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7/25/2019 Tugas Besar Kelompok 13 - Supersonic Intake Design
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FINAL TASK REPORT
SUPER SONIC INTAKE DESIGN
AE3210 AERODYNAMICS 2
By:
GROUP 13
Bintang A.S.W.A.M [13613019]
Irvan Hakim F. [13613025]
Muhammad Fazlur R [13613039]
M. Rafiqi Sitompul [13613057]
Due Date : Wednesday, 18th May 2016
AERONAUTICS AND ASTRONAUTICS MAJOR
FACULTY OF MECHANICAL AND AEROSPACE
ENGINEERING
INSTITUTE TECHNOLOGY OF BANDUNG
2016
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Supersonic Intake Design
Department of Aerospace Engineering, Bandung Institute of Technology,
Indonesia
Abstract: This research’s goal is to perform aerodynamic design of an intake for
supersonic fighter aircraft. The intake is design to be have the most efficient
external-compression inlet. The analysis is taken at one flight condition that is
operate at 41000 ft. The air is assumed to be inviscid and calorically perfect. The
flow also assumed to be two dimensional and for the nominal design we will assume
the plane is flying level (angle of attack = 0). The calculations will be done
assuming gamma = 1.4. The Mach number after the normal shock at the entrance
to the subsonic diffuser is required to be M = 0.8 and the maximum angle that the
flow can turn prior to entering the diffuser is 25 degrees. The range of Mach
numbers to be considered from 3.1 ≤ ≤ 3.5.
Key words:
External-compression, inviscid, efficient
Introduction
Aircraft which is flying at supersonic speed or above the speed of sound has
a lot of problems for engineers. One of the problems faced is the design of Intake.
The Intake is a duct before the inlet of compressor. Because of the combustion
process occurs quite well at subsonic velocity air flow. Therefore, the incoming air
flow in Combustion Engines are required to become subsonic. To meet these
requirements, the flow must be slowed from supersonic to less than Mach 1.
In this project, the author assumes that ideally the flow enters the divergent
nozzle at the front of the engine intake should be about M = 0,8. The high number
Mach is reduced in divergent nozzle for the special needs of the machine. In this
project the author will design some design Intake for Mach numbers between M =
3,1 and M = 3,5. The final design is expected to be able to slow down the flow into
M = 0.8 and also have the highest Pressure Recovery.
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The aircraft engine requires a supply of uniform hogt total pressure recovery
air for good performance. It also related to total pressure loss which effects to the
engine thrust and consequently the fuel consumption. Thefore it is important to
maximaze the total pressure recovery at the inlet of engine. The total pressure
recovery is noted as ratio of the total pressure of the airflow before entering the
compressor to the freestream.
Proposed Method and Implementation
The preliminary design is divided into two following subpart. The first task is to
determine the number of ramps or oblique shocks and then select the subsonic
diffuser. The second task is to determine the total pressure recovery of the selected
design and choose the most efficient design.
A. Inlet Configuration and Properties
1. Inlet Preliminary Design
In preliminary design, we try to deciding the basic configuration of the inlet
such as theta angle. After that we calculated the Oblique Shock Wave (OSW)
properties for each stage interrelated so that in the end of the stage we have a normal
shock wave. The flight atmosphere condition at 41000 feet from ISA has static
pressure 17864.42 Pa, Temperature 216.65 K.
Assumption in calculation :
Adiabatic
No slip wall
No relative pressure along farfield and inlet
a. Using Manual Iteration Method
The preliminary design first and second Inlet (Inlet 1 & 2) is informed in
this table below.
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Table 1 inlet preliminary design 1
inlet
first
OSW
Secont
OSW
Third
OSW
Fourth
OSW
fifth
OSW
Normal
Shock
theta 10 10 10 10 8
beta 25.46 29.4 34.3885 41.31 49.9651
mach number 3.3 2.7468 2.2995 1.9077 1.5537 1.2692 0.8
Table 2 inlet preliminary design 2
Beta angle is calculated by doing iteration for known theta. The relation between
theta, beta and mach number for OSW is noted below
θ = flow deflection angle
β = oblique shock wave angle
γ = the ratio of specific heat
M1 = Mach number in fort of oblique shockwave
The geometry of oblique shock wave is demonstrated as the figure below
Figure 1. oblique shock wave geometry
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We have to know the Mach number behind the oblique shock wave (M2) for the
next stage oblique calculation. M2 can be determined by the formula :
Pressure static ratio between behind and in front of oblique is calculated by
To calculate the total pressure ratio, we need to know the ratio of temperature. With
isentropic assumption we can determine the total pressure. To get the temperature ratio
we have to calculate the density property of air flow.
Density ratio :
Total pressure ratio :
After all the properties known, Inlet Pressure Recovery (IPR),
Pi2 = Total pressure after the normal shock wave
Ptoo = Total pressure of free streeam
854919 1022180 0.8364
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b. Using Optimation Approaching Method
Geometry Approach for Failed Result
Consider for Normal Shocks:
Previous Stage: M j-1, P j-1,Pt j-1, T j-1
P j-1 = Pq; Pt j-1 = Ptq; T j-1 = Tq;
Notice that Mq = M j-1 represents upstream Mach Number , meanwhile M j
represents downstream Mach Number.
Governing Equation :
1− 22− 1 ;
Next Stage: M j, P j,Pt j, T j
Consider for Oblique Shocks:
Previous Stage: Mq-1, Pq-1, Ptq-1, Tq-1 ;
Governing Equation(s) :
+ − −(+)
[−−][−+]
tan
co− ++− ;ℎ 1,2,3, … … … , 1,
Optimum criteria for flow trajectory:
− sin() ≡ ;ℎ 1,2,3, … … … … . . . , 1,
For more convenience , let define
≡ 1−sin
1; ℎ 1,2,3, … … … … . . , 1,
2 + 1 1
+ 2 ;
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4 + 1
+ 1 ;
Thus,
− ;
By doing algebraic manipulation in order to obtain the results simultaneously and
also a little bit creativity to modify the governing equation, and then we’ll get
A = 19.66291154; B = 26.67264491; C= 14.32838557; (exists for
range freestream Mach Number between 1.7-5.4 )
Notice that n expresses the number of oblique shock(s) that occur.
Next Stage: Mq, Pq, Ptq, Tq ;
Equation of Total Pressure Ratio for oblique shock:
For k = 1-3,
[ +−+]
[ +−−]
Equation of Total Pressure Ratio for normal shock:
For j = k+1,
[ +−+]
[ +−−]
Thus, total pressure recovery (TPR) is
∏ × ∏
=+
=
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Table 2 inlet preliminary design 3
B.
Design of The Inlet
1. 2D Geometry Modelling
The geometry of inlet is sketched using CATIA based on preliminary design we
have known.We sketch 2D inlet after that extrute becoming 3D.
Draw all theta and beta angle using CATIA
Figure design 1
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Figure design 2
Obtained geometry length of the lines as follows
Figure design 1
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Figure design 2
2. 3D Modelling
Extrude 2D sketch
Figure design 1
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Create the Farfield, Inlet and Outlet
Figure design 1
Figure design 2
3. ANSYS Simulation
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Import geometry that has been created using CATIA to ANSYS in igs. Format
Figure design 1
Create and define sections/parts such as inlet, outlet, farfield and wall.
Figure design 1
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Figure design 2
Repair geometry
Figure design 1
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Figure design 2
Meshing Steps
Figure design 1
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Figure design 2
Figure design 1
Figure design 2
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Input ti CFX, then input Boundary Condition, Domain and Initial Condition.
Figure design 1
Input condition in Inlet, Outlet, Wall and Farfield
Figure design 1
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Figure design 1
Figure input condition design 2
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Figure input condition design 2
Figure input condition design 2
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Figure input condition design 2
4. Running the Simulation
Figure running design 1
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Figure running design 1
Figure running design 1
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Figure running design 2
Figure running design 2
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Figure running design 2
Figure running design 2
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C. The Inlet Design Result
Figure result design 1
Figure result design 1
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Figure result design 1
Figure result design 1
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Figure result design 1
Figure result design 2
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Figure result design 2
Figure result design 2
Total pressure at the end = 854919 pa
Mach number = 0.8
Comparison betweet analytical result and cfd total pressure = 1:1
-Mach 3.1 free slip wall (off design)
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-Mach 3.5 free slip wall(off design)
The result :
TP awal 1362560
M2 1.127735 sebelumnormal
Tp2 900350.6
mach akhir 0.8966
takhir/t2 1.078
Tp akhir 898730
p0akhir/p02 0.9982
IPR 0.659589
Table : result design 2
Figure result design 4
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Figure result design 4
Figure result design 4
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Failed Result
Figure Shock Detach
Figure Shock Detach
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Figure result design 3
Figure result design 3
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Figure result design 5
Figure result design 5
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Figure result design 5
Figure result design 5
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Analysis
To design an external air intake, as possible as Oblique Shock Wave hit
the cowl in order OSW not reflected in the internal intake. In the design, first sought
smallest theta and successively enlarged so that the total pressure loss can be
minimized. It happens because the total pressure before the first oblique so is high
so that the pressure loss is also high. In designing intake, as much as possible using
small angle because in addition to avoid the total pressure loss is high it can also
avoid detached-shock.
Before passing through the normal shock wave, we suggest not to
encourage area because it can thwart occurrence the normal shock wave. This case
happens if the rearing area before normal shock will accelerate the speed of the
flow.
To minimize the total pressure loss in normal, as much as possible the
flow rate approaching one. Because of the greater Mach number then normal shock
that occurs is getting stronger. To minimize the total pressure loss in the normal
shock wave, as much as possible about Mach Upstream Downstream Mach 1.1-1.3
order of about 0.81 - 0.94
Boundary layer effects caused by the walls of the intake was not overly
affect the angle beta and theta. This is evident from the results of CFD analysis has
been done, the result is the same as the analytical results that we did. As for the
boundary layer effects on speed, overall do not affect if the geometry is not too
large. However, the boundary layer remains influential on the speed around the
wall. To see a normal shock wave, is better viewed with the pressure difference.
Because if you use the speed difference, there will be the effect of the boundary
layer on the wall. So, there is the velocity distribution in the direction perpendicular
wall.
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Conclusion
The design of Inlet 1 has great Inlet Pressure Recovery which is 0.8364 as
same as design 2. But, the design 1 is diffrent kind with design 2. Design 1 is two-
dimentional sigle duct, in the other hand Design 2 is two-dimentional bifurcated
duct.
Actually we have 4 design as we can see in the result of this report. But, 2
kind design has been failed so far, the problem is, first theta is to high so that it
occure shock detach. And the last design is fail because of technical problem using
ANSYS, we guess that for the Axisymmetric outward-turning has different
threarment in the meshing and define the boundary condition.
References
Anderson, John D. 1990. Modern Compressible Flow. Second Edition.
Singapore : McGraw-Hill Publishing Company
Ran, Hongjun and Dimitri Mavris. 2015. Paper : Preliminary Design of a
2D Supersonic Inlet to Maximize Total Pressure Recovery. Atlanta : Georgia
Institute of Technology