Isı Bilimi ve Tekniği Dergisi, 35, 2, 53-62, 2015

J. of Thermal Science and Technology

©2015 TIBTD Printed in Turkey
ISSN 1300-3615

SIZING OF A TURBOPROP UNMANNED AIR VEHICLE AND ITS PROPULSION

SYSTEM
Ali DİNÇ
Tusaş Engine Industries Inc., Çevreyolu No: 356, 26003 Eskişehir, Turkey, alidinc@yahoo.com

(Geliş Tarihi: 29.05.2014, Kabul Tarihi: 01.03.2015)

Abstract: In this study, a genuine computer code for sizing of an unmanned aerial vehicle (UAV) and its turboprop
engine by analytical method is developed. Payload and fuel weights are primary factors affecting UAV size and weight
(which need to be fulfilled in terms of flight performance parameters, e.g. lift and drag). The parameters within mission
profile such as altitude, speed of aircraft are keys for selecting engine type. Engine specific fuel consumption (SFC)
defines the total fuel amount to be stored and carried during the flight, which affects the general dimensions and the
gross weight of the aircraft. Some engine parameters namely, compressor pressure ratio and turbine inlet temperature,
have direct influence on SFC. Turboprop engine subroutine code developed in this study is within ±1% agreement with
commercial engine cycle analysis software “GasTurb” for shaft power, propeller thrust and SFC etc. values, at all mission
points of UAV. Calculated weight, size, endurance values of UAV are seen to be close to literature values. Literature values
indicate some inconsistencies especially for endurance and empty weight of aircraft. Considering most dependable
references and logical combinations of inputs, error in calculation of UAV weight and size is about ±1,5% and for
maximum operating altitude is around ±3%. Moreover, calculated endurance values are within range of literature values.
Keywords: unmanned air vehicle design, aircraft sizing, propulsion, turboprop, gas turbine engine, cycle analysis

BİR TURBOPROP İNSANSIZ HAVA ARACININ VE İTKİ SİSTEMİNİN

BOYUTLANDIRILMASI
Özet: Bu çalışmada, bir insansız hava aracının (İHA) ve turboprop motorunun analitik yöntemle boyutlandırılması
için özgün bir bilgisayar yazılımı geliştirilmiştir. Faydalı yük ve yakıt ağırlıkları, İHA boyut ve ağırlığını (taşıma ve
sürükleme kuvvetleri gibi temel uçuş performans parametreleri açısından karşılanması gereken) etkileyen başlıca
faktörlerdir. Görev profilindeki irtifa, uçak hızı gibi parametreler, motor tipini seçmek için anahtardır. Motor özgül
yakıt tüketimi (SFC) uçağın genel boyutlarını ve toplam ağırlığını etkileyen ve uçuş sırasında depolanıp taşınacak
olan toplam yakıt miktarını tanımlar. Kompresör basınç oranı ve türbin giriş sıcaklığı gibi bazı motor
parametrelerinin, SFC üzerinde doğrudan etkisi vardır. Bu çalışma kapsamında geliştirilen turboprop motor alt
yazılımı İHA’nın tüm görev noktalarında mil gücü, pervane itkisi, SFC vb. değerler için ticari motor çevrim analiz
yazılımı olan "GasTurb" ile ±%1 uyum içerisindedir. İHA’nın hesaplanan ağırlık, boyut, dayanıklılık değerlerinin
literatür değerlerine yakın olduğu görülmüştür. Literatür değerleri, özellikle azami uçuş süresi (endurance) ve uçak
boş ağırlığı için tutarsızlıklar göstermektedir. En güvenilir referanslar ve girdilerin mantıksal kombinasyonları göz
önüne alındığında, İHA ağırlığı ile boyutundaki hesaplama hatası yaklaşık ±%1,5 ve azami uçuş irtifası ±%3 olarak
hesaplanmıştır. Ayrıca, hesaplanan azami uçuş süresi değerleri literatürde verilen aralıklardadır.
Anahtar Kelimeler: insansız hava aracı tasarımı, uçak boyutlandırma, itki, turboprop, gaz türbin motoru, çevrim analizi

NOMENCLATURE Cfht composite fudge factor for horizontal tail

[dimensionless]
ai speed of sound at engine station “i” [m/s] Cfvt composite fudge factor for vertical tail
A7 exhaust nozzle area [m2] [dimensionless]
A7,c nozzle exit area for choking condition [m2] Cfw composite fudge factor for wing [dimensionless]
AIAA American Institute of Aeronautics and CL lift coefficient [dimensionless]
Astronautics CLmax maximum lift coefficient [dimensionless]
AR aspect ratio of wing [dimensionless] CL, TO max. lift coefficient for take of [dimensionless]
c chord length [m] CPC specific heat for compressor [kJ/kg K]
C specific fuel consumption [kg/(kN.h)] CPT specific heat for turbine [kJ/kg K]
Ccorr engine airflow correction factor [dimensionless] CPW power coefficient of propeller [dimensionless]
CF thrust coefficient of propeller [dimensionless] d propeller diameter [m]
Cfe surface friction coefficient [dimensionless] dcl climb distance (in horizontal) [km]
Cff composite fudge factor for fuselage [dimensionless] D drag force [N]
e Oswald efficiency for wing [dimensionless] tB block time [h]
E Endurance [h] tCL climb time [h]
EPW Equivalent shaft power [kW] tCR cruise time [h]
ESFC Equivalent power specific fuel consumption tD descent time [h]
[kg/(kW.h)] tF flight time [h]
f fuel air ratio [kg/kg] tGR ground maneuver time [h]
ffus.vol fuselage volume utilization factor [m3/m3] tidle idle time before takeoff [h]
ffus.vol.m fuselage volume margin [m3/m3] Tinst installed thrust (propeller+jet)
flg landing gear weight ratio [kg/kg] tloi loiter time [h]
fmisc other misc. equipment weight ratio [kg/kg] tpow time of maximum power for take-off [min]
fnac.wet nacelle wet area factor [m2/m2] ttot total mission time [h]
Fnet net total thrust [N] twait maximum waiting time spent on air before
fresfuel landing reserve fuel ratio [kg/kg] landing [min]
fsys avionics and flight systems weight ratio [kg/kg] Tamb ambient static temperature [K]
fth.ins thrust installation loss factor [N/N] Tbleed Compressor middle stage total temperature
fw.eng engine installation weight factor [kg/kg] for bleed [K]
fwm empty weight margin [kg/kg] Ti total temperature at station “i” [K]
FA nozzle thrust [N] TSFC thrust specific fuel consumption [gr/(kN.s)]
GC Power required for compressor [kW] Tuninst uninstalled thrust (propeller+jet) [N]
GHPT power generated by high pressure turbine [kW] ufus unit weight for aircraft fuselage and nacelle
GLPT power generated by low pressure turbine [kW] [kg/m2]
h aircraft flight altitude [km] uhtail unit weight for the horizontal tail [kg/m2]
h0 altitude for takeoff (point 1) [km] uvtail unit weight for the vertical tail [kg/m2]
h10 altitude for landing (point 10) [km] uwing unit weight for wings [kg/m2]
h34 altitude at the end of climb or at the start of UAS unmanned air system
cruise (points 3−4) [km] UAV unmanned air vehicle
h56 altitude at the end of cruise or at the start of V0 aircraft flight velocity [km/h]
loiter (points 5−6) [km] Vclimb climb speed [km/h]
h78 altitude at the end of loiter or at the start of Vstall stall speed [km/h]
cruise (points 7−8) [km] Vcr maximum cruise speed [km/h]
h9 altitude at the end of return cruise or at the start Vfuel fuel volume [m3]
of descent (point 9) [km] Vwing wing volume [m3]
J advance ratio [dimensionless] Vloi max. speed (Mach number) for loiter and climb
L lift [N] Vstall stall speed [km/h]
L/D lift to drag ratio [N/N] Vref reference speed for takeoff [km/h]
Lfus total fuselage length [m] Vwind headwind [km/h]
Lfus.b fuselage back side cone length [m] wi engine airflow at station “i” [kg/s]
Lnac nacelle length [m] wf fuel flow [kg/s]
M0 flight Mach number [dimensionless] W/S wing loading [kg/m2]
n propeller speed [rev/s] W aircraft weight during flight stages [kg]
neng number of engines W0 maximum aircraft gross weight [kg]
nfus number of fuselage of aircraft Waf airframe weight [kg]
Ncd nozzle discharge coefficient We aircraft empty weight [kg]
Ncx nozzle thrust coefficient Weng engine weight [kg]
Pi total pressure at station “i” [kPa] Wf total fuel weight [kg]
Pamb ambient static pressure [kPa] Wfe external fuel weight [kg]
PLAidle engine idle throttle percent [%] Wfi internal fuel weight [kg]
PSFC power specific fuel consumption [kg/(kW.h)] Wfus fuselage weight [kg]
PW uninstalled engine shaft power [kW] Wh.tail horizontal tail weight [kg]
R range [km] Wlg landing gear weight [kg]
RASCE Rapid Air System Concept Exploration tool Wmisc miscellaneous weight [kg]
RCL climb distance [km] Wnac nacelle weight [kg]
RCR cruise range [km] WPe external payload weight [kg]
RD descent distance [km] WPi internal payload weight [kg]
RF range [km] WPL total payload weight [kg]
Rop operation radius of UAV [km] Wsys systems weight [kg]
S total wing area [m2] Wv.tail vertical tail weight [kg]
Sh horizontal tail area [m2]
Sv vertical tail area [m2] Greek Symbols
t/c thickness to chord ratio of wing [m/m] ∆Pb total pressure loss in combustion chamber
tave average thickness of wing airfoil [m] [kPa/kPa]
tmax maximum thickness of wing airfoil [m]

54
∆Pin intake total pressure loss [dimensionless] The suitability of UAVs in “dull, dirty and dangerous”
∆Pj.pipe total pressure loss of the low pressure turbine missions, the increasing success of UAVs in service and
jet pipe [dimensionless] demonstration. Therefore, UAV is an aircraft which can
∆Pt.duct total pressure loss in the duct between high and autonomously fly or can be remotely controlled to perform
low pressure turbines [kPa/kPa] a specific mission without flying crew inside. Unmanned
Øeng engine diameter [m] air vehicles have a lot of varieties including micro, tactical,
Øfus fuselage diameter [m] strategic and combat types (Chaput, 2004).
Ønac nacelle diameter [m]
Ønac nacelle diameter [m] An over-simplistic view of an unmanned aircraft is that it
Øprop propeller diameter [m] is an aircraft with its aircrew removed and replaced by a
β1 compressor middle stage bleed air ratio computer system and a radio-link. In reality it is more
[dimensionless] complex than that, and the aircraft must be properly
γ specific heat ratio [dimensionless] designed, from the beginning, without aircrew and their
δ atmospheric pressure ratio [=Pamb/PSL] accommodation, etc. (Austin, 2010).
ε1 compressor middle stage air extraction ratio for
low pressure (LPT) duct cooling Design of a UAV is similar to that of a manned aircraft to
ε2a cooling air ratio for high pressure turbine some extent. A lot of tools and methods were developed
ε2b cooling air ratio for low pressure turbine duct in the past for aircraft design. Some of the well-known
ε3 cooling air ratio nozzle guide vanes (NGV) text books published in literature belong to Raymer
η2 compressor isentropic efficiency (1999), Roskam (1990) and Nicolai (2010). Sizing codes
η416 high pressure turbine isentropic efficiency were generated to speed up the conceptual design phase
η46 low pressure turbine isentropic efficiency using the developed aircraft design methodology. Some
ηb combustor efficiency are simple in-house parametric sizing codes and some are
ηC,pol compressor polytrophic efficiency sophisticated commercial programs (Internet, 2014).
ηm mechanical efficiency Parametric codes are typically used for concept
ηp propeller efficiency (constant or function of J) exploration. One example is Rapid Air System Concept
ηprop,d propeller dynamic efficiency Exploration (RASCE) tool which is a physics-based,
ηprop,s propeller static efficiency unmanned air system conceptual level design and
ηT,pol turbine polytrophic efficiency analysis system. RASCE is originally developed as an
θ atmospheric temperature ratio [=T amb/TSL] educational tool to support undergraduate student
λ tip to hub chord length ratio of wing exploration (Chaput, 2010).
ΠC compressor total pressure ratio
ΠN exhaust nozzle pressure ratio (design input) For aircraft design, mission profile definition consists of a
ΠT turbine total pressure ratio very important set of parameters. A mission profile is a
ρeng uninstalled engine density scenario that is required to establish the weight, fuel,
ρfuel fuel density payload, range, speed, flight altitude, loiter and any other
ρlg landing gear density operations that the aircraft must be able to accomplish.
ρpl payload density The mission requirements are specific to the type of the
ρsys systems density aircraft (Curtis et al., 2009). In other words, mission
Ωprop propeller speed [rpm] profile is a scheme of aircraft’s flight segments and
detailed description of aircraft activities in flight. For this
INTRODUCTION reason, it is very important for the design of the aircraft.
Figure 1a, shows a typical reconnaissance unmanned air
It has been more than a century since Wright brothers vehicle mission profile. A representative mission profile
realized the first controlled, powered and sustained is assumed for this study and shown in Figure 1b. A total
heavier-than-air human flight in 1903. Aviation industry is of 11 mission points are defined from start to end.
one of the emerging high-tech fields, starting from the Descriptions of all those 11 points are given in Table 1.
period of First World War and studies in this field indicate
that importance of the aviation sector will increase in the
coming years (Genç et al., 2008). Many types of aircraft
with different complexity have been designed and built
over the past century as the technology advanced.

In addition to manned aerial vehicles, unmanned or

uninhabited air vehicles (UAV) have been developed as
well. As defined in the AIAA Committee of Standards’,
“Lexicon of UAV, ROA (remotely operated aircraft)
Terminology, a UAV is “an aircraft which is designed or
modified, not to carry a human pilot and is operated Figure 1a. A typical reconnaissance unmanned air vehicle
through electronic input initiated by the flight controller or mission profile (Federation of American Scientists, 2014).
by an onboard autonomous flight management control
system that does not require flight controller intervention”.

55
 AIRCRAFT PERFORMANCE

Although intensive test data and accumulated experience

are needed for accurate estimates of performance in the
later stages of design, the equations used in this study will
be simple, basic and are sufficient for the preliminary
design stage. Flight durations and range calculations
(Buğdaycı, 1980) are given in equations section.

PROPULSION

Figure 1b. A representative mission profile for Predator B. Different propulsion systems are used to power aircraft to
fly. Piston, electric, gas turbine, ramjet, scramjet and even
Table 1. Predator B assumed mission profile. nuclear engines are used in commercial and military
Mission aviation. Turbofan, turbojet, turboshaft and turboprop
Description Mission Altitude
Point engines are in the group of gas turbine engines. Gas
0 Idle 0 turbine engines, based on terrestrial and aeronautical, are
1 Taxi and take off 0 used for a wide range of power generation applications,
End of take-off, start of including aerospace, cogeneration, power plants and the
2 0
climb like (İlbaş and Türkmen, 2012).
3 End of climb 6,096 km (20 kft)
4 Start of cruise climb 6,096 km (20 kft) The general energy supply and environmental situation
5 End of cruise climb 9,144 km (30 kft) requires an improved utilization of energy sources.
6 Start of loiter 9,144 km (30 kft) Therefore, the complexity of power-generating units has
7 End of loiter 9,144 km (30 kft)
increased considerably. This requires thermodynamic
calculations of high accuracy (Şahin et al., 2011).
8 Start of cruise descent 9,144 km (30 kft)
9 End of cruise descent 6,096 km (20 kft) As explained by Chaput (2010), satisfying propulsion
10 Start descent 6,096 km (20 kft) data requirements can be a problem, particularly for
11 Landing 0 students. Air vehicle performance codes typically require
tabular inputs of installed thrust and fuel flow. Generating
the installed data can be time consuming and/or involve
SIZING OF UNMANNED AIR VEHICLE use of proprietary engine company codes. One solution to
the problem is to use an integrated multi-discipline
For the sizing of UAV, general aircraft weight formulae parametric design system that has the fidelity of a
can be used by omitting crew and passenger weights conceptual point design and analysis system and the
(Raymer, 1999). In Figures 2a-2b, a simplified geometry flexibility of a parametric sizing code.
is created genuinely for this study by inspiration from real
UAVs such as “Predator B”. The UAV in Figures 2a-2b Structuring an aero-thermal model of a gas turbine engine
consists of basic cylinder, sphere, cone etc. shapes which is the first step to simulate engine performance in a
make calculations (e.g. volume, area) simpler. The dynamic manner (Uzol, 2011). The object of parametric
volume and area values for UAV components (wing, cycle analysis is to obtain estimates of the performance
fuselage etc.) are used for weight estimation. Related parameters (power/thrust and specific fuel consumption)
component weight formulae and densities are taken from in terms of design limitations (such as maximum
Chaput (2004) and Raymer (1999). allowable turbine temperature and attainable component
efficiencies), the flight conditions (the ambient pressure,
temperature and Mach number) and design choices (e.g.
compressor pressure ratio) (Mattingly et al., 2002).
Parametric analysis determines the engine performance
under different flight conditions, different design choices
(e.g. compressor pressure ratio) and design constraints
(e.g. burner exit temperature); whereas the performance
analysis allows the calculation of performance for
different flight conditions and power level of the engine
with determined specific values (Turan et al., 2008).
Figure 2a. A simplified geometry UAV, 3-dimensional view.
"Predator-B" UAV uses TPE331−10 turboprop engine
(Honeywell, 2014). Therefore in this section, a detailed
turboprop engine on-design cycle analysis is studied,
although in RASCE tool developed by Chaput (2010),
turboprop performance is simply modeled as a turbofan
Figure 2b. A simplified geometry UAV, side view.
of very high bypass ratio.

56
In Figure 3, schematic of a two spool turboprop engine is match the overall performance parameters. For example,
shown. Main components are intake (A), compressor (B), engine parameters like ηC,pol , ηT,pol , ηb , ηm , ε1 , ε2a , ε2b , ε3
burner (C), high pressure turbine (D), low pressure turbine etc. are best guesses to match the power and SFC values.
(E) and finally exhaust nozzle (F). Methodology in Walsh
and Fletcher (2004) is used for turboprop cycle analysis in For a different run, if any of the inputs in Table 3 is
general, with the exception that similar to method in changed, then both aircraft and engine size and
Kurzke (2007), burner exit temperature is taken as design performance parameters change. For example, if payload is
parameter instead of stator outlet temperature. changed as an input, aircraft size changes to carry this
additional weight by increasing the wing area (for the
A two spool turboprop engine design point calculations given wing loading). Then engine size change as a result
are given in Equations section. for new size aircraft needs, after iterations. Similarly, if
fuel weight is changed as an input, aircraft size changes for
fuel storage and similarly engine size and performance
(endurance, range etc.) change as well, again after
automatic iterations.

There are some checkpoints in the code and gives

warning messages if any of the following illogical cases
occur. Then inputs should be reviewed:

- Too low thrust to balance drag, for the input altitudes

(decrease altitude or increase engine power)
Figure 3. Schematic view of two spool turboprop engine
(Walsh and Fletcher, 2004).
- Too low input velocity to produce required lift (increase
cruise or loiter Mach number)
A CODE FOR SIZING UAV AND ITS ENGINE - Complex numbers in results (check engine and aircraft
parameters)
Putting together all information (aircraft and engine - Negative numbers in flight time segments, climb rate
sizing, performance equations), a genuine code for etc. (check engine and aircraft parameters)
sizing a turboprop UAV is written for both aircraft and
its engine. Some empirical data and correlations are also
used for some parameters such as weight estimation of
engine, UAV components and systems. The code is
about 1000 lines including engine subroutine.

In Figure 4, flowchart of UAV and Turboprop engine

sizing code is given. Code starts with an initial size for
aircraft and engine, but updates (decrease or increase)
the size of both aircraft and engine through iterations for
the given inputs. At the end it converges to a final size,
weight, performance of engine and UAV.

More than 100 parameters (related with mission profile

aircraft and engine design options/selection/limits etc.)
are used in the input file (see Table 3).

Conf. 1 input values are given in Table 3. There are 3

configurations of UAV as follows:

- Conf.1: With external fuel tanks and 1/3 of weapons

- Conf.2: Clean configuration (no external fuel tanks or Figure 4. Flowchart of UAV and turboprop engine sizing code.
weapons)
- Conf.3: With full weapons (no external fuel tanks) RESULTS AND DISCUSSION

In Table 3, input data and configuration information are For an assumed Predator-B configuration (per inputs in
collected from literature (Chaput, 2004; Raymer, 1999; Table 3), code is run and results are given in Table 2 and 4.
Honeywell, 2014; General Atomics Aeronautical Systems, Literature and calculated values are compared and are
2014; Defense Technical Information Center, 2003; found to be close. Literature values are somehow
Executive Aircraft Maintenance, 2014) and sub grouped inconsistent especially for the endurance, empty weight,
into mission, aircraft weight, aircraft aero, aircraft internal fuel storage amount which does not make clear
components and finally engine specific segments. Some targets for the computer model or code to match, mainly
data are not available in literature and are best guessed to due to official detail data is limited and there are a lot of

57
versions developed in time for the Predator B and also close to benchmark model values (Predator B). This
external payload and fuel amount may vary in kind of fast tools can be beneficial especially in the
configurations of UAV. However most dependable conceptual design iterations before going into detailed
references and logical combinations are checked and design. The code has a detailed engine cycle analysis
presented. In Table 2, only overall parameters are given as subroutine and can be used both aircraft designers and
a summary and comparison is made with literature values. engine designers. Aircraft designers can use this
Mathematical model is robust and for 3 different detailed engine subroutine for propulsion calculations
configurations calculated error for UAV weight and size is (when they need a so-called rubber engine) and engine
about ±1,5%, maximum operating altitude is around ±3% designers can use it to start a new engine design by
and endurance is within given literature intervals. seeing direct effects of engine design parameters on
aircraft sizing and performance. Similarly students can
In Table 4, selected resulting parameters are listed after use this kind of educational tool to see the effect of an
calculations. Aircraft sizing (e.g. weight, wing area) and engine or aircraft parameter on the overall UAV size
performance (endurance, range, drag etc.) parameters and performance. Code can be improved by adding
are calculated. In addition, engine related parameters empirical data, more detailed aero and structural models
such as power, SFC and all pressure and temperature which are specific to companies in the extent of their
values at different stations are given. Those aircraft and experience, tested products and matured technologies.
engine related parameters are close to benchmark model
Predator B and its engine TPE331−10. Table 3. Input parameters and values for Conf.1 (Chaput, 2004;
Raymer, 1999; Honeywell, 2014; Defense Technical Information
Table 2. Predator-B data comparison (General Atomics Center, 2003; Executive Aircraft Maintenance, 2014)
Aeronautical Systems, 2014; UK Royal Air Force, 2014; Mission Parameters Value
Defense Technical Information Center, 2003; Department of h0 0 [km]
the Air Force Headquarters Air Force Civil Engineer Support h10 0
Agency, 2009; Department of Defense, 2009). h34 6,096
Literature Calculated Error h56 9,144
Parameter
Value value % h78 9,144
4763 kg 4770 kg 0,15% h9 6,096
Max Gross
(Conf. 1-3)
Takeoff Rop 472,3
3454 kg 3403 kg 1,5%
Weight tidle 20
(Conf. 2)
1863- 2227 kg PLAidle 10 %
Empty Weight 2184 kg in range tpow 5
(Conf. 1-2-3)
Dimension: twait 20
20,11 m 20,13 m fresfuel 0,03
Wingspan
10,97 m 10,97 m Weight Parameters Value
Length 0,1%
0,91−1,13 m 1,02 m WPi 363
Fuselage
(Conf. 1-2-3)
diameter WPe 454
Max Wfi 907
15,24 km
Operating 14,78 km 3,1% Wfe 862
(Conf. 2)
Altitude flg 0,043
Maximum 27-32 (Conf. 1) 31,28 fsys 0,12
Endurance 20-24 (Conf. 2) 22,26 in range fmisc 0,02
(hours) 12-14 (Conf. 3) 12,96 fwm 0,03
Engine ufus 19
704 kW 702,2 kW
Parameter: uwing 32
0,325 kg/kW/h 0,3258 0,25%
EPW
(Conf. 1-2-3) kg/kW/h uhtail 18
ESFC
uvtail 18
ρfuel 801
Additionally, calculated engine parameters by code are
ρeng 449
compared with a commercial engine cycle analysis software
ρlg 256
GasTurb (Kurzke, 2007) and maximum differences are ρsys 256
given in Table 5. For each mission point (see Table 1), code ρpl 272
makes the on-design calculations and results are within ±1% Cff 0,9
agreement with GasTurb as can be seen in Table 5. Cfw 0,85
Cfht 0,83
CONCLUSION Cfvt 0,83
Aerodynamic Parameters Value
As the unmanned air vehicles become more common Cfe 0,0041
and useful systems in both military and civil area, fast CLmax 1,8
and efficient design tools are needed in every stage of CL, TO 1,49
development process. The genuine computer code e 0,75
developed in this study can be used in the preliminary Vwind 0
stage of design for initial sizing of aircraft. Input Vclimb/Vstall 1,25
parameters are calibrated and the results obtained are

58
Vloiter/Vstall 1,1 ηT,pol 0,86
Vcr 370 ηb 0,999
Vloi 0,4 ηm 0,995
Vref 185,2 ∆Pin 0
Airframe Parameters Value ∆Pb 0,03
nfus 1 ∆Pt.duct 0,025
Lfus/Øfus 10,75 ∆Pj.pipe 0,02
Lfus.b/Lfus 0,1 Ncd 1
ffus.vol 0,7 Ncx 0,99
ffus.vol.m 1,3 ΠN 1,03
Wing Parameters Value β1 0
AR 17,92 ε1 0
W/S 210,8 ε2a 0,05
λ 0,444 ε2b 0
t/c 0,13 ε3 0,05
tave/tmax 0,6 Øprop 2,8
Vfuel /Vwing 0,5 Ωprop 1591
Tail Parameters Value ηprop,d 0,8
Sh/S 0,08 ηprop,s 0,7
Sv/S 0,135
Nacelle Parameters Value Table 5. Turboprop engine cycle calculations comparison.
Lnac/Ønac 2,7 Engine Parameter Max difference from GasTurb
Ønac/Øeng 1,25 PW 0,37%
fnac.wet 0,5 EPW 0,10%
Lnac/Lfus 0,5 ESFC 0,51%
Engine parameters Value TSFC 0,85%
neng 1 Fnet 0,46%
fth.ins 0,9
Pi 0,11%
fw.eng 1,3
Ti 0,13%
Fnet/W0 0,3359 [kgf/kg]
T4 1368,7
ΠC 10,37
ηC,pol 0,795

Table 4. Result parameters and values.

Aircraft output parameters
W0 4770
S 22,61
R 9308
tloi 28,96
tCR 2,32
ttot 32,20
E 31,28
dcl 85,20
Mission points #1 #2 #3 #4 #5 #6 #7 #8 #9 #10
h 0,000 0,000 6,096 6,096 9,144 9,144 9,144 9,144 6,096 0,000
M0 0,151 0,182 0,268 0,326 0,339 0,329 0,270 0,339 0,326 0,145
CL 1,274 0,878 0,870 0,590 0,825 0,878 0,878 0,556 0,381 0,878
V0 185,2 222,6 304,9 370,4 370,4 359,0 294,7 370,4 370,4 177,8
L/D 22,46 24,03 24,03 22,24 23,99 24,03 24,03 21,73 17,55 24,03
W 4762 4743 4697 4697 4618 4618 3112 3112 3037 3037
D 2078 1935 1917 2070 1887 1884 1269 1404 1696 1235
Engine output parameters
PW 712 719 462 476 356 354 343 356 476 711
EPW 739 752 485 504 375 372 358 375 504 737
PSFCuninstalled 0,324 0,323 0,282 0,278 0,265 0,266 0,269 0,265 0,278 0,325
ESFCuninstalled 0,312 0,309 0,269 0,263 0,252 0,253 0,258 0,252 0,263 0,313
Tinst 11310 9503 4425 3730 2787 2862 3399 2787 3730 11757
TSFCuninstalled 5,673 6,782 8,184 9,864 9,408 9,133 7,562 9,408 9,864 5,453
wi 3,540 3,561 1,803 1,840 1,246 1,241 1,215 1,246 1,840 3,536
P0 103,0 103,7 48,9 50,1 32,6 32,4 31,7 32,6 50,1 102,8
P2 103,0 103,7 48,9 50,1 32,6 32,4 31,7 32,6 50,1 102,8
P3 1067,7 1075,2 507,6 519,7 337,9 336,3 328,3 337,9 519,7 1066,3
P31 1067,7 1075,2 507,6 519,7 337,9 336,3 328,3 337,9 519,7 1066,3

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P4 1035,6 1043,0 492,4 504,1 327,8 326,2 318,4 327,8 504,1 1034,3
P41 1035,6 1043,0 492,4 504,1 327,8 326,2 318,4 327,8 504,1 1034,3
P416 261,4 262,5 151,3 153,5 110,6 110,2 108,5 110,6 153,5 261,3
P44 261,4 262,5 151,3 153,5 110,6 110,2 108,5 110,6 153,5 261,3
P46 256,2 257,2 148,2 150,4 108,4 108,0 106,3 108,4 150,4 256,0
P48 104,9 104,9 48,2 48,2 31,1 31,1 31,1 31,1 48,2 104,9
P5 104,4 104,4 48,0 48,0 31,0 31,0 31,0 31,0 48,0 104,4
T0 289,5 290,1 252,1 253,8 234,0 233,7 232,1 234,0 253,8 289,4
T2 289,5 290,1 252,1 253,8 234,0 233,7 232,1 234,0 253,8 289,4
T3 660,0 661,3 579,3 583,0 539,4 538,7 535,1 539,4 583,0 659,8
T31 660,0 661,3 579,3 583,0 539,4 538,7 535,1 539,4 583,0 659,8
T4 1368,7 1368,7 1368,7 1368,7 1368,7 1368,7 1368,7 1368,7 1368,7 1368,7
T41 1334,9 1335,0 1331,5 1331,7 1329,9 1329,9 1329,7 1329,9 1331,7 1334,9
T416 1007,0 1006,3 1047,0 1045,1 1066,4 1066,7 1068,5 1066,4 1045,1 1007,1
T44 990,9 990,4 1025,6 1024,1 1042,5 1042,8 1044,4 1042,5 1024,1 991,0
T46 990,9 990,4 1025,6 1024,1 1042,5 1042,8 1044,4 1042,5 1024,1 991,0
T48 818,5 817,3 807,4 803,5 800,0 800,8 804,8 800,0 803,5 818,7
T5 818,5 817,3 807,4 803,5 800,0 800,8 804,8 800,0 803,5 818,7
T7 818,5 817,3 807,4 803,5 800,0 800,8 804,8 800,0 803,5 818,7

Equations P3  P2 C (17)

For the sizing of UAV, general weight formulae: Compressor exit airflow:
W0 = We + WPL + Wf + Wmisc (1) w 3  w 2 - w 2 ( 1  1 ) (18)
We = Waf + Wlg + Weng + Wsys (2) T3  T2
Waf = Wfus + Wnac + Wh.tail + Wv.tail (3) Tbleed  (19)
2
Flight durations and range calculations: G C  w 3 C PC (T3 - T2 )  (w 2 - w 3 )C PC ( Tbleed - T2 ) (20)

RF = RCL + RCR + RD (4) Compressor exit diffuser:

tF = tCL + tCR + tD (5)
tB = tF+ tGR (6) Total temperature does not change, total pressure
reduces by a ratio. Moreover, air mass flow rate is
RCR = V0 L ln Wi1 (7)
C D Wi
reduced due to bleed and cooling air extraction.
tCR = L / D ln Wi1 (8) T31  T3 (21)
C Wi
P31  P3 (22)
A two spool turboprop engine design point calculation: w 31  w 3 - w 2 (  2a   2b   3 ) (23)
   1  2  (9)
T0  Tamb 1   M  Combustor and NGV station:
  2  

P4  P31 1  Pb  (24)
    1  2   1 (10)
P0  Pamb 1   M  f, fuel air ratio calculation:
  2  

Intake: f1 = 0,10118 + (2,00376e-05)(700 - T31) (25)

f2 = 3,7078e-03 – (5,2368e-06)(700 - T31)-
T2  T0 (11) (5,2632e-06)T4 (26)
P2  P0 * 1  Pin  (12) f3 = (8,889e-08) abs(T4-950) (27)
 f = (f1 -(f12 + f2)½ - f3)/ ηb (28)
    1  2   1 (13) wf = f (w31 + w2 ε3) (29)
 1   M 
  2   w4 = w31+ wf (30)
Ccorr  1
     1  2  2 w41 = w31 + w2 ε3+wf (31)
 1    M 
(w 4 C PT T4  w 2  3 C PC T31 )
   2  
T41  (32)
P0 / PSL (14) w 41 C PT
Ccorr 
T0 / TSL 
1
2 P41  P4 (33)

Compressor section: No total pressure loss is assumed at NGV.

 C 1

2 
 C    1 C (15) High pressure turbine:
 1
 C  *  1
C

C C , pol GHPT = GC/ ηm (34)

1  1
(16) T416= T41 – GHPT/(w41 CPT) (35)
T3  T2 (1  ( C   1))
2

60
HPT isentropic efficiency can be calculated by assuming ΠT Nozzle pressure ratio in choked condition:
(Turbine total pressure ratio) an initial value (such as 4):  C 
- 
 1   C    C 1 
(1 -  T )  T , pol P5 /P7s,c   (58)
1 - T T  2 
(36)
 416  1-  T If, design nozzle pressure ratio ΠN > P5/P7s,c than the
1 - T T nozzle is choked and M7 is equal to 1. Choked nozzle
 T -T 
exit static temperature:
log 1 - 41 416 
  416 T41   
e (37)   (59)
T  
T5
1-  T T7s,c
 T 1
 1 
T  2 

a7   T RT7s,c 
1
If equations (4.46) and (4.47) are iterated 3-5 times, ΠT 2 (60)
and η416 values converge. (61)
0.001w5
high pressure türbin total pressure: A7, c =   T 
P5   T    T    2( T )
1/ 2

N cd    
P416 = P4/ ΠT (38) (T5 )1/2  R   2 
w416 = w41 (39)
FA= w5 a7 + A7,c Ncd (P7s,c - Pamb) 1000 (62)
Rotor cooling air addition is done numerically at Station
44 (Kurzke, 2007): If nozzle is not choked, M7 is not 1 and can be
calculated as follows:
w44 = w416 + w2  2a (40) 1
 1 T 
 2
M7 =  2  Pamb  T (63)
(w 416C PT T416  w 2  2a C PC T31 )    1
T44  (41)   T  1  P5  
   
w 44 C PT
P44 = P416 (42)  
 T5  (64)
T7s  
Turbine duct (between HPT and LPT):  T 1 2 
 1 M7 
 2 
Total temperature does not change, but total pressure is
a7   T RT7s 2
1
reduced. (65)
FA= w7 a7 M7 (66)
T46 = T44 (43)
P46 = P416 (1 - ∆Pt.duct) (44) Propeller Thrust Calculation:
w46 = w44 (45)
V0 (67)
Low pressure turbine: J d
n
P5 = ΠN Pamb PW (68)
(46) C PW 
n3d 5 
Where ΠN, exhaust nozzle pressure ratio is a design input.  prop,d C PW
CF  (69)
P48 = P5 /(1 - ∆Pj.pipe) (47) J
ΠT = P46/ P48 (48)
(1- T )  T , pol for static conditions (V0=0) :
T
1 - T (49)
46  1- T C F  ( prop,s C PW ) 2 / 3 ( / 2)1 / 3 (70)
T
1 - T C F PW (71)
1- T
FP 
C PW nd
T48  T46 (1  46 (1 - T  T )) (50)
GLPT = (T46 - T48) (w46 CPT) (51) Total Thrust and SFC:
PW = GLPT ηm (52)
V0 = M0 (γC R Tamb)1/2 (72)
Low pressure turbine exit: Fnet = FP +FA Ncx - w2 V0 (73)
TSFC = wf /Fnet (74)
w48 = w46 (53) PSFC = wf /PW (75)
w5 = w48 + w2 ε1+ w2 ε2b (54) EPW = PW+V0 FA/ηp (76)
w 48 C PT T 48  w 2   C PC T bleed w 2  2b C PC T31 (55) ESFC = wf /EPW (77)
T5 
w 5 C PT
Exhaust Nozzle: REFERENCES

T7= T5 (56) Austin R., 2010, Unmanned Aircraft Systems, Wiley,

w7=w5 (57) West Sussex, UK.

61
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Approach (Third Ed.), AIAA Series, Washington, DC.
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DISCLAIMER
Honeywell, 2014, Product Brochures,
http://www51.honeywell.com/aero/common/documents/ The views and opinions expressed in this article are
myaerospacecatalog-documents/BA_brochures- those of the author and do not represent the official
documents/TPE331-10_PredatorB_0292-000.pdf policy or position of Tusaş Engine Industries Inc. or any
other company and institution.

Ali DİNÇ

Ali DİNÇ was born in 1970. He graduated from the department of Aeronautical Engineering,
Middle East Technical University (METU) in 1992. He earned his MSc degree in 1995 within
the scope of a cooperation project between METU and Tusaş Engine Ind. Inc. (TEI). Then, he
worked for TEI in different positions of design engineering and management levels in aircraft
engine components design area. In this context, he has worked for TEI in the design offices of
General Electric-USA and ITP-Spain for a total of 5 years for the development of T38/J85
engine exhaust module design and A400M/TP400 turboprop engine development projects,
respectively. He received his PhD degree at Anadolu University, School of Civil Aviation in
2010. He currently works as a senior engineer in Chief Engineering Office of TEI.

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