Last Year`s Sample Report

Transcription

A Conceptual Design
Approach
to a turbofan engine
Advisor: Assis. Prof. Dr. Onur Tuncer
Team Kuzgun- TurkJet-1
Burak Özkahya
Cihat Akın
Coşku Çatori
2
3
ABSTRACT:
The aim of this project is to design a conceptual engine to be fitted into a half-sized
model of Lockheed Martin F-35 Lightning II multirole fighter UAV.
TJ-1 (TurkJet-1) provides the necessary high power extraction emerged by the
growing technology of UAV’s. In order to minimize the cost and dimensions of the A/C,
smallest possible engine dimensions have been investigated. To increase the power, spools
have been designed to rotate coaxially. Therefore, power generation unit is moved to the inner
part of the engine, removing the problematic mechanic power generation. TJ-1, in this
manner, is a baseline engine for the current demand of all-electrical A/C’s. In addition, for
engine inlet, two symmetrical inlet and s ducts has been designed to ensure stealth,
considering the military aspect of the JSV. TJ-1 also includes a two dimensional variable
nozzle, thus increasing the maneuverability of the A/C and reducing the fuel consumption
during cruise. Keeping mind of current technological limits, each spools are supplied with a
single-stage turbine.
Special Thanks to:
Assoc. Dr. Onur Tuncer
Aykut Dağkıran
Ufuk Inan Bayrı
Oğuz Eren
4
CONTENT:
1
Introduction ..................................................................................................................................... 7
2
Cycle Analysis............................................................................................................................... 16
3
Inlet Design ................................................................................................................................... 41
4
Combustion Systems ..................................................................................................................... 45
5
Compressor.................................................................................................................................... 50
6
Turbine .......................................................................................................................................... 63
7
Nozzle Design ............................................................................................................................... 72
8
Electricity ...................................................................................................................................... 76
9
Appendix ....................................................................................................................................... 81
LIST of TABLES
Table 1-1: Aircraft minimum net thrust requirements .............................................................. 8
Table 1-2: Similar engine specifications ................................................................................... 9
Table 1-3: Projected design space ........................................................................................... 10
Table 1-4: Aircraft performance requirements ........................................................................ 11
Table 1-5: Primary mission profile ......................................................................................... 13
Table 1-6: Aircraft specifications ............................................................................................ 15
Table 2-1: Engine design variables ......................................................................................... 17
Table 2-2: Sensitivity analysis ................................................................................................ 29
Table 2-3: Engine specifications ............................................................................................. 36
Table 2-4: Engine thrust and TSFC values ............................................................................. 37
Table 2-5: Fuel save (lbf) ........................................................................................................ 38
Table 3-1: Ramp angles ........................................................................................................... 43
Table 3-2: Duct specifications ................................................................................................. 44
Table 4-1: Stations dimensions ............................................................................................... 45
Table 4-2: Air partitions (Tg= 4174 0Ra, εPZ = 0.8)................................................................. 46
Table 4-3: Zones geometry...................................................................................................... 47
Table 4-4: Burner geometry .................................................................................................... 47
Table 4-5: Flow areas .............................................................................................................. 47
Table 4-6: Mixer +diffuser dimensions ................................................................................... 48
Table 4-7: Combustion parameters ......................................................................................... 49
Table 5-1: Fan input values ..................................................................................................... 54
Table 5-2: Fan output value..................................................................................................... 55
Table 5-3: HPC input values ................................................................................................... 57
Table 5-4: HPC output values of 1-3. Stages .......................................................................... 58
Table 5-7: Fan centrifugal stress ............................................................................................. 61
Table 5-8: HPC centrifugal stress ........................................................................................... 61
Table 5-9: Greek Ascoloy properties ...................................................................................... 62
Table 6-1: HPT design point parameters (1.1 M/20kft) .......................................................... 65
Table 6-2: Conditions of cold air rig test of single-stage turbine at NASA report[12] ............. 65
Table 6-3: Performance parameter values ............................................................................... 66
5
Table 6-4: Results of the TURBN.exe for HPT -1- ................................................................ 67
Table 6-7: Low-pressure turbine design point parameters (1.1M/20kft) ................................ 70
Table 6-8: Results of the TURBN.exe for LPT -1- ................................................................. 70
Table 6-9: Results of the TURBN.exe for LPT -2- ................................................................. 70
Table 6-10: Results of the TURBN.exe for LPT -3- ............................................................... 70
Table 7-1: Input values of nozzle calculations ........................................................................ 75
Table 7-2: Output values of nozzle calculations ..................................................................... 75
LIST of FIGURES, GRAPHICS and DRAWINGS
Figure 1-1 : Constraint analysis results ................................................................................... 12
Figure 1-2 : Reference weight fractions and Fuel used through mission ................................ 14
Figure 2-1 : A8/A9 ratio at dry conf. ........................................................................................ 18
Figure 2-2 : A8/A9 ratio at wet conf. ....................................................................................... 19
Figure 2-3 : Fan/LPC dry configuration .................................................................................. 20
Figure 2-4 : Fan/LPC wet configuration ................................................................................. 20
Figure 2-5 : Mixer pressure ratio dry conf. ............................................................................. 21
Figure 2-6 : Mixer pressure ratio wet conf. ............................................................................. 22
Figure 2-7 : Overall pressure ratio dry conf. ........................................................................... 23
Figure 2-8 : Overall pressure ratio wet conf. .......................................................................... 23
Figure 2-9 : Design bypass ratio dry conf. .............................................................................. 24
Figure 2-10 : Design bypass ratio wet conf. ........................................................................... 25
Figure 2-11 : Design Mixer Mach number dry conf. .............................................................. 26
Figure 2-12 : Design Mixer Mach number wet conf............................................................... 26
Figure 2-13 : Maximum burner exit temperature dry conf. .................................................... 27
Figure 2-14 : Maximum burner exit temperature wet conf. .................................................... 28
Figure 2-15 : Maximum reheat exit temperature dry conf. ..................................................... 28
Figure 2-16 : Total pressure ratio in flight .............................................................................. 30
Figure 2-17 : Specific thrust in flight (Mcrit = 0, altdesign=0 ft, θbreak=1) .................................. 31
Figure 2-18 : Specific thrust in flight (Mcrit = 1.157, altdesign=25,000 ft, θbreak=1.05) ............. 31
Figure 2-19 : TSFC in flight (Mcrit = 0, altdesign = 0 ft, θbreak = 1) ............................................ 32
Figure 2-20 : TSFC in flight (Mcrit = 1.157 altdesign = 25,000 ft, θbreak = 1.05) ........................ 32
Figure 2-21 : Mission Fuel saving percentages (contour lines) .............................................. 34
Figure 2-22: Engine Airflow requirements “contour lines for flow rate (lbm/s)” .................. 35
Figure 2-23 : Super cruise fuel consumption .......................................................................... 36
Figure 2-24 : TJ-1 weight fractions and Fuel used through mission ....................................... 38
Figure 2-25 : Power take-off variation at 0.9 M 35 kft ........................................................... 39
Figure 2-26 : Power extraction 0.9 M 35 kft ........................................................................... 40
Figure 2-27 : Power extraction 1.4 M 35 kft ........................................................................... 40
Figure 3-1 : Engine airflow requirement in flight ................................................................... 42
Figure 3-2 : Inlet geometry ..................................................................................................... 43
Figure 3-3 : Inlet pressure recovery ratio ................................................................................ 44
Figure 4-1 : Swirler layout ...................................................................................................... 46
Figure 4-2 : Mixer + diffuser layout ....................................................................................... 48
Figure 4-3 : Flameholders layout ............................................................................................ 48
Figure 6-1 : Turbine stages and velocity diagrams ................................................................. 64
Figure 6-2 : Smith chart for turbine stage efficiency[9] ........................................................... 68
Figure 6-3 : Creep Rupture Life (h) – Stress (MPa) graph of SC16[10]................................... 69
Figure 8-1 : Integrated started generated system .................................................................... 79
Figure 8-2 : Starting process ................................................................................................... 79
6
Acronyms
LPC: Low pressure compressor
HPC: High pressure compressor
LPT: Low pressure turbine
HPT: High pressure turbine
AB: Afterburner
β:
bleed air fraction
PTOL: Low spool power takeoff
PTOH: High spool power takeoff
ε:
Cooling air fractions
ef:
fan polytrophic efficiency
ecL: LPC polytrophic efficiency
ecH: HPC polytrophic efficiency
etH: HPT polytrophic efficiency
etL:
LPT polytrophic efficiency
ηburner: burner efficiency
ηAB: afterburner efficiency
ηmL: Low spool mechanical efficiency
ηmH: High spool mechanical efficiency
πcid: Compressor inter duct pressure loss
πbyid: Bypass inter duct pressure loss
πtid: Turbine inter duct pressure loss
πted: Turbine exit duct pressure loss
πm5: Mixer core duct pressure loss
πm1.6: Mixer bypass duct pressure loss
πintake: Intake diffuser inter duct pressure loss
πM max: Mixer inter duct pressure loss
πAB: Afterburner inter duct pressure loss
πnozzle: Nozzle inter duct pressure loss
πburner: Burner inter duct pressure loss
πf :
Fan pressure ratio
πcL: Low compressor pressure ratio
πcH: High compressor pressure ratio
α:
Bypass ratio
A8/A9: Nozzle throat ratio
Tt4: Burner exit temperature
Tt7: Afterburner exit temperature
7
1
Introduction
One of the main ongoing researches for military systems is UAV’s. UAV’s remove the
human limitations, human factor and offer more precise results. However, current technology
requires different solutions to onboard electric generation. TJ-1 is designed to offer a possible
solution to this problem.
Design process started with competitors’ study. Using competitors’ values with
accessory computer programs, necessary and desired engine parameters have been acquired.
Nozzle, inlet, compressor, burner(s) and turbine calculations have been conducted with
respect to each other. Investigating current technological trends, electrical system of the
engine has been formed. All designed engine parts are drawn with a CAD program and CAD
drawings are given in Appendix -E-. Detailed information will be given later on throughout
the report.
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1.1
Request for Proposal Requirements
As a part of Joint AIAA-IGTI Undergraduate Team Engine Design Competition
(UTEDC) 2011/2012 a Request for Proposal (RFP) paper has been published at August 24,
2011. RFP focused on supplying the required onboard power to A/C’s, emerged by the
growing power requirement for A/C systems, while having a minimum influence on total
performance.
RFP calls for a low bypass augmented turbofan engine concept which is powered to a
half scaled model of a Joint Strike Fighter (JSF) UAV. A/C’s envisaged flight regime varies
from static sea level condition to supersonic flight at 40,000 feet and 1.8 Mach. RFP provides
a table which contains minimum net thrust requirements for different flight conditions. While
delivering requested thrust, full augmentation usage is possible for takeoff condition even
though partial afterburner usage is encouraged to achieve the required thrust values stated in
Table (1-1). Maximum allowable temperature at after burner is 3200R.
Table 1-1: Aircraft minimum net thrust requirements
Designed system should deliver 67 hp auxiliary power under all flight conditions for
general purposes. Additional 300 hp power will be required throughout the mission, except
take-off, during combat maneuvers, which can occur at both subsonic and supersonic
segments, for avionics and weaponry. That secondary power need can be achieved via either
HP, LP spool or some combination. As a secondary objective, maximum power extractable
and how this is split between spools without compromising aircraft performance must be
determined in two mission conditions which are 0,9 Mach at 35,000 feet and 1,4 Mach at
35,000 feet.
Design progress should carry out two additional goals. One of them is aerodynamic
similarity, which must be maintained with respect to the baseline engine model, F100-PW229, by controlling RPM numbers of the spools and secondly fuel consumption should be
minimized considering cost & logistics of the military operation.
9
1.2
Design Drives
High and efficient power extraction while providing the required thrust is the main
intent of the concept. Some insights about major design drives can be obtained after a
thorough examination of the RFP. As stated, powered unit will be a half size model of the
JSF. Therefore, engine dimensions must be very carefully determined to ensure engine fits in.
Peak points of net thrust demand are; sea level static condition(take-off) and 1.8 Mach at
40,000ft.These peak points seems reasonable compared to the similar aircraft’s but the
temperature limit of 3200R is lower than the current available engine technology limits. As it
has been stated in Walsh[1] augmentation thrust gain ratio respect to dry thrust higher at higher
Mach numbers and augmentation limit given could stress the design.
1.3
Technological Stand Point
First step in the design was to form a database using existing concepts in order to grasp
the technological and economical limits of the requested design. Studying F135-PW-100,
F125-GA-100, EJ200 Mk.100 engines with addition to baseline engine, reference points has
been provide, which will be used later on. General specifications of the engines used in the
database are given in Table 1-2.
Table 1-2: Similar engine specifications
F100-PW-229A
EJ200 Mk.100
F135-PW-100
F125-GA-100
DRY
WET
DRY
WET
DRY
WET
DRY
WET
Thrust (lbf)
17,800
29,000
13,500
20,000
28,000
43,000
6,250
9,250
TS FC (1/h)
0.785
1.667
0.726
1.91
0.7
1.95
0.785
2.06
Airflow (lbm/s)
254
163
200
92.6
OPR
32
26
28
21
Bypass Ratio
0.36
0.4
0.57
0.49
Compressors
3L,10H
3L,5H
3L,6H
3L,4H+1C
Turbines
2H,2L
1H,1L
1H,1L
1H,1L
Diameter (inc)
46.5
-
51
23.3
Length(inc)
191.2
157
220
140.2
Weight (lbm)
3740
2180
3750
1360
10
Envelope limitations are not given in RFP so the definition of a reasonable design space
and size limits are obtained by scaling referenced engines to the thrust values, required in
Table 1-1, using Equation (through 1 to 3) as stated in Raymer[3]. Scaled engines maximum
and minimum values have been taken as the design space’s upper and lower limits and given
in Table 1-3.
(
)
( )
(
)
( )
(
)
( )
Table 1-3: Projected design space
TSFC (1/h)
Airflow (lbm/s)
OPR
Bypass Ratio
Compressors
Turbines
Diameter (in.)
Length(in.)
Weight (lbm)
1.4
Lower
DRY
WET
0.7
1.667
42.7
21
0.36
3L,5H
1H,1L
23.75
117.06
721
Upper
DRY
WET
0.785
2.06
90.1
32
0.57
3L,10H
2H,2L
26.75
142.3
1417
Aircraft
Generally design of an aircraft engine is a process which takes place simultaneously
with A/C design if not later. Great amount of A/C data is required to design a competent
engine.
Absence of the A/C data and mission profile in RFP, a common conceptual fighter
mission profile is taken in the design study. Constraint boundary analysis’s performance
requirements and mission profile definition at AIAA Engine Design [2] taken as a baseline and
modified considering aircraft specifications.
11
Table 1-4: Aircraft performance requirements
Take-off
2000ft PA, 100oF, STO = SG +SR ≤ 1500 ft
Acceleration
kTO =1.2, µTO = 0.05, max power
Rotation
VTO, tR = 3 s, max power
Supersonic penetration & Escape dash
1.4 M / 35 kft, no afterburning
Combat
35,000 ft
Acceleration
Landing
0.81.6 M, t ≤ 50 s, max power
2000 ft PA, 100oF, SL = SFR +SBR ≤ 1500 ft
Free roll
kTO =1.15, tFR = 3 s, µTO = 0.18
Breaking
Drag chute diameter 3,79 ft , deployment ≤ 2.5 s
Maximum Mach number
1.8 M/ 40 kft, max power
One of the main drives of ongoing development increment on UAV trend is to make A/C’s
unbound to human limitations which has a heavy impact on aircraft endurance and allows
higher G maneuver capabilities. Referring to this, increasing aircraft’s sustainable value at
cruise altitude, 35,000ft, to 7-9 G is considered at first. However, detailed investigation
showed that Air Combat Maneuvering (ACM) has diminishing usage in air warfare due to
improvements in Beyond-Visual-Range (BVR) arms. Having said that, it is seen that
increasing thrust to weight ratio (T/W) with fixed-thrust just to sustain higher maneuver
ability is not a viable trade off. Therefore sustainable g-turn constrains does not increase
beyond what can sufficient to accomplish other requirements. Thus T/W and Wing Loading
(W/A) have been chosen accordingly with design space as; 1.05 and 68 lbf/ft2. Table 1-4
shows the definition of the requirements. Results of the constraint boundary analysis are given
in Figure 1-1.
12
1,40
1,30
1,20
Typhoon
F-15
1,10
Thrust Loading
Refale
Design Point
F-16
F-22
Takeoff
Landing
1,00
F-18
0,90
F-35
Mirrage 2000
Aceleration
Max Mach
Supercruise
0,80
1.6M/35Kft--6 G
0.9M/35kft--3.7 G
0,70
Design Point
REAL
0,60
0,50
0,40
40,00
50,00
60,00
70,00
Wing Loading
80,00
90,00
(lbf/ft2)
Figure 1-1 : Constraint analysis results
Two different sets of coefficients are needed for calculations; one to define
aerodynamic characteristics and the other for engine performance characteristics.
Aerodynamic coefficients are found via AIAA Aircraft Engine Design[2].
Engine performance characteristics, “α” -available thrust or thrust lapse- and Specific
Thrust Fuel Consumption (TSFC) can be expressed via simple algebraic models by using nondimensional temperature (θ and θo), non-dimensional pressure (δ and δo) and throttle ratio
(TR). Coefficients of these models vary from engine to engine and standard values have been
given[2]. Along with these values, new coefficients have been derived for each engine
evaluated in off-design cycle analysis. Reference equations are given below:
Reference Engine
Maximum power:
{
(
(
)√
)
}
13
Military power:
{
(
(
)
}
)√
In the light of the constrain boundary analysis and setting the range similar to JSF[4],
which is 584 nmi combat radius, mission profile is constructed. Primary mission profile is
given in Table 1-5
Table 1-5: Primary mission profile
1-2 Warm-up and takeoff
A--Warm-up
B--Acceleration
C--Rotation
2-3 Acceleration and climb
D--Acceleration
E--Climb/acceleration
3-4 Subsonic cruise climb
4-5 Descend
5-6 Combat air patrol
6-7 Supersonic penetration
F--Acceleration
G--Penetration
7-8 Combat
H--Fire AMRAAMs
I--Tum 1
J--Turn 2
K--Acceleration
L--Fire AIM-9Ls & ½
ammo
8-9 Escape dash
9-10 Minimum time climb
10-11 Subsonic cruise
climb
11-12 Descend
12-13 Loiter
13-14 Descend and land
2000 ft PA, 100°F
60 s, mil power
kto = 1.2, µto = 0.05, max power
Mto, tR ---- 3 s, max power
Minimum time-to-climb path
Mto to McL/2000 ft PA, 100°F, mil power
Mcc/2000 ft PA, 100°F to BCM/BCA,mil power
BCM/BCA, As23 + As34 = 375 n miles
BCM/BCA to Mcae/35k ft
35 kft, 20 min
35 kft
McaP to 1.4M/35k ft, max power
1.4M, ASF + Asc = 125 n miles,no afterburning
35 kft
652 lbf
1.5M, one 360 deg Max g sustained turn,with afterburning
0.9M, two 360 deg Ma x g sustained turn,with afterburning
0.8 to 1.6M, At< 50 s, max power
657 lbf
1.4M/35 kft, As89 = 75 n miles,no afterburning
1.4M/35 kft to BCM/BCA
BCM/BCA, As10-11 = 200 n miles
BCM/BCA to Mloiter 10 kft
Mloiter 10 kft, 20 min
Mloiter 10 kft to 2000 ft PA, 100°F
14
Each and every engine designed is simulated using the mission profile above. Using
the simulation results was an important factor in selecting the final design. Mission
performance with weight fractions for the reference design is given in Figure 1-2. Takeoff
gross weight used in mission has been estimated as 9143 lbf by using T/W ratio calculated
before.
1,0
700
0,9
500
0,8
400
300
0,7
200
0,6
Landing
Loiter
Descend
Subsonic cruise
Climb
Escape dash
Acceleration
Fire AIM-9Ls
0.9M turn
0
1.6M turn
Supercruise
Acceleration
Air patrol
Descend
Subsonic cruise
Climb
Acceleration
Rotation
Acceleration
Taxi/Warm-up
0,5
Fire AMRAAMs
100
Figure 1-2 : Reference weight fractions and Fuel used through mission
From the Weight fraction it can easily be seen that subsonic and supercruise has the
heaviest impact on the total fuel consumption. Doing constrain boundary and mission analysis
main parameters for the conceptual aircraft are gathered and can be seen in Figure 1-2.
Fuel used (lbf) -- (red)
Total weigth fractions -- (blue)
600
15
Table 1-6: Aircraft specifications
Aircraft's Spesifications
Takeoff Gross Weight
9143 l bf
Thrust S ea Level
9600 l bf
Wing Area
134.4 ft2
Payload
2075 l bf
Empty Weight
4464 l bf
Fuel Weight
2064 l bf
Thrust to Weight Ratio
1.05
Wing Loading
68 l bf/ft2
Inspection of the gathered results shows that empty weight fraction of the plane is
48.83%. Which may seem small for a fighter thus another equation is used for comparison;
(4)
Equation 4[16] has been calculated as 44.36% which is even a smaller value. Therefore,
calculated weight is found reasonable in this respect.
16
2
Cycle Analysis
2.1
Introduction
Turbo-machinery cycle analysis consists of two different calculation phase. First phase
is called “parametric cycle analysis” which is developed in order to understand performance
characteristic of the engine with respect to its design point. In the scope of parametric cycle,
engine dimensions and physical specifications (i.e. Overall pressure ratio, maximum inlet
temperature, bypass ratio) have not yet been determined. Variation of these specifications is
determined by the designer’s will while each combination defines a different engine.
Parametric cycle only determines engine performance under design conditions and it is useful
to determine the limitations of the design space. It is also the prerequisite of the off-design
calculations. Off-design cycle analyses are conducted in the means of acquiring the
performance characteristic of a physically defined engine in various flight conditions.
17
Complete definition of the two spool low-bypass mixed turbofan engine with
afterburner concluded only after determining 36 engine quantities, which are categorized in
Table 2-1.
Table 2-1: Engine design variables
Aircraft system parameter
Design limitations
Coolant fractions
Polytrophic efficiencies
Component efficiencies
Total pressure losses
Design choices
̇
⁄
Determining that amount of quantities could be tedious. Furthermore if not chosen
appropriately, nonrealistic results may arise. Accurate usage of historical data and trends
reduces the number of unknown design parameters which is essential given the massive
number of unknowns.
Every single quantity under the group off design limitation defined to indicate figure
of merit of the subparts engine consist off and these values present the current technological
limitations. Efficiencies hold same as baseline engine at the early stages of the study and
revised along the process when a better suited values secured.
Aircraft systems parameters on the other hand are selected considering aircraft system
requirement. Generally two of the aircraft subsystems, pressurizing and anti-icing systems,
need bleed air to be operative. UAV does not require pressurization and selection of an
electrical anti-icing system is found more convenient taking note of the fact that bleed airless
engines are more efficient. While extracting power is the primary object of the design,
conducting the exploration of the promising cycle boundaries under 67 horsepower (hp)
continuous power load is found reasonable. Lastly, cooling air percentages have been taken
same with baseline values until real values are calculated at turbine design.
After these assumptions 10 design choice left out to determine. Before determination
that quantities some understanding of turbomachinery has to be applied with further subgrouping and ordering to explore that 10 dimensional design space in a reasonable manner.
18
All of the following graphs gathered by using base line engine at 25000 ft, 1.1566 M, 1.05
θbreak for referencing.
2.2 Pressure Group
2.2.1
Throat Area Ratio
Throat area is the definition of how much of the pressure formed in throat turns to
kinetic energy via expansion. Meaning, instead of using the area ratio, pressure at the nozzle
exit and ambient pressure can be used to express throat area.
Inlet Corr. Flow W2Rstd = 300 ... 0 [lb/s]
Input Parameter 4 = 0,8 ... 1,25
.908
0,8
.907
.906
Sp. Fuel Consumption [lb/(lb*h)]
1,25
.905
0,85
1,2
.904
1,15
.903
1,1
.902
.901
0,95
1,05
1244
1246
1248
1250
1
1252
1254
1256
Specific Thrust [ft/s]
28.03.2012
GasTurb 11
Figure 2-1 : A8/A9 ratio at dry conf.
In Figure 2-1 line approaching optimum and line returning from optimum
coincidences though it may not be visible at first. This means that there is a point for pressure
ratios to achieve minimum fuel consumption and maximum thrust and that pressure ratio is
actually analytically defining ‘perfectly expansion’ assumption for the nozzle.
19
.
1.884
0,8
1.882
1.88
1,25
1.878
0,85
1.876
1,2
1,15
1.874
1,1
0,95
1.872
1,05 1
1.87
2840
2845
2850
2855
28.03.2012
GasTurb 11
Figure 2-2 : A8/A9 ratio at wet conf.
Same situation is also valid when afterburner is on and same optimum value exists.
Referring to these facts and setting pressure ratio to ‘1’, throat ratio can be found and is no
longer a variable.
2.2.2
Inner Fan
Remaining design parameters can actually be reduced by using the dependency of the
parameters between each other. First of all, fan compression and low pressure compression
ratios cannot be evaluated separately because they are being carried out from different regions
of the same rotor. It can be assumed that Fan/LPC compression ratios will differ between
‘1.05’ and ‘1.25’.
20
Input Parameter 1 = 1 ... 1,28
.9032
.9028
1
.9024
1,04
.902
1,08
.9016
1,12
.9012
1,16
.9008
1,2
.9004
1,24
.9
.8996
1,28
1246
1250
1254
1258
1262
26.03.2012
GasTurb 11
Figure 2-3 : Fan/LPC dry configuration
As it can already be seen in Figure 2-3, increasing fan inner/outer compression ratio
increases thrust and decreases fuel consumption. However, it should be noted that this
increment is minimal. In addition, aerodynamically it is a real challenge to increase this ratio
further.
Input Parameter 1 = 1 ... 1,28
1.872
1.87
1
1.868
1,04
1,08
1.866
1,12
1.864
1,16
1.862
1,2
1.86
1,24
1.858
1.856
1,28
2860
2865
2870
2875
2880
26.03.2012
GasTurb 11
Figure 2-4 : Fan/LPC wet configuration
21
From Figure 2-4, it can be seen that situation remains the same while afterburner is on.
In this case, bypass/core pressure ratio should be raised as high as possible. Therefore,
bypass/core pressure ratio is chosen as ‘1.2’ –same as baseline engine- for initial study.
2.2.3
Outer Fan
Designed engine has a mixed-exhaust configuration. Therefore, bypass and core streams
cannot be investigated separately because two streams will mix later on by mixer.
.928
.924
0,6
1,32
.92
1%
.916
1,24
0,68
.912
.908
1,16
0,76
.904
0,84
1,08
0,92
.9
.896
1210
1220
1230
1240
1250
1
1260
1270
26.03.2012
GasTurb 11
Figure 2-5 : Mixer pressure ratio dry conf.
As it can be seen in Figure 2-5, pressure ratios have an optimum value around ‘1’.
Value of pressure ratio for dry configuration is not very important considering the little
change in TSFC and thrust with pressure ratio.
22
1.93
1.92
0,68
1.91
1.9
1,32
1%
0,76
1.89
1,24
0,84
1.88
1,16
1.87
0,92
1,08
1.86
1.85
2780
2800
2820
2840
1
2860
2880
26.03.2012
GasTurb 11
Figure 2-6 : Mixer pressure ratio wet conf.
Same trend can also be seen when Afterburner is opened in Figure 2-6. In this case
optimum value is around ‘1’ but changing the pressure with afterburner configuration has a
significant effect on TSFC and thrust. By choosing fan compression ratio as ‘1’ for each
engine design will remove this ratio as a variable.
2.2.4
High Pressures Compressor
Only remaining variable in pressure group remains high pressure compressor
compression ratio. Choosing total pressure ratio instead of HPC pressure ratio as a design
variable has great advantages considering evaluation and comparison. Total pressure ratio is
limited by the temperature values at the last stages of HPC and it is around 35-45. However, it
should be noted that increasing total compression means increment in length, weight and
number of stages.
23
Input Parameter 2 = 18 ... 36
.97
18
.96
20
.95
.94
22
1%
.93
24
26
.92
28
.91
30
32
.9
34
36
.89
.88
1200
1250
1300
1350
1400
26.03.2012
GasTurb 11
Figure 2-7 : Overall pressure ratio dry conf.
From Figure 2-7, it can be seen that increment in total compression while afterburner is
off, reduces TSFC and specific thrust. Nonetheless, increment in total pressure has a
diminishing return. As the pressure rises, the percentage loss of specific thrust per TSFC rises.
Input Parameter 2 = 18 ... 36
1.89
1.88
36
1.87
34
32
1.86
1%
30
1.85
28
26
1.84
24
22
1.83
1.82
18 20
2840
2860
2880
2900
2920
26.03.2012
GasTurb 11
Figure 2-8 : Overall pressure ratio wet conf.
24
While the afterburner is on, unlike the dry condition, increment in total compression
results in decrement in thrust and fuel efficiency. In Figure 2-8, it is seen that there is an
optimum pressure ratio. However, it is around ‘20’ so this value is over feasible limitations
considering the significant effect it will have on fuel consumption while the afterburner is off.
2.3 Bypass Group
2.3.1
Design bypass ratio
Bypass ratio is the ratio between the mass flow rate of the stream going into core and
bypassing it. Low bypass engines are a hybrid of jet and high bypass engines and bypass ratio
shows which side it is closer to. As the bypass ratio increases, engine increases it is fuel
efficiency by getting close to high bypass. Decreasing it increases the thrust efficiency.
Therefore, bypass ratio should be chosen very carefully.
Design Bypass Ratio = 0,2 ... 0,76
1
0,2
.99
.98
0,28
.97
0,36
1%
.96
0,44
.95
0,52
.94
0,6
.93
0,68
.92
0,76
.91
1300
1400
1500
1600
1700
1800
1900
26.03.2012
GasTurb 11
Figure 2-9 : Design bypass ratio dry conf.
As it has been very mentioned, while the engine is in non-augmented mode, increment
in bypass ratio leads to better fuel consumption, worse specific thrust. ‘1’ unit percentage loss
in specific thrust equals to ‘3’ times reduction in fuel consumption. Therefore increasing the
bypass ratio is very tempting.
25
Design Bypass Ratio = 0,2 ... 0,76
1.84
0,76
1.82
0,68
1.8
0,6
1.78
1%
0,52
1.76
0,44
1.74
0,36
1.72
0,28
1.7
0,2
1.68
1.66
2880
2960
3040
3120
3200
26.03.2012
GasTurb 11
Figure 2-10 : Design bypass ratio wet conf.
While the afterburner is on, situation changes a bit. Increment in bypass ratio negatively
effects both fuel consumption and thrust generation. Ratio of percentage rise in fuel
consumption and decline in specific thrust equals ‘0.9’. This would be a negative result of
choosing a high bypass ratio. These two opposite trend, as it has been mentioned above,
makes choosing a bypass ratio more and more significant and difficult to choose.
2.3.2
Design Mixer Mach Number
Changing the design mixer Mach number is only possible under certain limitations. As
mixer Mach number plays an important role in establishing the Mach number of the stream
coming from the bypass and core. If the mixer Mach number goes upon a certain limit, bypass
Mach number will exceed 1 and enter supersonic region. This would lead to unwanted results
in engine. Also, afterburner limits the design mixer Mach number further. It would be
reasonable to limit the Design mixer Mach number between ‘0.1’ and ‘0.3’.
26
Design Mixer Mach Number = 0,1 ... 0,66
.9032
.9028
0,62
.9024
0,54
.902
.9016
0,46
.9012
0,38
0,3
.9008
0,22
0,14
.9004
1255
1256
1257
1258
1259
26.03.2012
GasTurb 11
Figure 2-11 : Design Mixer Mach number dry conf.
Figure 2-11 is generated to show the effects of mixer Mach number on the engine are
insignificant, in dry mode, even if it reaches as high as ‘0.6’. Graphic also shows how the
decline in mixer Mach number effects fuel consumption and thrust positively for the dry
condition.
Design Mixer Mach Number = 0,1 ... 0,66
1.92
1.9
0,3
1.88
1%
0,26
1.86
0,22
1.84
0,18
0,14
1.82
0,1
1.8
2800
2820
2840
2860
2880
2900
2920
2940
2960
26.03.2012
GasTurb 11
Figure 2-12 : Design Mixer Mach number wet conf.
27
Figure 2-12 shows that decline in mixer Mach number, with afterburner configuration,
results positively for the engine. This result supports the decision of choosing mixer Mach
number as lows as possible. However, lowering the mixer Mach number increases Mixer area,
in other words engine diameter. Also mixer Mach number will differ from off-design
conditions. In the light of all these facts and considering all operating stages of the engine,
mixer Mach number is envisaged ‘0.2’.
2.4 Temperatures
2.4.1
Maximum Burner Exit Temperature
Burner Exit Temperature = 2600 ... 3500 [R]
Inlet Corr. Flow W2Rstd = 0 ... 661,387 [lb/s]
1
.98
3500
1%
3400
.96
3300
.94
3200
3100
.92
3000
2900
.9
2800
2600
.88
800
2700
1000
1200
1400
1600
1800
2000
26.03.2012
GasTurb 11
Figure 2-13 : Maximum burner exit temperature dry conf.
Investigating the effects of maximum inlet temperature on engine (see Figure 2-13),
increasing the temperature while the afterburner is on results in better specific thrust and
worse fuel consumption. Even if higher fuel consumption is undesired, 5% rise in thrust for
1% increase in fuel consumption is beneficial.
28
Burner Exit Temperature = 2600 ... 3500 [R]
2.1
2600
2
2700
1%
1.9
2800
2900
1.8
3000
3100
3200
1.7
3300
3400
3500
1.6
1.5
2600
2800
3000
3200
3400
26.03.2012
GasTurb 11
Figure 2-14 : Maximum burner exit temperature wet conf.
In Figure 2-14, afterburner configuration, increasing the temperature results in better
fuel consumption and better thrust. Therefore, maximum inlet temperature is taken as high as
material allows which is 3000R.
2.4.2
Maximum reheat temperature
Reheat Exit Temperature = 2600 ... 3500 [R]
2.1
2
3500
3400
1%
1.9
3300
3200
3100
1.8
3000
2900
1.7
2800
2700
1.6
1.5
2600
2400
2600
2800
3000
3200
26.03.2012
GasTurb 11
Figure 2-15 : Maximum reheat exit temperature dry conf.
29
Maximum inlet temperature for afterburner naturally only effects while the afterburner
is on. Figure 2-15 shows that increment in temperature, increases both thrust and TSFC
equally. Increment in temperature with dry configuration has no harmful effects on but
engine dimensions are limited by military thrust values, meaning engine dimensions are
affected by the increase in temperature. Decreasing the engine dimensions reduces weight and
cost. Therefore, it is very desirable. Considering all that maximum inlet temperature for the
afterburner is chosen as 3200R as limited by RFP. In addition, sensitivity analysis is
conducted and the results are given in Table 2-2.
Table 2-2: Sensitivity analysis
Sensivity Analysis
Range
TSFC (lb/(lb*h))
Dry
Specific Thrust (ft/s)
Afterburner
TSFC (lb/(lb*h))
Specific Thrust (ft/s)
Temperatures (R)
Inlet temperature
2600 -- 3500
+ 0.081
+ 0.444
- 0.219
+ 0.202
Reheat temperature
2600 -- 3500
-
-
+ 0.200
+ 0.204
Pressures
Overall pressure ratio
18 -- 36
- 0.082
- 0.095
- 0.026
+ 0.021
Inner/outer fan pressure ratio
1.04 -- 1.28
- 0.003
+ 0.013
- 0.047
+ 0.047
Bypass/core pressure ratio
0.6 -- 1.32
0.026
0.036
0.031
0.028
Bypass
Bypass ratio
0.2 -- 0.76
- 0.076
- 0.25
+ 0.077
- 0.085
Mixer Mach number
0.1 -- 0.3
+ 0.001
- 0.001
+ 0.047
- 0.046
2.5 Off-design Group
2.5.1
Design altitude
Design Mach number and altitude are the most important parameters deciding the offdesign performance of the engine. Design reaches its peak -highest temperature and
compression ratio- at its predetermined design point. How these two variables effect the
engine is determined by investigating the nondimensional temperature which called theta
break (θbreak).
30
36
32
24
t=
Al
0
00
35
t=
Al
0
00
30
0
0
00
25
00
20
t=
Al
0
t=
Al
0
00
15
00
10
t=
Al
00
50
0
20
t=
Al
t=
Al
t=
Al
Overall Pressure Ratio P3/P2
28
16
12
0
.5
1
1.5
2
Mach Number
24.03.2012
GasTurb 11
Figure 2-16 : Total pressure ratio in flight
In previous regions than θbreak performance limited by P3/P2,after design point T4max will
start to limiting engine as can be seen in Figure 2-16.
Relationship between altitude and θbreak is linear and weak; meanwhile relationship
between altitude and Mach number is strong and parabolic. Therefore, operational engine
altitude has been chosen fixed at sea altitude.
2.5.2
Design Mach
After setting the design altitude at sea level, Design Mach number has been left as the
only variable affecting θbreak. Two important criteria have been set in designing TJ-1. A first
criterion was to minimize fuel consumption with dry configuration. Other criterion was to
maximize thrust generation with wet configuration. Therefore, effect of θbreak on engine has
been investigated with these criteria. In the following graphics θbreak values are marked with
red lines.
31
Figure 2-17 : Specific thrust in flight (Mcrit = 0, altdesign=0 ft, θbreak=1)
In Figure 2-17, specific thrust graphic with wet configuration can be seen. θbreak at this
configuration has been set to 1. Breaking points from the line trends can easily be seen in the
figure. These points are the θbreak’s of the engine. Lines linearize after θbreak values but
important thing to notice is the negative slope of the line. This requires θbreak to be higher than
1 in order to reduce thrust loss in high Mach numbers.
Figure 2-18 : Specific thrust in flight (Mcrit = 1.157, altdesign=25,000 ft, θbreak=1.05)
32
Figure 2-18, shows that with increasing θbreak, low Mach specific thrust values decline
while high Mach specific thrust values increase.
Figure 2-19 : TSFC in flight (Mcrit = 0, altdesign = 0 ft, θbreak = 1)
In Figure 2-19, dry TSFC values at θbreak=1 are given. TSFC values changes linearly
before and after θbreak. Slope declines a bit around θbreak.
Figure 2-20 : TSFC in flight (Mcrit = 1.157 altdesign = 25,000 ft, θbreak = 1.05)
33
In Figure 2-20, fuel consumption is lower while θbreak is 1.05. However, at the selected
design point, fuel consumption is higher. Engine is planned to conduct supercruise and thrust
requirement at maximum speed (1.8 Mach) will be very high. Therefore, θbreak should be
higher than 1.
2.6 Flow Rate (ṁ)
Airflow requirement of the engine is generally not considered as a variable in
parametric analysis. After a promising cycle is selected for engine, ṁ is scaled to provide
necessary thrust. However, this linear relationship is only valid for engines with high ṁ. As ṁ
declines axial compressor efficiency also declines. This effect can be seen in F125-GA-100
which has 4 axial and 1 centrifugal compressor stages. In order to operate with a flow rate
where axial compressor is still fuel efficient, ṁ is chosen to be higher than 110 lbf.
2.7 Summary
Following the investigation above, 3 important unknown parameters for the engine
which are; θbreak, bypass and overall pressure ratios. It should be noted that overall pressure
ratio is not a very independent parameter. Overall pressure is greatly depended on data
acquired from compressor design and engine dimensions. Considering fuel consumption and
engine dimensions, overall pressure is limited within a certain range. Using historical values,
it would not be a bad estimate for the ratio to be between 26 and 30. Therefore, overall
pressure ratio has been chosen as 28.
There will be a thorough investigation for the remaining parameters. A proper cycle
selection can’t be made without considering fuel consumption. Therefore off-design data,
acquired from all combinations of θbreak and bypass ratio have been used to determine the
coefficients in TSFC and thrust lapse equations.
In order to remove the negative effects of other design parameters (Fan pressure ratio,
mixer Mach number, core/bypass pressure ratio), Figures 2-24, 2-25, 2-26 show the engines
that have been optimized by GASTURB for minimum fuel consumption with Mach 1.4 and at
35000ft altitude.
34
Figure 2-21 : Mission Fuel saving percentages (contour lines)
Figure 2-24, shows contour lines defining fuel consumption percentagewise for the
baseline engine. As it can be seen, lowest fuel consumption is around critical Mach number
(0.8). There are two optimum values at this line; first one is around 0.2 bypass ratio which
decreases fuel consumption at augmented mission segments and the other around 0.46 which
decreases fuel consumption at non-augmented mission segments. It is preferable to choose the
optimum point at 0.46 since in real world the aircraft will not engage to a combat in every
mission.
35
Figure 2-22: Engine Airflow requirements “contour lines for flow rate (lbm/s)”
Figure 2-25, has been created to calculate air flows required to meet the thrust values at
RFP. As it can be seen, there exists a band around Mach 1 where engine has minimum flow
rate. Above this band, thrust ratio reaches desired values “9600/7400 (sea level static thrust /
1.8 Mach 40000 ft thrust)” and this allows the engine to be smaller.
Cross-referencing Figure 2-24 and 2-25, Mcrit has been chosen as 0.88 and the design
bypass ratio as 0.46 in order to achieve optimum fuel consumption and meet airflow
requirement.
Design altitude for the engine remains. Analytically, engine is expected to perform
constantly on the same θbreak line -neglecting altitude and Mach changes-. However,
simulation with GASTURB showed that due to mixer effects, thrust ratio of the engine
(9600/7400), changes at different altitude at constant θbreak. In this stage, design altitude has
been fixed at 6,000m / 20,000ft in order to minimize supercruise TSFC (Figure 2-26).
Afterwards to keep the thrust ratio of the engine fixed, θbreak has been rearranged to 1.1.
36
Mach Number = 0 ... 1,8
Altitude = 0 ... 11000 [m]
Contours: Off Design Sp. Fuel Consumption [g/(kN*s)]
12
.6
27
28
.8
32
30
28
.4
10
28
29
.6
3
29
.2
*10
8
31.6
27,6
28
28,4
28,8
29,2
29,6
30
30,4
30,8
31,2
31,6
32
6
4
30.8
30
.4
Altitude [m]
31.2
2
0
-2
-4
0
.5
1
1.5
2
Mach Number
16.03.2012
GasTurb 11
Figure 2-23 : Super cruise fuel consumption
Parametric cycle calculations have given the engine specifications that are shown at Table 23.
Table 2-3: Engine specifications
Overall pressure ratio
Design Mach number
Design altitude
Fan pressure ratio
LPC pressure ratio
HPC pressure ratio
Mixer Mach number
Maximum inlet temperature
Maximum afterburner temperature
28
1.1
20,000 ft
4.150
3.587
7.991
0.248
3000 oR
3200 oR
37
Table 2-4: Engine thrust and TSFC values
Mach
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
Altitude
(ft)
0
0
5000
10000
15000
20000
25000
30000
35000
40000
Dry Thrust
(lbf)
8076
7599
6364
5663
5321
5232
5270
5117
4784
3835
Wet Thrust
(lbf)
12159
11797
10347
9595
9317
9371
9551
9558
9281
7826
Dry TSFC
(lbm/(lbf*h))
0.670
0.736
0.788
0.838
0.884
0.924
0.962
0.985
1.007
1.054
Wet TSFC
(lbm/(lbf*h))
1.533
1.612
1.683
1.715
1.713
1.688
1.654
1.639
1.637
1.678
Inlet
Drag
0.0413
0.0130
0.0009
0.0028
0.0111
0.0198
0.0262
0.0293
0.0319
The equations to calculate TSFC at various mission points are given below:
Selected Engine
Maximum power:
{
(
)
(
}
)√
Military power
{
(
(
)
}
)√
Total fuel usage throughout the mission has been calculated for the engine whose
specifications were given at Table 2-3. The values have been compared with the fuel usage of
the reference engine.
38
1,0
700
500
0,8
400
300
200
0,6
Landing
Loiter
Descend
Subsonic cruise
Climb
Escape dash
Acceleration
Fire AIM-9Ls
0.9M turn
1.6M turn
Fire AMRAAMs
Supercruise
Acceleration
Air patrol
Descend
Subsonic cruise
Climb
Acceleration
Rotation
Acceleration
Taxi/Warm-up
100
0
Figure 2-24 : TJ-1 weight fractions and Fuel used through mission
Fuel difference for different mission segments, between reference and selected engines
are given in Table 2-5.
Table 2-5: Fuel save (lbf)
Taxi/Warm-up
Acceleration
Rotation
Acceleration
Climb
Subsonic cruise
Air patrol
Acceleration
56.44
-4.46
-0.41
20.23
27.00
115.52
53.60
0.4
Total
Supercruise
1.6 M turn
0.9 M turn
Acceleration
Escape dash
Climb
Subsonic cruise
Loiter
2.56
-14.56
-22.94
-9.20
1.94
1.06
32.72
63.96
323.86
Fuel used (lbf) -- (red)
Total weigth fractions -- (blue)
600
39
2.8
Power extraction
RFP requires calculations for maximum power that can be taken-off at the following
flight conditions; 1.4 M – 35000 ft and 0.9 M – 35000 ft.
Power Offtake = 0 ... 550 [kW]
Load Shaft Power Requ. = 0 ... 550 [kW]
PWX = 0
PWX = 50
PWX = 100
PWX = 150
PWX = 200
PWX = 250
PWX = 300
PWX = 350
PWX = 400
PWX = 450
PWX = 500
PWX = 550
Net Thrust [kN] = 6,5...14,5
700
6,5
7
7,5
8
8,5
9
9,5
10
10,5
11
11,5
12
12,5
13
13,5
14
600
12.5
500
12
Power Offtake [kW]
400
11.5
5
10.
11
300
10
200
7
8.5
100
7.5
9
0
Not converged points are marked red
0
100
9.5
8
200
6.5
6
300
400
500
600
700
Load Shaft Power Requ. [kW]
01.04.2012
GasTurb 11
Figure 2-25 : Power take-off variation at 0.9 M 35 kft
Figure 2-21 shows the engine thrust values at subsonic cruise segment of the mission.
Horizontal axis shows the power which is taken-off from low-pressure spool and the vertical
axis shows the power which is taken-off from high-pressure spool. As it can be seen in Figure
2-21, using HP-spool for power extraction is more efficient than using LP-spool. However, in
order to extract the maximum power without compromising the engine performance, power
should be taken-off equally from both spools. Furthermore, it has been decided to implement
this method to TJ-1 design and the power values have been given in Figure 2-22, 2-23.
40
33
21
32
20
30
29
28
27
Sp. Fuel Consumption [g/(kN*s)]
26
16
15
14
17
Net Thrust [kN]
18
31
19
44
40
36
32
28
24
Fan Surge Margin
22
21
20
19
18
HPC Surge Margin
23
48
24
52
25
0
200
400
600
800
01.04.2012
GasTurb 11
Figure 2-26 : Power extraction 0.9 M 35 kft
Figure 2-22, shows how much power can be extracted from the engine at 0.9 Mach
and 35000 ft. Maximum power extraction is 600 kW considering HPC and fan stall margins.
30
13.6
29
13.2
28
12.8
26
25
24
23
22
Sp. Fuel Consumption [g/(kN*s)]
27
12.4
12
11.6
11.2
10.8
10.4
Net Thrust [kN]
42
40
38
36
34
32
Fan Surge Margin
18
16
14
12
10
HPC Surge Margin
20
44
22
46
24
48
26
0
100
200
300
400
500
600
700
800
01.04.2012
GasTurb 11
Figure 2-27 : Power extraction 1.4 M 35 kft
Figure 2-22, shows how much power can be extracted from the engine at 1.4 Mach and
35000 ft. Maximum power extraction is 650 kW considering HPC and fan stall margins.
41
3
Inlet Design
3.1 Introduction
Even if TJ-1 is able to operate in both subsonic and supersonic regions, compressor’s
axial speed will remain constant. Inlet transmits the air to compressor at a specific speed,
independent from flight conditions. Inlets are designed to reduce the air speed to engines
operable conditions while minimizing the pressure loss. In subsonic region, 1% of pressure
loss in inlet approximately equals 1% loss in thrust generation. In supersonic region thrust
loss increases nonlinearly[5].
As RFP requested, TJ-1 is designed as 2 dimensional variable ramp with 2 external
oblique shocks. Moreover, considering stealth requirements, fan face of the engine shouldn’t
be hit directly by radio waves. In order to achieve that, two symmetrical ramps at each side of
the A/C are proposed to be attached to engine with s-ducts.
In order to acquire the dimensions of the inlet, mass flow rate of the engine
throughout the flight is needed. Corresponding area to that flow rate may also be used. Figure
3-1 is generated via GASTURB -considering MIL-E-5008B standards- and it shows the area
required with Mach number at different altitudes.
42
Figure 3-1 : Engine airflow requirement in flight
There is a 4% of safety margin at Figure 3-1 and boundary layer bleed requirements are
included (0.8 at 0% and 1.8 at 4%, linearly changing). As it can be seen, maximum required
inlet area by the engine is 2.96 ft2.
Isentropic flow equations are used for inlet design. In order to achieve maximum
pressure recovery, a definition introduced by Oswatitsh used [6].Oswatitsh states that in a
system with (n-1) oblique shocks and (1) normal shock, maximum recovery is achieved when
the oblique shocks have equal power. This definition is formalized in Equation (5).
(5)
The designed 2 ramped system has full variable geometry. In each and every Mach
number, both ramps will optimize their angles in order to maintain best pressure recovery. As
a safety margin, maximum Mach number is increased by 0.03 and taken as 1.83. Inlet
geometries corresponding to 1.4 and 1.8 Mach numbers are listed in Table 3-1.
43
Table 3-1: Ramp angles
1.4 Mach
51.13
4.006
60.28
3.608
First shock angle
First ramp angle
Second shock angle
Second ramp angle
1.83 Mach
41.79
8.777
53.12
9.023
After normal shock, stream enters into the transition zone of the inlet. Transition zone
ensures boundary layers to be re attached. According to Crosthwait [7], transition zones length
is twice the diameter of the throat. After the transition zone, stream enters into diffuser.
Diffuser geometry differs from a rectangular cross section to a circular cross section by super
elliptical cross sections in transition.
Figure 3-2 : Inlet geometry
At zero flight speed pressure recovery is calculated as 0.804 which is under acceptable
limits. Therefore, auxiliary air inlets should be included into the design. Reverse calculations
are made for ƞr = 0.95 and Mach number at throat is found as 0.285. Required extra auxiliary
air inlet area for that Mach number is 2.854 ft2. TJ-1 includes two identical inlets, thus
dividing by 2 the resulting area has been found as 1.427ft2. These auxiliary inlet doors will
ensure the desired pressure recovery. Detailed drawing of inlet is given in APPENDIX -B-
44
Figure 3-3 : Inlet pressure recovery ratio
Pressure recovery performance of the designed inlet is given at Figure 3-3 along with
the MIL-E standards. In addition, dimensions and geometry of the inlet are given in
APPENDIX -B-;
Table 3-2: Duct specifications
γ
Mi
Me
Ae/Ai
L/H
η
Pte/Pti
T (K)
Inlet duct
1.399
0.846
0.5
1.311
4
0.91
0.988
310
LC duct
1.390
0.55
0.47
1.117
1.8
0.6
0.986
465
HC duct
1.347
0.25
0.06
4.026
3
0.89
0.996
866
HT duct
1.317
0.33
0.2
1.586
2
0.65
0.986
1270
LT duct
1.331
0.6
0.25
2.030
3
0.89
0.985
1054
Duct calculations have been performed in Inlet section and the results are given in Table 3-2.
45
4
Combustion Systems
TJ-1 has 2 different combustion modules. First -main burner- is between compressor
and turbine, and active throughout the flight. Second -afterburner- is between turbine and
nozzle and activated when necessary. Design and optimization of the combustion system is
done using equations in Mattingley[8] and using EXCEL as a calculator.
4.1
Burner
Generally burners are divided to three (can, cannular, annular). TJ-1 uses annular
combustion room, which is lighter and has a lower pressure loss.
Design point for burner is chosen as 1.25 Mach at sea level. This point also includes the
maximum dynamic pressure for the engine. Results of the compressor and turbine analyses
are required for burner design. Data’s used in design are listed in Table 4-1.
Table 4-1: Stations dimensions
Station 3.1
Station 4
Burner
Router (in.)
6.69
10.35
9
Rinner (in.)
6.02
9.76
7.04
Rmean (in.)
6.38
10.05
8.02
H (in.)
0.669
0.591
1.956
46
In order to have an efficient combustion, air coming from HPC has to be slowed down
to operable speeds. Diffuser of burner should be able to complete this by minimal pressure
loss.
Size criteria’s for diffuser forces burner to have 2 splitters. No dump, flat wall
geometry is selected for diffuser and length of the diffuser is calculated as 4.425 inch using
the equations in Mattingley[8]. In addition, dome radius is calculated as 1.394 inch. Also for
an adequate mix, total pressure loss is 2.84 psi which is 50% of the allowable value (5.59 psi).
For air partitioning calculations, liner material is chosen as any Hastalloy able to
withstand 20000R. Also, liner cooling is chosen as transpiration cooling. Results of the
calculations are given in Table 4-2.
Table 4-2: Air partitions (Tg= 4174 0Ra, εPZ = 0.8)
Total
Primary
Secondary
Transpiration
Dilution
Zone
zone
Cooling
Zone
Air flow (lbm/s)
170.2
63.65
27.23
42.04
37.10
Mass fractions
1.00
0.374
0.160
0.247
0.218
Figure 4-1 : Swirler layout
47
Swirler blades have been chosen as airfoil cross-sections with 0.64 drag coefficient
o
and 35 blade angle. After calculations, swirlers are arranged as in Figure 4-1. S’ swirl
number is found as 0.61 (which is just a bit more than minimum value 0.6).
Table 4-3: Zones geometry
Nprimary
35
Lprimary (in.)
0.847
Nsecondary
512
Lsecondary (in.)
2.789
Ndilution
250
Ldilution (in.)
2.091
Zone calculations are conducted via Mattingley[2] equations. Results are listed in Table 4-3. Finally,
the burner dimensions are given in Table 4-4.
Table 4-4: Burner geometry
4.2
Length (in.)
5.726
Diameter (in.)
1.956
Total Volume (ft3)
2.246
Combustion Zone(ft3)
1.601
Afterburner
Afterburner radius is selected as 13.78 in., which is the maximum diameter of the engine.
Stream leaving turbine is mixed with bypass stream before entering afterburner. Physical properties of
different streams are listed in Table 4-5. These values are acquired from the parametric study.
Table 4-5: Flow areas
Station 5
Station 6
Station 13
Station
Station
16
6A
Station 7
Router (in.)
11.42
13.28
13.78
13.78
13.78
Rinner (in.)
9.64
8.28
13.28
8.24
0
H(in.)
1.78
5.00
0.5
5.54
13.78
A (ft2)
0.814
2.354
0.293
2.660
4.028
48
In mixer design, in order to keep the dimensions as small as possible, a mixer-diffuser design
has been chosen. In addition, diffuser efficiency is chosen as 0.9 as it was stated in Mattingley[2] for
flat wall and dump diffuser’s. Table 4-6 lists the general specifications of diffuser.
Table 4-6: Mixer +diffuser dimensions
Station 6A
Station m
Station 6.1
Router (in.)
13.78
13.78
13.78
Rinner (in.)
8.24
4.36
0
Rmean (in.)
11.01
9.07
6.89
H(in.)
5.54
9.42
13.78
A (ft2)
2.660
3.728
4.028
Figure 4-2 : Mixer + diffuser layout
Vee-gutter angle of flameholders are chosen as (2θ = 300). In addition W/H is chosen as 0.4 in
order to keep the pressure loss minimum. Ring number is chosen as 2, considering size limitations.
Afterburner geometry is given in Figure 4-3.
Figure 4-3 : Flameholders layout
49
Finally using the equations in Mattingley [2], important data’s for burner and afterburner are calculated.
Results with design guideline are given in Table 4-7.
Table 4-7: Combustion parameters
Combustor loading
(kg/s atm1.8 m3)
Combustor intensity
(MW/atm m3)
Combustor loading
(kg/s atm1.8 m3)
Mafterburner
Burner
Design guideline
0.5
Maximum 10
30
Maximum 60
Afterburner
Design guideline
5.16
Maximum 100
0.23
Maximum 0.3
50
5
Compressor
The very first thing that must be determined is the compressor type and number. Aircraft
engines usually employ two types of compressors -axial and centrifugal-. These two types
have different pros and cons and they may also be used together. Centrifugal compressors are
generally utilized at smaller flow rates than TJ-1 has. All of these configurations are analyzed
thoroughly.
5.1
General Information
RFP demands two spool low bypass turbofan, however this does not mean only two
compressors must be used. More than one compressor on a same shaft is also possible.
For TJ-1, centrifugal compressor is considered instead of last four stages of high pressure
compressor. However, efficiency calculations showed that efficiency of a centrifugal
compressor -which is placed instead of last four stages of axial compressor-, will be between
0.75 and 0.8. Although their tolerance to rapid flow rate change is lower, axial compressor has
been found the most suitable type considering its relatively high corrected flow rate and better
efficiency.
51
Before starting the compressor design, one must determine which types of assumptions
are going to be made, consequently how much the results of this design are going to be
compatible with real life compressor behavior.
Since there are no detailed data for the compressors’ design at this level, some
simplifications and assumptions are necessary. For high pressure compressor; repeating row,
repeating stage, mean line design, meanwhile for fan; constant tip radius, repeating row and
repeating stage design has been found appropriate for preliminary calculations. The flow
properties are assumed constant throughout circumferential location and span-wise direction.
Calculations are only going to be performed for the mean line properties. Axial velocities are
constant and air is assumed to be calorically perfect gas with constant γ and known R. Swirl
angles are also constant along the stages. Free vortex swirl model has been used.
Besides these assumptions, blade tip Mach number is selected as 1.27 for fan and 0.93
for high pressure compressor. These values are same as baseline model’s blade tip Mach
numbers. This has been done in order to fulfill the requirement of using replica of blades,
which have also been used for baseline model. It is clear that blade tip Mach numbers are the
first constraint for the designed compressor. Second constraint is the dimensions of the engine
itself. This engine is going to be mounted on a half scaled model of JSF. Data from the
parametric study are used in order to determine the lowest dimensions possible for obtaining
thrust needed. Dimensions are not the only challenge, stress in the blades and rims must be
considered. In addition, blade and rim stresses are strongly depended on the dimensions of the
compressor.
Another thing that must not be forgotten is fan and high pressure compressors’ RPMs
have to be same with low and high pressure turbines. In order to design an efficient engine, all
the engine parts must work efficiently with another. Therefore, turbines and compressors have
been designed separately, and then optimized together. This may decrease the turbine and
compressor efficiencies separately, however overall efficiency is the more important issue.
As the first step of the design process, data from parametric study are gathered. Then
stage counts have been determined, paying attention to stage loading and flow coefficients.
Considering today’s technological limits, stage loading is limited to 0.7[1] and flow coefficient
generally varies from 0.45 to 0.55[8]. Main goal of this design process is to introduce the most
efficient compressor without exceeding limits and constraints given previously.
52
5.2
Fan design
Fan is one of the most crucial parts of the engine. To be able to design the fan properly,
all constraints and requirements must be met. Some initial parameters are selected regarding
today’s trends to begin the designing process. Combining gathered data from parametric study
and initial parameters, fan is designed.
Since fan is the widest part, rim and blade stresses are very important. During design
process, maintainability and costs are also considered. Therefore titanium alloys are used.
Taper ratio of 0.8 is also used for rotor blades to reduce stress at blades. Again in order to
reduce stress levels in the blades, hollow fan blades are considered to put in use.
Beginning of the process includes a number of estimated values such as solidity,
polytrophic efficiency and diffusion factor. Solidity is simply a measure of blades’ closeness.
A design with closer blades will have more loss due to boundary layer separations. Having
very less solidity (chord length divided by pitch length) causes stages’ compression to be
insufficient. Even though the baseline model has a very high solidity value, solidity of 1.1 is
employed for designed fan. The design has already more loss due to boundary layer issues
because of lower dimensions. If blades are placed very closely, amount of losses will affect
the performance significantly. Accordingly, diffusion factor indicates danger of boundary
layer separation. Values over 0.6 lead to dramatic pressure losses. Higher values mean less
stage number, consequently less weight. Considering today’s technology level diffusion factor
is chosen as 0.55. Polytrophic efficiency is actually isentropic efficiency for infinitively small
volume. Hence higher efficiency means better performance. However efficiency is limited by
technology. Knowing that, a realistic value of polytrophic efficiency is assigned which is
0.89. Generally highest limits of general trends are chosen for design process.
All of the flight conditions are known from the parametric study and mission profile. As
the high pressure compressor and fan are designed for a specific design point, only the
calculations for this point are performed. Test rig data is required for real life compressor
behavior and this situation is beyond the scope of this project.
First stage hub to tip ratio of the fan is found after a literature check to be around 0.4 or
0.5 for military low by-pass turbofans. The present design has been decided to have a hub to
tip ratio of 0.51 to avoid any material failure. More importantly, constant tip radius design
allows the back of the low pressure compressor (fan) to be wide. This spacing is decided to be
used for the integrated starter–generator system. By this way even if the required gap for
53
starter generator system is not provided, engine’s total length will be affected very little for
the installation of starter generator system. All stages’ hub to tip ratios are arranged to keep
tip radius constant.
Typical values of the blade aspect ratios vary 1.5 to 3.5[1]. Increasing aspect ratio of fan
causes operation point to move closer to the surge line and costs more[1]. Low aspect ratio
means higher chord length, thereby leading a longer fan. Considering all of these, aspect
ratios of the rotor and stator blades are selected as 2.7. Main reason for this choice is the
request of using replica blades. In other words, aspect ratio of blades has to be same as
baseline model if present design’s blades are just scaled copies.
Besides that blade gapping is also an important issue for the axial compressors. Blades
of rotors and stators are prone to interact with each other. This situation requires leaving a gap
between blades large enough to minimize vibratory excitation. Gap is generally at least 20%
chord length of related blades[1]. Increasing blade gap enormously is not a good choice since
length of the engine is increasing. A conservative approach is made for present design and
blade gap value is selected 25% chord length of the related blades.
Inlet guide vanes are decided to be utilized for the design. In the absence of inlet guide
vanes, fan blades have to turn faster in order to compensate the lack of pre-directed flow. Inlet
guide vanes also increases the compressor inlet Mach number which increases the
performance of the fan. The flow through guide vanes is assumed to reach Mach 0.6 before
reaching the first stage of fan.
Design of the fan is carried out using AEDsys software and Microsoft Excel.
GASTURB is also an alternative however, AEDsys is found more suitable for part designing
process. All values of the fan is calculated via EXCEL and validated with AEDsys. After
validation, some of the values required by AEDsys are also found from calculations done on
EXCEL. Furthermore, mean line properties are given only in consequence of mean line
design assumption. Input and output values of AEDsys software is shown in Table 5-1.
54
Table 5-1: Fan input values
Fan
Number Of Stages
3
Mass Flow Rate (lbm/s)
105.41
Rotor Angular Speed (rad/s)
1275
Inlet Total Pressure (psia)
14.56
Inlet Total Temperature (R)
557.12
Entry Angle (degrees)
31.60
Entry Mach
0.60
Diffusion Factor
0.55
Rotor Chord/Height Ratio
0.59
Stator Chord/Height Ratio
0.59
Polytrophic Efficiency
0.89
Solidity
1.10
Exit Angle For Last Stage
31.60
Exit Mach For Last Stage
0.48
Ratio of Specific Heats
1.40
Gas Constant (J/kg K)
53.34
Rotor angular speed is dependent on tip speed, mean radius and hub to tip ratio of the
stage. Consequently, using replica blades sets the angular speed. Entry Mach number was
estimated from historical data[1].
55
Table 5-2: Fan output value
Fan
Total Temperature (R)
Static Temperature (R)
Total Pressure (psia)
Static Pressure (psia)
Mach
Velocity (ft/s)
Radius (in.)
Flow Area (in2)
Hub to Tip Ratio
Delta Tt (K)
Tt3/Tt1
Pt3/Pt1
Rotor Chord Length (in.)
Stator Chord Length (in.)
Stage Loading Coefficient
Flow Coefficient
Rotor Blade Number
Stator Blade Number
Rotor Mean Speed (ft/s)
Isentropic Efficiency
Reaction Factor
AN2 (1010)
Total Length (in.)
Total Compression Ratio
Rotor Entry
Mean
557.10
519.70
14.56
11.42
0.60
670.00
9.72
448,00
0.45
First Stage
Rotor Exit
Mean
651.10
556.80
24.06
13.92
0.92
1064.00
10.36
393.81
0.55
93.96
1.17
1.63
2.48
1.94
0.53
0.55
29.00
39.00
1032.50
0.89
0.40
5.84
Stator Entry
Mean
651.10
613.70
23.81
19.36
0.55
670.00
11.16
312.09
0.67
Rotor Entry
Mean
651.10
613.70
23.81
19.36
0.55
670.00
11.16
312.09
0.67
Second Stage
Rotor Exit
Mean
745.00
660.90
36.55
24.03
0.80
1005.00
11.52
270.69
0.72
93.96
1.14
1.52
1.52
1.27
0.40
0.48
53.00
65.00
1185.90
0.88
0.50
4.01
9.94
3.59
Stator Entry
Mean
745.00
707.60
36.23
30.25
0.51
670.00
11.84
230.26
0.77
Third Stage
Rotor Entry Rotor Exit Stator Entry
Mean
Mean
Mean
745.00
839.00
839.00
707.60
758.60
801.60
36.23
52.72
52.30
30.25
37.05
44.59
0.51
0.73
0.48
670.00
983.00
670.00
11.84
12.06
12.23
230.26
201.53
176.97
0.77
0.80
0.83
93.96
1.13
1.44
1.07
0.92
0.36
0.45
78.00
92.00
1257.90
0.88
0.54
2.99
56
As can be seen from Table 5-2 flow coefficients and stage loading coefficients are
consistent with allowable margins. Moreover, as a result of the assumptions that made before,
mean axial velocity, tip radius, rotor mean blade speed, stage loading coefficient, flow
coefficient and increase in the total temperatures do not change inter stages. Furthermore,
degree of reaction is very close to 0.5. This is very satisfactory because, sharing the burden of
static temperature rise is expected from stators and rotors in general case. By this means,
excessive values of diffusion factor are also avoided[2].
There is one more point that must be emphasized. Usually bypass air stream has a
slightly higher compression ratio than core stream because of increasing blade speeds near the
blade tip. AEDsys software, does not have such calculation mode, however one can realize
outer fan compression ratio is 1.157 times inner fan compression ratio from parametric study
data.
5.3
High Pressure Compressor Design
High pressure compressor design is carried out, using the same approximation and
assumptions with only one difference. Mean line design is preferred since blade height is low
relative to mean radius. This type of design gives better results for this situation.
Same process that has been carried out for fan design is repeated. First, the stage count
is determined from the total temperature rise and 9 stages are found suitable. Besides,
polytrophic efficiency, solidity and diffusion factor are assumed to be 0.89, 1.1 and 0.55
respectively. These values are almost at the limit of the general trends.
Hub to tip ratio of the first stage is chosen to satisfy the given tip radius. RPM of the
high pressure compressor is only depend on tip radius and tip speed. Tip speed is found from
the circumferential tip Mach number constraint. With the radius known, RPM is calculated
easily. Throughout this process high pressure compressor and high pressure turbine is
calculated together. Since the high pressure turbine is more demanding because of the high
temperatures, compressor’s tip radius is determined from required high pressure turbine RPM
and radius. Aspect ratio of the compressor blades have to be same as the blades used in
baseline model which is 2.7.
Entry angle of the high pressure compressor and exit angle of the fan is different.
Furthermore, fan and the high pressure compressor is counter rotating. In consequence of this,
57
there is a guide vane is placed between the two. This guide vane basically removes the swirl
of the fan and adjusts the entry angle of high pressure compressor.
All of the entry properties such as entry Mach number, total temperature and total
pressure are known from the last stage fan properties. Flow rate is decreased due to bypass air
flow.
HPC input values can be seen in Table 5-3.
Table 5-3: HPC input values
High Pressure Compressor
Number Of Stages
9
Mass Flow Rate (lbm/s)
72.20
Rotor Angular Speed (rad/s)
1970
Inlet Total Pressure (psia)
52.24
Inlet Total Temperature (R)
835.92
Entry Angle (degrees)
29.4
Entry Mach
0.47
Diffusion Factor
0.55
Rotor Chord/Height Ratio
0.37
Stator Chord/Height Ratio
0.37
Polytrophic Efficiency
0.89
Solidity
1.10
Exit Angle For Last Stage
29.4
Exit Mach For Last Stage
0.33
Ratio of Specific Heats
1.40
Gas Constant (J/kg K)
53.34
All results of HPC calculations can be seen in Table 5-4. In addition, cross sections of
Fan and HPC are given in APPENDIX -C-;
58
Table 5-4: HPC output values of 1-3. Stages
High Pressure
Compressor
Mach
Velocity (ft/s)
Flow Area (in2)
Hub to Tip Ratio
Tt3/Tt1
Pt3/Pt1
Rotor Blade Number
Stator Blade Number
AN2 (1010)
Flow Coefficient
Reaction Factor
Radius (in.)
Delta Tt (K)
Total Length (in.)
First Stage
Rotor Entry
Mean
835.90
800.60
52.24
44.91
0.47
652.00
120.95
0.61
Rotor Exit
Mean
922.90
840.00
71.81
51.64
0.70
998.00
110.36
0.64
1.10
1.36
1.70
1.53
27.00
30.00
0.89
3.91
Second Stage
Stator Entry
Mean
922.90
887.60
71.27
62.16
0.45
652.00
96.87
0.68
Rotor Entry
Mean
923.00
887.60
71.27
62.16
0.45
652.00
96.87
0.68
Rotor Exit
Mean
1010.00
927.00
95.22
70.54
0.67
998.00
89.16
0.70
1.09
1.33
1.37
1.24
33.00
36.00
0.89
3.16
0.48
0.54
0.45
6.36
87.02
1043.60
12.20
7.96
Third Stage
Stator Entry
Mean
1010.00
974.60
94.56
83.47
0.43
652.00
79.22
0.73
Rotor Entry
Mean
1010.00
975.00
94.60
83.50
0.43
652.00
79.22
0.73
Rotor Exit
Mean
1097.00
1014.00
123.40
93.70
0.64
998.00
73.45
0.75
1.09
1.30
1.12
1.03
40.00
44.00
0.89
2.60
Stator Entry
Mean
1097.00
1062.00
122.60
109.30
0.41
652.00
65.92
0.77
59
High Pressure
Compressor
Mach
Velocity (ft/s)
Flow Area (in2)
Hub to Tip Ratio
Tt3/Tt1
Pt3/Pt1
Rotor Blade Number
Stator Blade Number
AN2 (1010)
Flow Coefficient
Reaction Factor
Radius (in.)
Delta Tt (K)
Total Length (in.)
Fourth Stage
Rotor Entry
Mean
1097.00
1062.00
122.60
109.30
0.41
652.00
65.92
0.77
Rotor Exit
Mean
1184.00
1101.00
156.70
121.50
0.61
998.00
61.48
0.78
1.08
1.27
0.94
0.86
48.00
52.00
0.89
2.18
Fifth Stage
Stator Entry
Mean
1184.00
1149.00
155.70
140.00
0.39
652.00
55.65
0.80
Rotor Entry
Mean
1184.00
1149.00
155.70
140.00
0.39
652.00
55.65
0.80
Rotor Exit
Mean
1271.00
1188.00
195.70
154.50
0.59
998.00
52.17
0.81
1.07
1.25
0.79
0.73
56.00
61.00
0.89
1.85
0.48
0.54
0.45
6.36
87.02
1043.60
12.20
7.96
Sixth Stage
Stator Entry
Mean
1271.00
1236.00
194.50
176.20
0.38
652.00
47.57
0.83
Rotor Entry
Mean
1271.00
1236.00
194.50
176.20
0.38
652.00
47.57
0.83
Rotor Exit
Mean
1358.00
1275.00
240.80
193.10
0.57
998.00
44.80
0.84
1.07
1.23
0.68
0.63
66.00
70.00
0.89
1.59
Stator Entry
Mean
1358.00
1323.00
239.50
218.40
0.37
652.00
41.10
0.85
60
High Pressure
Compressor
Mach
Velocity (ft/s)
Flow Area (in2)
Hub to Tip Ratio
Tt3/Tt1
Pt3/Pt1
Rotor Blade Number
Stator Blade Number
AN2 (1010)
Flow Coefficient
Reaction Factor
Radius (in.)
Delta Tt (K)
Total Length (in.)
Seventh Stage
Rotor Entry
Mean
1358.00
1323.00
239.50
218.40
0.37
652.00
41.10
0.85
Rotor Exit
Mean
1445.00
1362.00
292.50
237.90
0.55
998.00
38.85
0.86
1.06
1.22
0.59
0.55
76.00
81.00
0.89
1.38
Eighth Stage
Stator Entry
Mean
1445.00
1410.00
291.00
266.80
0.35
652.00
35.84
0.87
Rotor Entry
Mean
1445.00
1410.00
291.00
266.80
0.35
652.00
35.84
0.87
Rotor Exit
Mean
1532.00
1449.00
351.40
289.20
0.54
998.00
34.00
0.87
1.06
1.20
0.51
0.48
86.00
92.00
0.89
1.20
0.48
0.54
0.45
6.36
87.02
1043.60
12.20
7.96
Ninth Stage
Stator Entry
Mean
1532.00
1497.00
349.70
322.20
0.34
652.00
31.52
0.88
Rotor Entry
Mean
1532.00
1497.00
349.70
322.20
0.34
652.00
31.52
0.88
Rotor Exit
Mean
1358.00
1275.00
240.80
193.10
0.57
998.00
44.80
0.84
1.06
1.19
0.45
0.43
98.00
104.00
0.89
1.06
Stator Entry
Mean
1358.00
1323.00
239.50
218.40
0.37
652.00
41.10
0.85
61
5.4
Structural Consideration:
TJ-1 engine has higher RPMs than most of the turbofans available at market. Because of
this, special amount of consideration must be given for structure design part. Below,
centrifugal stresses and static temperatures are calculated for every stage and separately
tabulated for fan and HPC in Table 5-7.
Table 5-7: Fan centrifugal stress
Fan
Stage
1
2
3
Centrifugal Stress (ksi)
76.77
53.80
39.69
Centrifugal Stress (MPa)
529.29
370.96
273.64
T1 (R)
519.70
613.70
707.60
T1 (K)
288.70
340.90
393.10
Material choice strongly affects total weight. Therefore, the chosen material must be
strong enough to withstand the stresses while ensuring the lightest design. Greek Ascoloy has
been found convenient for being strong enough at all stages’ temperatures. Additionally,
Greek Ascoloy’s density is fairly low compared to nickel based alloys. Considering these,
Greek Ascoloy has been chosen for fan and HPC.
Table 5-8: HPC centrifugal stress
Stage
1
2
3
4
5
6
7
8
9
Centrifugal Stress (ksi)
49.52
39.75
32.57
27.15
22.96
19.66
17.01
14.86
13.08
Centrifugal Stress (MPa)
341.45
274.06
224.58
187.21
158.33
135.56
117.30
102.45
90.21
T1 (R)
800.60
887.60
975.00 1062.00 1149.00 1236.00 1323.00 1410.00 1497.00
T1 (K)
444.78
493.11
541.67
590.00
638.33
686.67
Properties of Greek Ascoloy are tabulated below for comparison;
735.00
783.33
831.67
62
Table 5-9: Greek Ascoloy properties
Yield Strength
Rankine
Yield Strength (MPa)
273.15
491.67
1000.00
145.04
1200.00
174.05
373.15
671.67
1095.00
158.82
1290.00
187.10
473.15
851.67
1170.00
169.69
1380.00
200.15
573.15
1031.67
1240.00
179.85
1470.00
213.21
673.15
1211.67
1270.00
184.20
1500.00
217.56
773.15
1391.67
1270.00
184.20
1510.00
219.01
873.15
1571.67
1220.00
176.95
1450.00
210.30
(ksi)
Ultimate Strength (ksi)
Ultimate Strength
Kelvin
(MPa)
63
6
Turbine
Conventional turbine designs for flow rates greater than 30 lbm/s are mostly axial
turbines. Thus, design process begins with the decision of both high pressure turbine (HPT)
and low pressure turbine (LPT) to be axial. In engines world, there is never a best turbine
design for a given application, it is always a tradeoff of several parameters, such as, rotor
stress, weight, outside diameter, efficiency, noise, durability, and cost, so that the final design
lies within acceptable limits for each parameter. The primary goal of the turbine design was to
develop turbine stage performance models which will meet the parametric design
requirements by juggling around these parameters.
RFP calls for an engine to mount on a half-sized model of JSF. This is the main
challenge in all aspects of the engine design process, thus turbine design process as well. The
expansion ratio of the turbines, the need of fewer stage numbers and the counter-rotating
spool dynamics, led to a brand new design in spite of being a replica of the baseline engine.
This choice reveals a bunch of challenges to be overcame in the design process. Challenges
have been mainly categorized as structural and aerodynamic.
64
The structural limitations may be subcategorized into; uncooled configuration of LPT,
material and cooling limitations for the high rotational speeds of HPT. The aerodynamic
limitations may be subcategorized as; choking turbine rotor entries while maintaining the
subsonic flow over the airfoil and correlating all these design decisions within the
compressors’ feasible operating RPM number. HPT and LPT specifications are selected in the
ranges of open literature sources and researches conducted by NASA. The design decisions
and results are given explicitly in HPT and LPT design sections. Typical turbine stage and
velocity diagrams are given at Figure 6-1. For better understanding, in further explanations,
subscripts of the quantities will be used due to the station numbers in Figure 6-1.
Figure 6-1 : Turbine stages and velocity diagrams
6.1 HPT design
As it is stated in the RFP, a half-size JSF will be mounted with the designed engine.
According to Raymer equations (see chapter 2.4), the design length of the engine will
approximately be 142.3 in (3.6 m). Therefore, a tradeoff should be made to check whether it is
better to design a 1-stage turbine, while forcing the design limits, or to design a 2-stage
turbine, as it is in the baseline engine F100-PW110-229. A simple comparison for these two
designs is made through rough estimations. Multi-staged design offers lower centrifugal
loading and higher efficiency while single stage design offers a large saving in initial engine
cost, weight, and maintenance cost because of the significant reduction in the number of
components, especially expensive cooled airfoils. Considering advantages of the designs, a
single-stage design is found appropriate.
65
Since it is impossible to generate a single-stage turbine design based on the baseline
engine, replication blades –as suggest in RFP- have been ruled out. A new approach will be
investigated to obtain the cycle analysis parameters listed in Table 6-1.
Table 6-1: HPT design point parameters (1.1 M/20kft)
τ = 0.78
Pt 4.1 = 393.2 psi
γ = 1.3
п = 0.299
Tt 4.1 = 3000 R
gcCp = 7445 ft2/(s2.0R)
N = 18812 RPM
ṁ 4.1= 68.08lb/s
R = 53.4 ft.lbf/(lbm.0R)
In this approach, a NASA report[12] has been used as a baseline. In the report, a singlestage uncooled turbine design -which has a rim speed higher than the suggested values in the
literature-, is tested in a cold air rig. The design and test parameters are given Table 6-2:
Table 6-2: Conditions of cold air rig test of single-stage turbine at NASA report[12]
Performance Parameter
Design conditions
Test condition
R (K)
3960 (2200)
518.7 (288.2)
Inlet total pressure, P0
psia (N/cm2)
560 (386.1)
14.7 (10.13)
Mass flow, ṁ
lbm/s (kg/s)
108.92 (49.41)
8.501 (3.856)
Turbine rotative speed, N
RPM
21772
8081
Blade tip speed, Ut
ft/s (m/s)
1900 (579.1)
705.3 (215)
0
Inlet total temperature,T0
The performance parameters for a turbomachine operating with compressible flow
mechanics, can be expressed as
{
}
̇
Further investigation of the isentropic relations between temperature and pressure, more
useful non-dimensional functions can be expressed.
{
̇√
√
}
The first parameter at the right-hand side is the equivalent mass parameter (ṁeq) and the
second one is the equivalent speed parameter (Neq) where δ is the ratio of total pressure at the
inlet to the U.S standard sea-level pressure and θ is the squared ratio of critical velocity at the
66
turbine inlet to the U.S. standard sea-level critical velocity. The stage loading of this design is
1.94 and the expansion ratio is 3.44. The parameters are listed in Table 6-3.
Table 6-3: Performance parameter values
Test values
Design Values
TJ1 design values
ṁeq
8.501
7.714
6.22
Neq
8081
8066
7970
In the report, designed turbine is validated by the tests while maintaining correlated
aerodynamic and thermodynamic similarities. In Table 6-3, it can be clearly seen that there is
a difference between TJ-1 design values and the test values. This means, correlations will not
be enough to consolidate a direct application of the designed turbine to TJ-1. Further
calculations and tests are needed to meet the parametric cycle data. Furthermore, it is not
acceptable for an equivalent design for TJ-1 but enough to reveal that the design is feasible.
Based on the report, a mean-line design has been conducted. In order to investigate the
conditions, TURBN.exe[11], a mean-line turbine design program, is used for obtaining
preliminary design quantities in aerodynamic and thermodynamic manner by using numerical
calculations. Since it is not possible to provide comprehensive methods for turbine design,
calculations are made under constant axial speed and adiabatic assumptions with selected
relative Mach number constraints for stator and rotor airfoils. However, these conditions
stated above are sufficient to analyze the turbine behavior. Analyses will reflect the engine
cycle assumptions and resemble real turbine designs.
An EXCEL sheet have been prepared with consideration of the formulas stated in
AIAA-Aircraft Engine Design[8], Turbine Aerodynamics section that is consistent with the
TURBN program, and also with the assumptions of; selection of M2 and M3R, 2D flow,
constant mean radius, adiabatic flow throughout the turbine stages, calorically perfect gas
with constant R and γ. The aim for this EXCEL sheet is to determine input parameter
quantities for the TURBN program by the SOLVER plug-in, to obtain desired output
parameters.
In order to meet HPT requirements, such as enthalpy drop -found from parametric
cycle analysis- and 1970 rad/s rotational speed -to match compressor design speed-,
parameters in Table 6-1 and SOLVER plug-in has been used. U3/u2 is decided to be 0.9 in
order to set a smaller value for M3R. In addition, remaining parameters are found as; etL=0.89,
67
M2 = 1.13, M3R = 0.93 and the stator exit angle α2 = 74.50. Using these parameters as input for
URBN.exe -keeping the stator choked and M3R subsonic-, following results have been found:
Table 6-4: Results of the TURBN.exe for HPT -11h
1m
1t
2h
2m
2t
2Rm
3Rm
3h
3m
3t
Tt
0
3000
3000
3000
3000
3000
3000
2061
2601
2341
2341
2341
T
0
R
2982
2982
2982
2466
2518
2562
2518
2308
2307
2308
2309
Pt
Psia
393.2
393.2
393.2
381
381
381
205.4
185.6
117.4
117.4
117.4
P
Psia
383.1
383.1
383.1
163
178.3
192.2
178.3
110.6
110.4
110.6
110.7
0.2
0.2
0.2
1.20
1.13
1.07
0.47
0.92
0.31
0.31
0.30
R
M
V
ft/s
516
516
516
2819
2680
2555
1115
2089
705
695
687
u
ft/s
516
516
516
716
716
716
716
644
644
644
644
v
ft/s
0
0
0
2726
2582
2453
855
1987
285
260
239
α
Deg
0
0
0
75.3
74.5
73.7
23.8
21.9
20.3
β
Deg
Radii
in.
9.60
10.52
11.44
10.09
10.52
10.95
9.96
10.52
11.08
50
72
10.52
10.52
Table 6-5: Results of the TURBN.exe for HPT -2Hub
Rt = 0.2409
A1 = 56.45 in2
Mean
Rt = 0.3178
A2 = 73.82 in2
Tip
Rt = 0.3835
A3 = 121.28 in2
Table 6-6: Results of the TURBN.exe for HPT -3Stage-load
Flow
Isentropic
Aspect Ratio
Solidity
Blade number
coefficient
coefficient
efficiency
Vane
Rotor
Vane
Rotor
Vane
Rotor
1.645
0.415
0.903
1
1.11
0.924
1.487
62
74
Three parameters controlling the efficiency (stage-load coefficient, flow coefficient
and reaction) are stated above and found consistent with the boundaries given in literature.
From Figure 6-2, uncooled total-to-total turbine efficiency (ηtH) is found as 0.905.
68
Figure 6-2 : Smith chart for turbine stage efficiency[9]
6.1.1
Structural analysis of HPT
After setting up the aerodynamic and thermodynamic design of the HPT, two
significant matter remains. First, to be able to withstand the stresses generated. Secondly, to
be able to operate for estimated working hours without material or mechanical failure.
In turbine structural analysis, the limiting factor would most likely be the creep
behavior, especially for the turbines operating at high temperatures and high rotational speeds.
TJ-1’s HPT is opposed to a higher creep risk due to the higher rotational speed. The
centrifugal stress is directly proportional to the material density. Thus, a semi-iterative method
should be followed while calculating the stresses and selecting the materials.
For the rotor, taper ratio (At/Ah) is selected within the recommended boundaries[9] as
0.7 to reduce the stress. The average annulus area (Aav) is 97.6 in2 and the rotational speed (ω)
is 1970 rad/s. Centrifugal stress (σc) is found as 37.7 ksi. SC 16 single crystal Nickel alloy has
been chosen for the nozzle guide vane (NGV) and the rotor blades of the HPT.SC 16 has 2000
hours creep rupture life at 15620F (8500C) for 40ksi (275MPa) operation stress as it can be
seen in Figure 6-3[10]. Oxidation and elongation performances have also been taken into
consideration in material selection.
For 74 rotor blades and a 10% airfoil thickness, disk stress (σd) has been found around
65.3 ksi (450 MPa) for disk shape factor (DSF) of 2.1. Disk stress is allowable by the selected
material since the disk temperature is 200-3500F (100-2000C) lower than the blade average.
69
The cooling effectiveness (Φ) is 0.62 for 15620F (8500C) material and 15080F (8200C)
coolant flow temperatures. As stated by Hess in “Laminated turbine vane design” NASA
report[13], with 6.17% coolant mass flux ratio and 0.62 cooling effectiveness, cooling from
28000F to 11960F (mass average temperature) had been succeeded. Thus, in TJ-1 HPT,
cooling mass flux ratio of 5% is found sufficient for the turbine rotor temperature to drop
from 24710F to 15620F. Nevertheless, NGVs are not exposed to high centrifugal stress that the
rotors are, coolant percentages should be examined for distribution along the turbine for better
engine performance. Further rotor creep rupture life can be procured by ceramic coating that
provides 1040F (400C) decrease on the airfoil surface. That means 1.2-1.3 times more rupture
life than uncoated. However, a cost tradeoff should be done between reducing service time for
coating control and maintenance for rupture.
Figure 6-3 : Creep Rupture Life (h) – Stress (MPa) graph of SC16[10]
6.2 LPT design
In LPT design, as it has been argued in Chapter 6.2, the same tradeoff between one
and two stage has been made again. It has been decided as one stage since the limiting
boundaries are not tight.
LPT design has a similar process with the HPT design, a mere difference is that the
rotational speed is known as a limiting factor because of the fan’s operating conditions which
is 1275 rad/s.
Low-pressure turbine design point parameters are given in Table 6-7:
70
Table 6-7: Low-pressure turbine design point parameters (1.1M/20kft)
τ = 0.85
Pt 4.4 = 118.1 psi
γ = 1.305
п = 0.451
Tt 4.4 = 2259 R
gcCp = 7351 ft2/(s2.0R)
N = 12175 RPM
ṁ 4.4= 73.77lb/s
R = 53.4 ft.lbf/(lbm.0R)
Parameters listed in Table 6-7, 1275 rad/s rotational speed, etL=0.9,M2 = 1.1, M3R =
0.9 and α2 = 600have been used as inputs for TURBN.exe, many design quantities have been
found and listed in Table (6-8, 6-9, 6-10):
Table 6-8: Results of the TURBN.exe for LPT -11h
1m
1t
2h
2m
2t
2Rm
3Rm
3h
3m
3t
Tt
0
2259
2259
2259
2259
2259
2259
2042
2042
1911
1911
1911
T
0
R
2228
2228
2228
1858
1907
1946
1907
1817
1816
1817
1818
Pt
Psia
118.1
118.1
114.6
114.6
114.6
114.6
74.4
70.8
53.3
53.3
53.3
P
Psia
111.4
111.4
111.4
49.6
55.5
60.5
55.5
43
42.9
43
43.1
0.3
0.3
0.3
1.19
1.10
1.03
0.68
0.90
0.584
0.58
0.577
R
M
V
ft/s
671
671
671
2429
2275
2146
1409
1817
1179
1171
1165
U
ft/s
671
671
671
1137
1137
1137
1137
1137
1137
1137
1137
V
ft/s
0
0
0
2147
1970
1820
831
1417
310
279
253
Α
deg
0
0
0
62.09
60
58
15.25
13.76
12.54
Β
deg
Radii
in.
9.63
10.72
11.81
9.85
10.72
11.59
9.84
10.72
11.60
36.15
51.26
10.72
10.72
Table 6-9: Results of the TURBN.exe for LPT -2Hub
Rt = 0.1188
A1 = 117.5 in2
Mean
Rt = 0.2575
A2 = 119 in2
Tip
Rt = 0.3658
A3 = 146.38 in2
Table 6-10: Results of the TURBN.exe for LPT -3Stage-load
Flow
Isentropic
Aspect Ratio
Solidity
Blade number
coefficient
coefficient
efficiency
Vane
Rotor
Vane
Rotor
Vane
Rotor
1.972
0.998
0.908
1
1
0.886
2.077
34
71
71
Stage-load coefficient, flow coefficient and turbine reaction are stated above. Results
are consistent with the limitations in literature while providing an uncooled total-to-total
0.908 turbine efficiency.
6.2.1
Structural analysis of LPT
As it has been in HPT, the same structural assumptions and calculations have been
made in LPT. The only difference is that LPT requires uncooled design. Parameters have been
selected or calculated as; At/Ah=0.7, Aav= 132.7 in2, ω=1275 rad/s and σc=22ksi. The same
material which has been used for HPT, SC16 has also been chosen for LPT. Average total
temperature at the parametric cycle is 17420F. SC16 has 1000h creep rupture life for 21.76ksi
at 17420F. This life is fairly low for a turbine design, but as it was explained in HPT design, a
ceramic thermal coating will drop the temperature of the blades about 100-1200F which gives
3 times more rupture life to the turbine. In conclusion, an uncooled low turbine design is
achievable at the parametric cycle temperatures.
72
7
Nozzle Design
Nozzle is one of the main parts that substantially change the operating conditions of an
engine and has a strong effect on thrust and specific fuel consumption. Therefore, nozzle
design is a very important part of the engine design process. Variable nozzle becomes
prominent since it alters the operating conditions of the engine during off-design operations.
Low corrected mass flow rates cause operating line to get closer to the surge line. This
is an undesired situation since transient operation can cause compressor to surge. Increasing
nozzle throat area balances the engine back pressure and also by increasing the corrected mass
flow rate, it moves the operating line away from surge line.
There are mainly two types of nozzles. One of them is the convergent nozzle which is
usually utilized for the subsonic aircrafts. Other one is the convergent divergent nozzle which
is generally used for the supersonic aircrafts. Main reason for this situation is that supersonic
engines’ operating requirements -exit velocity- are higher than the exit velocities provided by
convergent nozzles.
Engines with afterburner must have variable nozzle throat area for the sake of proper
back pressure control. Thus, nozzle exit area must be increased in order to balance exit total
pressure and back pressure. Besides that, the designed engine is able to operate at both
73
subsonic and supersonic flow regimes which mean completely different nozzle area ratios for
the engine.
In addition, variable nozzle enhances starting performance. Opening the nozzle exit
area reduces the back pressure on the turbines, leading to increased expansion ratio so that
turbines generate more power at lower turbine inlet temperatures. Hence smaller starter can be
employed to provide a lighter engine.
Theoretically, variable convergent divergent nozzle can expand the air till its static
pressure equalizes to the ambient pressure. However in real conditions this may not be
possible. Over-expansion or under-expansion may occur while operating. Over-expanded
nozzles have more pressure losses than under-expanded nozzles[14]. Thus some margin can be
put to use in order to avoid over-expansion since slight under-expansion is tolerable.
Additionally, present design has been decided to have two dimensional (pitch only)
thrust vectoring since it increases maneuverability and decreases SFC at certain flight
points[4]. Moreover thrust increment up to 7% is possible with thrust vectoring.
Designing process includes some decisions that are vital for the engine. Nozzle must
have high performance at all the mission segments as required by a tactical aircraft. Surely
performance is a very important parameter; however production and maintainability costs are
also need to be considered in order to achieve an efficient design. For a tactical aircraft, cost
and maintainability issues are less important than lack of performance. General trends are also
indicating that tactical aircrafts flying over Mach 1.5 require convergent-divergent variable
nozzle[3]. Furthermore the mission varies greatly over different segments. Hence, optimizing
nozzle for two design points, like the geometrically scheduled nozzle design does, is not an
option in order to keep the performance high. All of these lead to geometrically scheduled
nozzle to be omitted from design choices.
Passively scheduled nozzle has also been considered for the present design. This
arrangement employs divergent section flaps which are allowed to move through a range for
the same throat area. Internal pressure of the divergent section provides the necessary force to
move these flaps through their range of possible area ratios[14]. Floating flaps is less complex.
Therefore, light and floating flaps can fulfill the required performance. However there are
certainly some downfalls. Firstly, this type of nozzle significantly reduces the region of flight
envelope which aircraft operates at its off design conditions[3]. Secondly, two dimensional
74
thrust vectoring is not possible in the absence of full control over the divergent section. In the
light of all this reasons, fully variable nozzle is found to be the best solution for performance
needs in spite of being complex and heavy.
Rectangular nozzle design is preferred in order to perform two dimensional thrust
vectoring. Some additional pressure loss has also been anticipated because of the rectangular
shape. Pressure loss of the rectangular nozzle with an aspect ratio of 2, is found 1.2 times
more than the circular nozzle for the same area[2] and included in calculations.
Nozzle designing process includes a tradeoff between dimensions and performance.
High performances can be achieved via increased nozzle dimensions. On the other hand, this
may cause the nozzle to be very long with respect to rest of the engine. In addition, a very
long nozzle contributes substantially to total weight of the engine. In order to prevent this,
performance must be compromised without exceeding limits.
Design process of nozzle begins with calculating the pressure ratios, throat areas and
exit Mach numbers for all the mission segments using off-design analyses. Using these data
and assuming perfect expansion, area ratios have been found.
For primary nozzle half angle, calculations have been carried out with minimum flow
rate and for secondary nozzle half-angle, maximum pressure ratio is used in order to find
largest values of nozzle half angles. Discharge coefficient has been assumed to be 0.94 in
order to reduce total nozzle length. By assuming discharge coefficient, primary nozzle angle is
determined as 31.50 at minimum flow rate. Using exit Mach numbers, which have been found
from area ratios and pressure loss, the velocity coefficient is found as 0.997. Largest
secondary nozzle half angle is obtained as 130 at maximum flow rate. Primary nozzle length is
0.186 meters and secondary nozzle length is obtained as 0.519 meters.
Nozzle dimensions are obtained from nozzle inlet radius and calculated angles.
Pressure loss across the nozzle has also been assumed as 0.99 for circular nozzle which equals
to 0.988 for the present design.
All calculations have been performed for circular nozzle. After that, using the primary
and secondary nozzle lengths obtained from circular nozzle calculations, cross section has
been changed into rectangular shape with an aspect ratio of 2 for throat at maximum power.
Side walls of nozzle are not allowed to move and width of the nozzle is constant for
rectangular cross sectioned zone. Nozzle area ratio is limited to 2.285 because of enormously
75
increasing exit dimensions. This area ratio corresponds to Mach 1.6, mission segment. Higher
Mach numbers require larger area ratios therefore operations with higher Mach numbers will
not be perfectly expanded anymore. This situation is shown in the graphics below. Input and
output values are also given below; nozzle dimensions have been drawn with CATIA and are
given in APPENDIX -A-;
Table 7-1: Input values of nozzle calculations
1.6 Mach Pressure Ratio
Minimum Flow
Rate
Mass Flow Rate
44.59
17.48
Pressure Loss
0.988
0.988
Throat Mach Number
1.00
1.00
Ambient Mach Number
1.60
1.40
A9/A8
2.28
1.51
Discharge Coefficient
0.94
Table 7-2: Output values of nozzle calculations
1.6 Mach Pressure Ratio
Minimum Flow
Rate
Exit Mach Number (ideal)
2.34
1.90
Exit Mach Number (real)
2.33
1.89
Velocity Coefficient
0.9973
0.9954
Nozzle Primary Half Angle
20.75
31.50
Nozzle Secondary Half Angle
13.00
4.79
Throat Height
0.29
0.10
Throat Width
0.59
0.59
Exit Height
0.67
0.15
Exit Width
0.59
0.59
Primary Nozzle Length
0.19
Secondary Nozzle Length
0.52
Total Nozzle Length
0.70
76
8
Electricity
Most aircrafts use two separate systems for electric generating and engine starting.
Since weight is a strong factor which affects aircraft performance, lighter engine will have
better performance under same circumstances. Combining these systems saves weight as well
as decreasing the total volume that whole system needs. This system type is called integral (or
integrated) starter generator and chosen for TJ-1 design.
Aircrafts generate power using different scenarios. The most common schemes include
constant frequency, variable speed constant frequency and variable frequency power
generating. All of the techniques given above are evaluated for TJ-1 design.
Traditionally, aircrafts employ wound field synchronous machines in order to generate
400 Hz constant frequency 3 phase alternative current[1]. This system is known as constant
speed drive (CSD). A gearbox and a shaft which connects the main engine shafts to an
accessory gearbox is needed for this arrangement in order to reduce speed. This naturally
causes total weight to increase. Aside from the weight issue, reliability is also a considerable
matter if the complexity of the system is considered. Therefore, operational costs increase
dramatically as gearbox has to be checked before every flight[2].
77
Variable speed constant frequency consists of two different techniques. One of these
techniques is known as DC link system which converts variable frequency alternative current
to direct form and then converts it to 3 phase 400 Hz, 115 V alternative current again. The
other one utilizes a cycloconverter that converts variable frequency alternative current to
constant 400 Hz, 115 V alternative current. DC link is generally preferred for its simplicity
and reliability. On the other hand, cycloconverters are more efficient even though it requires a
fixed turns gearbox and sophisticated control[2].
Further investigating, variable frequency is another option that comes to mind. This
arrangement can extract power directly from the main shafts which the rotor disks are placed
on. Engine angular speed varies over a large margin when engine operates through different
mission segments. As a consequence of the situation, frequency oscillates in a wide range.
However this is merely a problem when omitting cumbersome gearbox and shaft are
considered. System is less complex and more reliable without shaft and gearbox which
significantly decreases operational costs and weight. On the other hand, there are some
downfalls of the system. Placing the generator inside the engine hub will most likely increase
the need for cooling. Moreover engine structure must be stiffened if the rotor of the generator
is only supported by main engine bearings because of the small air gap requirement for
generator. Thus, serious alterations must be done for engine pylons[2].
All of the schemes that stated before are analyzed to find the most suitable system for
TJ-1 design. Amongst all of the designs, variable frequency is the most suitable choice for
integrated starter generator design since it has the capability of meeting the high power
requirements in a wide rpm range. Thus variable frequency generator scheme has been
selected for TJ-1.
Besides, generator type is also a critical issue for power extraction. Reliability, ease of
operability, and production costs are highly dependent on generator type. For TJ-1 design,
induction, synchronous, switched reluctance and permanent magnet generation types are
considered because of their reliability and robustness.
Induction generators are usually used for the small applications such as cars or APU.
Their reliability and robustness are not deniable. However power density of the induction
generators is not sufficient for TJ-1.
78
Permanent magnet generators are also examined in order to obtain the feasibility for
TJ1. Besides the many advantages such as high efficiency, high volumetric, gravimetric
efficiency and ease of cooling, permanent magnets are also very reliable. However, high
temperature levels are intolerable for this type of generator[3]. Hence permanent magnet
generator is elected from the design choices.
Even though synchronous generators dominate today’s aircraft engines, they need
external excitation which reduces efficiency. This configuration is not preferred because of
the external excitation requirement.
Nevertheless, general trends of today have a tendency to use switched reluctance
generators because of their simple, robust structure[3]. Switched reluctance generators also
operate in a wide range of rpms which makes them preferable for integrated starter generator
design. That is why switched reluctance type starter generator is used for TJ-1 design.
Cranking is also an important concern for starting performance of the engine. Usually jet
engines crank the high pressure compressor instead of low pressure compressor. Because,
inertia of the LPC is larger, starting time increases which is an undesirable situation for
tactical aircrafts[3]. Moreover cranking LPC creates more pressure loss than cranking HPC.
Starter generator system, mentioned above, is placed in engine hub, between high
pressure and low pressure compressors. Main reason of this is, to benefit from counter
rotating shafts as much as possible. Additionally, starter generator must be placed close to the
HPC in order to eliminate additional link elements since it cranks the HPC.
The system consists of two parts; first one is a 70hp starter/generator switched reluctance
machine whose stator is connected to the casing and rotor connected to the high-pressure
spool. The second one is a 300hp generator whose stator is connected to the high-pressure
spool and stator is connected to the low-pressure spool. System can be seen in the Figure 8-1.
Red shaft symbolizes the HPC spool and yellow shaft stands for LPC spool.
79
Figure 8-1 : Integrated started generated system
Pink and orange surfaces are the stator and the rotor of starter/generator (S/G)
respectively. S/G’s first responsibility is to provide 150 Nm constant torque on the HP-spool
for starting the engine to 25% of its maximum speed (5000 RPM). The second responsibility
of the S/G is to generate 67 hp for sub-systems throughout the mission while the engine is
self-sustaining its idle RPM (see Figure 8-2).
Figure 8-2 : Starting process
80
Blue and brown surfaces are the stator and the rotor of a main variable frequency
generator which provides 300 hp when needed. As one can see orange part is extend of high
pressure spool in order to crank the HP-spool only. It avoids electromagnetic interaction with
the LP-spool which offers better startup performance. The blue stator uses, the coils (brown
part) which are attached to LP-spool, as its rotor. Thus, it favors the counter-rotating
mechanism by obtaining 31500 relative RPM at full power and generates 300 hp from a
compacter electrical generator. A converter circuit and a control program are necessary to
manage the variable frequencies and shutting down the generator-only part for starting
process.
.
81
9
Appendix
82
APPENDIX -A- Layout of nozzle
83
APPENDIX -B- Layout of inlet
84
APPENDIX -C- Cross sections of LPC and HPC respectively
85
APPENDIX -D- Cross sections of LPT and HPT respectively
86
Appendix E – CAD Drawings

Inlet

Fan
87

88

Burner
89

High Pressure Turbine
90

Afterburner
91

Nozzle
92

Complete CAD render of design
93
94
References
[1] Philip P. Walsh, Paul Fletcher, 2004, Gas Turbine Performance
[2] Mattingle, Heiser, Pratt, AIAA Engine Design, 2004
[3] Raymer, D. P. (1992). Aircraft Design: A Concept Approach. Washington: American
Institue of Aeronautics and Astronautics, Inc.
[4] Official web site of Lockheed
http://www.lockheedmartin.com/content/dam/lockheed/data/aero/documents/f35/collateral/f35_ctol_a11-34324f001.pdf Retrieved 03.03.2012
[5] Whitford R, “Design for Air Combat”, Janes, London, 1987
[6] Goldsmith E. L., Seddon J, Practical Intake Aerodynamic Design, AIAA Education series,
1993
[7] Crosthwait, E.L, Kennon I.G et al, “Preliminary design Methodology for Air-Indutction
Systems”, Technical Report SEG-TR-67-1, Systems Engineering Group, 1967
[8] Mattingley, Elements of propulsion: Gas turbines and rockets.
[9] Dixon, S. L., & Hall, C. A. (2010).Fluid mechanics and thermodynamics of
turbomachinery.
[10] Bachelet, E. (1990).High temperature materials for power engineering.
[11] TURBN.exe(2005). “Multistage Axial-Flow Turbine Design V5.21”
[12] Moffitt, T. P.; et al. (1980).Design and cold-air test of single-stage uncooled core turbine
with high work output.NASA TP-1680.http://ntrs.nasa.gov/ archive/nasa/casi.ntrs.Nasa.gov
/19800016842_1980016842.pdf Retrieved 03.03.2012
[13] Hess, W. G. (1979).Laminated turbine vane design and fabrication.NASA CR159655.<http: //ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19790025033_1979025033.pdf>
Retrieved 03.03.2012
[14] http://www.spiritech.cc/images/AIAA-2004-3923_Nozzle%20Design%20Criteria.pdf
[15] http://ftp.rta.nato.int/public//PubFullText/RTO/MP/RTO-MP-051///MP-051-PSF-11.pdf
[16] Sobester, A., Scanlan, J. et al, “Conceptual design of UAV airframes using a generic
geometry service”, Report AIAA 2005-7079, University of Southampton, 2005.

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