the three blades tip clearance sizes. For example, at 1.7% blades tip
clearance, the total pressure ratio and stall margin are increased by 4.2% and 3%,
respectively.

Future Scope
 Improving stability: more works needs to be done to accurately estimate the secondary
flow losses and compressor stability to achieve stable operation.
 Reduced complexity in design as more advanced and more workability is induced.
 Due to high pressure ratio and high multistage capacity, it can be utilized for in-home
built projects for aerodynamics and fluid mechanics institutions.
 Noise and Vibration Reduction: Advances in reduction acoustic and vibrational
impacts to meet stricter regulations and improve user’s comfort.
 Designing turbomachinery system that are integrated with advanced control systems
for real time optimization and improve efficiency.

Research Gap:
 Obtaining the engine parameters lead to assumptions of various parameters.
 Finding the best solution for the CFD modelling.
 The various parameters such Aspect ratios, Tip clearance, Blade height and blade
pitch were not provided. Designing require the ability to decide the appropriate fitting
parameters and its desired results.

Objectives
 Optimization of Flow by varying various flow parameters such as mass flow rate.
Also, by varying other parameter such as blade pitch, varying aspect ratio.
 To increase pressure rise of a compressor stage.

Velocity Triangle for a compressor stage

 Calculation Analytical

Ca = 150 m/s α1 = 25° R = 50 % Dm

= 0.624m
T0 = 288 k N = 10,000 rpm Cp = 1005 J/kg k
Y= 1.4
dt = 0.47 m dh = 0.5
dt
dh = 0.5 x 0.47 = 0.235 m
dm = dh + dt = 0.47 + 0.235 = 0.3525 m
2 2

m = ρ ACa
20 = ρ(2 Π rm x h) Ca
20 = 1.066 x 2 x Π 0.3525h x 150
2
h = 0.112946 m

U = Π DmN = Π ×0.3525×10000 = 184.56 m/s

60 60
C1 = Ca = 150 = 165.9423 m/s
cosα1 cos25°

Cθ1 = C1 sinα1 = 165. 9423×sin25° = 69.9423m/s

T1 = T01 − C1² = 288− (165.50) ²

2Cp 2×1005
T1 = 288 − 13.629 = 274.37K
R = Ca [ tanβ1 + tanβ2 ]
2U
[ α1 = β2 ]
0.5 = Ca [ tanβ1 + tan25° ]
2U
0.5 = 150 [ tanβ1 + tan25° ]
2×184.56
tanβ1 = 1.2304 − 0.4663

∴ β1 = 37.4°
tanβ1 = 0.7641

C1 = Ca = 150 = 188.7679 m/s

cosα1 cos37.4°

( )
y
T1
P1 = T 01
y−⊥

P01
P1 = 1 × ( 288 )
1⋅ 4
274.37 0.4

P1 = 0.84 bar or 0.84×10 pascal 5

ρ1 = P1 = 0.84×10^5 = 1.066 kg/m³

RT1 287×274.37

Seven Stage;

( PP06 )x( PP05 ) x ( PP04 ) x( PP 0403 ) x ( PP02 ) x( PP01 )=4.3

07 06 05 03 02

( ) P02
P 01
=¿ 4.3 0.166 = 1.275

( )
y
P 02 T 02 y−⊥
( )=¿
P 01 T 01

∴
T 02
T02 = 308.70 K
0⋅ 4
( 1.275 ) 1.4 = T 01

METHODOLOGY ANALYSIS

REVIEW LITERATURE AND RESEARCH PAPERS.

IDENTIFY THE ACCOMPANY WITH THE JET
PROBLEM ENGINE

ACCQUIRE DATA
DEFINE RESEARCH GAP AND
IDENTIFY PROBLEM
STATEMENT.

SELECTION OF ENGINE FOR

COPTIMIZING
PERFORMANCE.

OBTAIN INFORMATIONOF
ENGINE COMPRESSOR.

ANALYSE

OPTIMIZE
 COMPARE

ANALYSE ENGINE PARAMETER

BASED ON VELOCITY TRIANGLE

OBTAIN ANALYTICAL SOLUTIONS OR

NUMERICAL SIMULATION

COMPARE WITH THE BASE PAPER FOR VALIDATION

METHODOLOGY DESIGN AND SIMULATION

SELECTION OF BLADE GEOMETRY DESIGN

SELECTION OF AIR, BLADE, MESH GENERATION
AND STAGGER ANGLE.

FLOW ANALYSIS DEFINE BOUNDARY LAYER

(INITIAL CONDITIONS)

OBTAIN RESULT AND VALIDATE.

REFERENCES
1) Govardhan M, Krishnakumar OG, Sitaram N (2007). Computational Study of
Effect of Sweep on the Performance and Flow Field in an Axial Flow
Compressor Rotor. I.MechE Part A: Journal of Power and Energy 221:315–
329.
2) Lakshminarayana B, Pouagare M, Davino R (1982). Three-Dimensional Flow
Field in the Tip Region of a Compressor Rotor Passage— Part 1: Mean
Velocity Profiles and Annulus Wall Boundary Layer. Transactions of ASME
104:760–771.
3) G. G. Adkins, Jr. L. H. Smith, Jr. Aircraft Engine Engineering Division General
Electric Company Cincinnati, Ohio 45215
4) S J Gallimore Rolls-Royce plc, PO Box 31, Derby DE24 8BJ, UK
5) By A. R. Howell, M.A.*, and R. P. Bonham, A.M.I.Mech.E.*
6) Avwunuketa, Ayedun Alex¹ , Adamu, Mohamed Lawal²,Ajao,Tofunmi
Ayodele³ Aircraft Engineering Department, Faculty of air Engineering, School
of Postgraduate studies, Ai rForce Institute of Technology, Kaduna Nigeria .
7) P. V. Ramakrishnaa & M. Govardhana a Thermal Turbomachines Laboratory,
Department of Mechanical Engineering, Indian Institute of Technology
Madras, Chennai 600 036, India Published online: 19 Nov 2014.
8) L. Porreca1 e-mail: luca.porreca@ch.manturbo.com A. I. Kalfas2 R. S. Abhari
Turbomachinery Laboratory, Swiss Federal Institute of Technology, CH-8092
Zurich, Switzerland
9) Deutsches Zentrum für Luft- und Raumfahrt e.V., Institut für
Antriebstechnik, Abteilung Turbulenzforschung, Müller-Breslau-Str. 8,
10623 Berlin, Germany
10) Salalah Ali and Joodakib aSchool of Mechanical Engineering, Arak
University of Technology, Arak, 38181-41167, Iran bUniversity of
Ayatollah Alozma Boroujerdi, Mechanical Engineering, Borujerd, Iran
11) W. JANSEN Senior Project Engineer, Northern Research and Engineering
Corporation, Cambridge, Mass. W. C. MOFFATT 1 Associate Professor of
Mechanical Engineering, Royal Militar y College of Canada, Kingston,
Ontario, Canada
12) M. inoue IVI. Kuroumaru M. Fukuhara Department of Mechanical
Engineering, Power Division, Kyushu University, Fukuoka, Japan