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Design of UAV Systems. c 2002 LM Corporation . Propulsion. Lesson objective - to discuss Propulsion and propulsion parametrics including … Rationale Applications Models. Expectations - You will understand when and how to use parametric propulsion relationships. 18-1.

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  1. Design of UAV Systems c 2002 LM Corporation Propulsion • Lesson objective - to discuss • Propulsion and propulsion parametrics • including … • Rationale • Applications • Models Expectations - You will understand when and how to use parametric propulsion relationships 18-1

  2. Design of UAV Systems c 2002 LM Corporation Lesson 5a - Air vehicle parametrics Definitions • From Webster’s New Collegiate Dictionary • Parameter – any set of physical properties whose value determine the characteristics or behavior of a set of equations • Our definition • Propulsion parametric – fundamental design parameter whose value determines the design or performance characteristics of an engine • Usually (but not always) a multi-variable relationship • e.g., SFC (WdotF/Bhp), Bypass ratio (BPR), etc. • Parametric model – Parametric based design approach to define, size, estimate performance and do trade offs on propulsion systems • Different from the traditional approach 18-2

  3. Design of UAV Systems • Simple models that correlate thrust or power, weight, fuel flow, speed and altitude • - Most based on historical and technology trend data • - Others based on simple definitions • - Some based on non-dimensional analysis (See RosAP6.2.5) • Internal combustion (IC) engines are easiest to model • - Simple (and generally independent) variables • - One complexity is fixed pitch propeller performance • Jet engines are more difficult to model parametrically • - Many interrelated design and operating variables • Turboprop (TBProp) parametric models are in between Propulsion parametrics can be used for pre-concept design but engine company models should be used as soon as they are available c 2002 LM Corporation Lesson 5a - Air vehicle parametrics Parametrics models 18-2a

  4. Design of UAV Systems c 2002 LM Corporation Propulsion Key parametrics • Engine Power-to-weight ratio (HP0/Weng) • - HP0 = Maximum power (hp, uninstalled, sea level static) • - Weng = Engine weight (lbm, uninstalled) • Engine thrust-to-weight ratio (T0/Weng) • - T0 = Maximum thrust (lbf, uninstalled, sea level static) • - Weng = Engine weight (lbm, uninstalled) • Specific Fuel Consumption (SFC) • SFC = Fuel flow/Power (WdotF/HP) • SFC0 = WdotF0/ HP0 (lbm/hp-hr) • Thrust Specific Fuel Consumption (TSFC) • TSFC = Fuel flow/Thrust available(WdotF/Ta) • TSFC0 = WdotF0/T0 (lbm/hp-hr) • Specific Thrust (Fsp) • - Fsp = Thrust/Airflow (T/WdotA) • - Fsp0 =T0/WdotA0 (lbf-sec/lbm) or or Note - “0” postscript indicates sea level static (V=0) conditions 18-3

  5. Design of UAV Systems c 2002 LM Corporation Lesson 5a - Air vehicle parametrics Engine size • Any number of performance requirements can drive engine size (See RayAD 5.2) • - Takeoff • - Ground roll and distance over an obstacle • - And/or balanced field length (BFL) • - Time and/or distance to climb • - Cruise altitude and/or speed • - Acceleration and/or turn performance • - Engine out performance (for multi-engine aircraft) • Historical thrust-to-weight or power-to-weight data can be used for first pass sizing • - See RayAD Tables 5.1 and 5.2 • UAV historical data is limited and we will use takeoff requirements for initial sizing • - See RayAD Figure 5.4 18-4

  6. Design of UAV Systems c 2002 LM Corporation Propulsion Sizing - manned vs. unmanned • Raymer (power-to-weight) • GA Single 0.07 • GA Twin 0.17 • Twin turboprop 0.20 • Raymer (thrust-to-weight) • Trainer 0.4 • Bomber 0.25 • Transport 0.25 UAV data from various sources including Janes, Unmanned Air Vehicles 18-5

  7. Design of UAV Systems c 2002 LM Corporation Lesson 5a - Air vehicle parametrics IC engines • Engine power available (BHP) • Output per unit size varies by type • - Small piston engines run at high RPM, have higher output per unit size. Same for rotary engines • For given engine at given altitude and RPM • - Almost no variation with speed • At given RPM, manifold pressure varies with altitude • Max power varies with air density ratio (see RayAD Eqn13.10) • - Bhp = Bhp0*(8.55*-1)/7.55 (18.1) • Engine SFC • Runs high for small and rotary engines • For given engine varies slightly with power available • - Typically 5-10% lower at cruise condition • Engine power-to-weight varies with engine type 18-6

  8. Design of UAV Systems c 2002 LM Corporation Propulsion IC engine parametrics Compilation of data from various sources including: Roskam, Aerodynamics & Performance (RosAP); Janes, Aero Engines; Janes, Unmanned Air Vehicles; www.tcmlink.com/producthighlights/www.lycoming.textron.com/main 18-7

  9. Design of UAV Systems c 2002 LM Corporation Propulsion IC engine size = 20 pcf = 30 = 40 (a) (b) • Charts 18-7/8 show only contemporary IC engines • Earlier engines were much larger Compilation of data from various sources including: - Roskam, Aerodynamics & Performance (RosAP) - Janes, Aero Engines - Janes, Unmanned Air Vehicles -www.tcmlink.com/producthighlights/ -www.lycoming.textron.com/main (c) 18-8

  10. Design of UAV Systems c 2002 LM Corporation Lesson 5a - Air vehicle parametrics Propellers • Two basic types of propellers (See RayAD10.4 & 13.6) • - Fixed pitch and Variable pitch • Efficiency (p) varies with design and installation • - Blockage and flow “scrubbing” generate losses • Fixed pitch efficiency varies with advance ratio (J) • - See RayADFig13.13 • - Props generally designed for climb or cruise • Variable pitch efficiency typically constant over range of design speeds • - Use a nominal p = 0.8 for an initial guess • Thrust horsepower (THP) defined by • - Thp = Bhp*p =Ta*V(fps)/550=Ta*V(KTAS)/325.6 • Where ……….Ta = Thrust available • Therefore…….. • - Thrust available decreases with speed • Ta = 325.6*BHP*p /KTAS (18.2) 18-9

  11. Design of UAV Systems c 2002 LM Corporation Lesson 5a - Air vehicle parametrics Propeller size • Key sizing constraint is tip speed (Mach number) • - Linear function of engine RPM (V=R*) • Lightest-most reliable design is direct drive • - No gear reduction • Some UAVs use high RPM engines with belt systems for speed reduction - Manned aircraft prop sizing can be used for UAVs - There will always be exceptions for special application aircraft (e.g.Altus) Compilation of data from unpublished sources plus: - Janes, Unmanned Air Vehicles 18-10

  12. Design of UAV Systems c 2002 LM Corporation Propulsion Propeller parametrics Nb = Number of blades Rp = Prop radius (in) Weight/Total blade length(in) ≈ 0.8 lb/in Raw data from Janes All the Worlds Aircraft 18-11

  13. Design of UAV Systems c 2002 LM Corporation Lesson 5a - Air vehicle parametrics Turboprop engines • Have typical prop characteristics (and constraints) plus jet exhaust for thrust augmentation • Thrust component is added for equivalent shaft HP, the difference between Shp and EShp (see RayAD 13.7) • Power available decreases with altitude • Decreases with pressure  (see RayAD Table E.3) • SFC variation with altitude less severe than jet • Typical Lth/Dia = 2 - 3 • Typical density = 22 pcf • Typical diameter = [4*Vol/(*L/D)]^1/3 • - From Vol(cyl) = [/4][D^2]L[D/D] 18-12

  14. Design of UAV Systems c 2002 LM Corporation Propulsion Typical TBProp parametrics Turboprop engines Data compiled from RosAP and Janes Aero Engines 18-13

  15. Design of UAV Systems c 2002 LM Corporation Propulsion Jet engines • Includes turbojet (TBJet) and turbofan (TBFan) engines • Thrust, weight, airflow and TSFC depend on design and technology content and operating conditions • Difficult to capture in general purpose parametrics • Thrust available decreases with altitude - all engines • TSFC decreases with altitude - all engines • Effect of speed on thrust varies with bypass ratio (BPR) • Low BPR - thrust generally increases with speed • High BPR - thrust decreases with speed • See RayAD Appendix E for specific examples • Thrust-to-weight varies with engine size, BPR and advanced technology content • Physical geometry primarily a function of design BPR • Physical installation reduces thrust by 5-20% 18-14

  16. Design of UAV Systems V Gas Generator (GG) Fan GG Airflow GG thrust Fan Airflow Fan thrust • Bypass ratio (BPR) = Fan airflow/total airflow • WdotAgg  WdotA*1/(BPR+1) • WdotAfn  WdotA*BPR/(BPR+1) Turbojet (TBJ) BPR = 0, Turbofan (TBFan) BPR < 10; TBProp BPR >> 10 c 2002 LM Corporation Propulsion Generic jet engine 18-15

  17. Design of UAV Systems SFC0 - Small turbojets 1.5 1 0.5 0 1000 2000 3000 T0 (lbf) c 2002 LM Corporation Propulsion Typical TBJet parametrics Data from Roskam, Aerodynamics & Performance (RosA)P and Janes, Aero Engines All engines (more later) 18-16

  18. Design of UAV Systems TSFC 1 0.75 0.5 0.25 0 0 2 4 6 8 10 BPR c 2002 LM Corporation Propulsion Typical TBFan parametrics (lbm/hr-lbf) Data from Roskam, Aerodynamics & Performance (RosA)P and Janes, Aero Engines 18-17

  19. Design of UAV Systems c 2002 LM Corporation Propulsion TBJet and TBFan size • Despite its apparent simplicity, Raymer’s engine size parametric correlates well with our database • - One difference is that Raymer bases his correlation on engine inlet diameter • - Our data shows that it correlates with overall engine diameter • Parametrics are for uninstalled engine weight • - Nominal TBFan T0/Weng = 5.5 • - An installation factor of 1.3 is applied to estimate installed engine weight 18-18

  20. Design of UAV Systems c 2002 LM Corporation Propulsion TBFan parametrics – cont’d Data from Roskam, Aerodynamics & Performance (RosA)P and Janes, Aero Engines 18-19

  21. Design of UAV Systems c 2002 LM Corporation Propulsion Afterburning • Way to augment jet engine performance to meet peak thrust requirements such as….. • Takeoff….combat maneuvers….supersonic flight • Works by injecting fuel into engine exhaust to react residual oxygen and increase temperature/jet velocity • Inefficient turbojet engines and turbofans can achieve high augmentation ratios - lots of air to “burn” • Efficient turbojets achieve low augmentation ratios • - Most of the air already “burned” • Essentially a ramjet on the rear of the engine • Only works for low-to-moderate BPR turbofans • - High BPR fans have insufficient overall pressure ratio • Relatively light weight but very fuel inefficient • High noise levels limit civil applications 18-20

  22. Design of UAV Systems c 2002 LM Corporation Propulsion A/B parametric data Data from Roskam, Aerodynamics & Performance (RosAP) and Janes, Aero Engines 18-21

  23. Design of UAV Systems c 2002 LM Corporation Propulsion TBProp baseline • We will use parametric data to make a first pass engine size estimate for our example UAV (see chart 15-40) • We assume a nominal TBProp UAV wing loading (W0/Sref) = 30 psf, a typical plan flap Clmax =1.8 (See RayAD Fig 5.3) and a standard Vto/Vs = 1.1 • From RayAD Figure 5.4, required takeoff parameter (TOP) for a 1500 ft takeoff • ground roll = 220 or: • 220 = • [W0/Sref]/[Clto*T0/W0] • For Clto = Clmax/(Vto/Vs^2) = 1.49 and W0 = 1918 lbm • BHp0/W0 = 0.092 • BHp0 = 176.5 BHp • BHp0/W0correlates well with our parametric data 18-22

  24. Design of UAV Systems c 2002 LM Corporation Propulsion Application – TBProp • Chart 18-13 shows turboprops of this small size class should be available at 2.25 Shp/lb • - Nominal weight would be 78.4 lbm • At a density of 22 lb/cuft, volume would be 3.6 cuft • At nominal Lth/Diam = 2.5, engine diameter (Deng) = [4*Vol/(*Lth/Deng)]^1/3 ≈ 1.22ft and length (Leng) = 3ft • SFC0 would about 0.65 lbm/hr-Bhp • In reality, however, there are no TBProps this small 18-23

  25. Design of UAV Systems c 2002 LM Corporation Propulsion TBFan alternative • We will also use parametric data to make a first pass engine sizing for the TBFan alternative • We assume a nominal TBFan UAV wing loading (W0/Sref) = 40 psf, a typical plan flap Clmax =1.8 (See Raymer AD Fig 5.3) and a standard Vto/Vs = 1.1 • From RayAD Figure 5.4, required takeoff parameter (TOP) for a 1500 ft takeoff • ground roll = 100 or: • 100 = • [W0/Sref]/[Clto*T0/W0] • For Clto = Clmax/(Vto/Vs^2) = 1.49 and W0 = 2939 lbm • T0/W0 = 0.269 • T0 = 790 Lbf • T0/W0correlates well with our parametric data 18-24

  26. Design of UAV Systems c 2002 LM Corporation Propulsion TBFan • From charts 18-16/17 we estimate T0/Weng ≈ 5.5 • - Our TBFan would weigh 144 lbm • - At BPR = 5, WdotAmax = 790/30 = 26.3 pps • - From charts 18-16/17 Deng ≈ 12 in, Leng ≈ 24 in • From charts 18-17/19 TSFC0 ≈ 0.4 and TSFCcr ≈ 0.65 • But unfortunately, there are no BPR = 5 turbofans of size class (see ASE261.Engine database.xls) • However, there might be some under development 18-25

  27. Design of UAV Systems c 2002 LM Corporation Propulsion Overall results • The TBProp parametrics show the SFC0 value assumed in the Lesson 15 example is low (0.4 vs 0.65) • Small engines are less efficient than larger ones • The performance impact will significant but it will make make the results fit better with our sizing parametrics • And there are no engines available at the size required • The TBFan parametrics show that Raymer’s values of cruise TSFC are optimistic (which we already knew) and that there are also no small BPR = 5 TBFan engines • - Size effects are likely to reduce TSFC also • However, we will continue our study as if engines were available since we really don’t know yet what size air vehicle we will end up with • - We will, however, note these issues as development risk items and consider the implications when we select our final configurations 18-26

  28. Design of UAV Systems c 2002 LM Corporation Propulsion Next subject • From Webster’s New Collegiate Dictionary • Parameter – any set of physical properties whose value determine the characteristics or behavior of a set of equations • Our definition • Propulsion parametric – fundamental design parameter whose value determines the design or performance characteristics of an engine • Usually (but not always) a multi-variable relationship • e.g., wing loading (W0/Sref), Swet/Sref, etc. • Parametric model – Parametric based design approach to define, size, estimate performance and do trade offs on propulsion systems • Different from the traditional approach 18-27

  29. Design of UAV Systems c 2002 LM Corporation Propulsion Parametric models • In the absence of real data, engine parametric models can be used to provide reasonable trends for use in pre-concept and conceptual design • For example, Equation 5.4 (RayAD Eq. 13.10) captures IC engine altitude effects and is useful for initial design • No similar effects are captured in traditional jet engine parametric models such as RayAD Eq. 10.5-10.15 • Mach and altitude effects are absent • More general purpose thrust and fuel flow models are needed and none exist • Therefore, we will have to develop our own jet engine models (both TBJet and TBFan) • We will use the engine performance charts in RayAD, Appendix E as the basis for these models 18-28

  30. Design of UAV Systems c 2002 LM Corporation Propulsion IC parametric model • Use Equation 18.1 (RayAD Eq. 13.10) to calculate maximum power as a function of altitude • BHP = BHP0*(8.55* -1)/7.55 (18.1) • where • BHP0 = maximum power, SLS (sea level static) •  = air density ratio • Calculate cruise performance at 75% takeoff power • Assume nominal 80% propulsion efficiency (p) • Estimate thrust available from • Ta = 325.6*BHP*p/KTAS (18.2) • Estimate fuel flow from • WdotF = SFC*BHP • where • SFC assumed constant (use SFC0 from chart 18-7) • Estimate engine weight from chart 18-7 • For supercharged engine, assume pressure ratio/stage = 2 • See Raymer Figure 13.10 (page 394) • Adjust engine weight as appropriate 18-29

  31. Design of UAV Systems V Gas Generator (GG) Fan GG Airflow GG thrust Fan Airflow Fan thrust • Assumptions • GG Fsp = constant = Fsp-gg • Fan Fsp = (V0/V)*Fsp-fn(SLS) • BPR = Constant Bypass ratio (BPR) WdotAgg  WdotA*1/(BPR+1) WdotAfn  WdotA*BPR/(BPR+1) Turbojet BPR = 0, Turbofan BPR < 10; Propfan BPR >> 10 c 2002 LM Corporation Propulsion Jet parametric model 18-30

  32. Design of UAV Systems • Assume jet engine thrust (subsonic) can be modeled as the sum two (2) simplified components • - “Gas generator”(gg) - the core engine of a turbofan or turboprop • - “Fan” (fn) - the fan or propeller • Assume core engine thrust varies with core airflow (constant core engine Fsp - a very simple approximation) • Assume fan Fsp varies inversely with speed ratio (V0/V) • - Like a propeller • - Select non-zero V0 to fit data as appropriate • Assume fan bypass ratio remains constant • Assume all other engine parameters follow “corrected” performance relationships (See P&W handbook, page 129) • - Ta = Ta0*(18.3) • - WdotF = WdotF0**sqrt(^k) (18.4) • - SFC = SFC0*sqrt( ^k) (18.5) • - WdotA = WdotA0* /sqrt() (18.6) • - f/a = (f/a0)*sqrt()*sqrt( ^k) ≈ (f/a0)* ^.75 (18.7) k = f(cycle) but ≈ 0.5 c 2002 LM Corporation Propulsion Jet parametric model 18-31

  33. Design of UAV Systems c 2002 LM Corporation Propulsion Jet parametric model (cont’d) • Using these simplifying assumptions, thrust available can be estimated from: • Ta = WdotA*Fsp = WdotA*[Fsp-gg/(1+BPR) + • Fsp-fn*(V0/V)*(BPR/(1+BPR)] (18.8) • where • WdotA = Total airflow (note :WdotA-gg = gg airflow) • Fsp-gg = Core engine Fsp • Fsp-fn = Fan Fsp (varies with number of fan stages) • V0 = Non-zero reference speed (select to fit data) • BPR - Fan bypass ratio (given by design) • Estimate installation losses at 5-20% (more to follow) • Estimate airflow from • WdotA = WdotA0*delta/sqrt() (18.9) • where •  = ( @h)*(1+.2M^2)^3.5 (18.10) •  = (@h)*(1+.2M^2) (18.11) • Estimate fuel flow from WdotA-gg*corrected fuel/air ratio • = WdotA/(1+BPR)*(f/a0)* ^.75 18-32

  34. Design of UAV Systems RayAD Appendix E engines (Table E.1 - Low BPR, Table E.2 - High BPR and Table E.3 - TBP) were modeled parametrically using the following values: Core Fsp (sec) Fan Fsp (sec) V0 (KTAS) BPR Fuel/air ratio LBPR HBPR TBP 90 90 90 66 25 5 100 100 50 0.4 8 133 0.0292 0.0292 0.0292 Even though our parametric model is unique to this course, engine companies often provide generic “cycle decks” which produce similar (but more accurate) results c 2002 LM Corporation Propulsion RayAD model matching 18-33

  35. Design of UAV Systems - Value (90 sec) selected from chart 18-17 for BPR = 0 - LBPR value selected to match data, HBPR Fsp scaled based on fan pressure ratio differences (1.6 vs. 4.3), TBP Fsp estimated at 20% HBPR Fsp. - Values selected to match data - Given values used except for TBP which was selected to match data - Value selected to match data Core engine Fsp Fan Fsp V0 BPR Fuel/air ratio c 2002 LM Corporation Propulsion Model input rationale 18-34

  36. Design of UAV Systems c 2002 LM Corporation Propulsion Model correlation - LBPR (lbm/hr-lbf) (lbm/hr-lbf) 18-35

  37. Design of UAV Systems c 2002 LM Corporation Propulsion Model correlation - HBPR (lbm/hr-lbf) (lbm/hr-lbf) 18-36

  38. Design of UAV Systems c 2002 LM Corporation Propulsion Model correlation - TBP (lbm/hr-lbf) (lbm/hr-lbf) 18-37

  39. Design of UAV Systems Fsp-fn parametric Fsp-fn parametric (Assumed Fsp-gg = 90) (Assumed Fsp-gg = 80) 80 100 Calculated values Calculated values 80 Model values Model values 60 60 Est. upper bound 40 40 20 20 0 0 0 2 4 6 8 10 0 2 4 6 8 10 BPR BPR c 2002 LM Corporation Propulsion Database comparison - TBF • Even though model Fan Fsp values generally match Raymer’s models at BPR = 0.8 and 8.0 by definition • - We have no idea what Fan Fsp might look like at intermediate BPR values • - And we have no idea how they correlate with real ones • We can get answers by assuming values of Fsp-gg and use Eq 18.18 to calculate Fan Fsp for typical engines • The Fsp-gg = 80 data looks like it provides a better fit 18-38

  40. Design of UAV Systems c 2002 LM Corporation Propulsion Other data comparisons • Actual engine performance data can also be used to check and/or calibrate parametric model estimates • For example, parametric model performance estimates for typical TBProp and TBFan engines can be compared to actual engines under the same flight conditions • But because of design differences, even real engines will show performance variations • Nonetheless, the comparisons can be used to generate multipliers to ensure the model estimates match actual engine performance ranges • Comparisons with database TBProp and TBFan engine performance are shown in the following chart • TBProp model thrust available and SFC are seen to fit within the data spread, albeit somewhat optimistically • TBFan thrust fits the data but TSFC is about 15% high • A 0.87 TBFan TSFC multiplier will compensate for it 18-39

  41. Design of UAV Systems TBProp SFC at 250 kts) TBProp power ratio at 250 kts 0.60 1.1 TPE331-14 TPE331-14 PT6A-41 (Flat rated) PT6A-41 (Flat rated) 0.55 0.9 Other Other 0.50 0.7 0.45 0.5 0.40 0.3 10 20 30 40 50 10 20 30 40 50 Altitude (Kft) Altitude (Kft) TBFan Cruise TSFC (35-40Kft) TBFan Cruise thrust ratio (35-40Kft) 1 M = 0.7-0.85 0.3 0.9 M = 0.7-0.85 0.8 0.25 0.63/36Kft 0.7 0.2 0.6 0.63/36Kft 0.15 0.5 0 2 4 6 8 10 0 2 4 6 8 10 BPR BPR c 2002 LM Corporation Propulsion Performance correlations Data from Roskam, Aerodynamics & Performance (RosA)P and Janes, Aero Engines - Propulsion model estimate 18-40

  42. Design of UAV Systems c 2002 LM Corporation Propulsion A note about Turboprops • Raymer’s Appendix E.3 TPB model is somewhat unique in that performance is expressed in terms of thrust and TSFC, not Shp or Eshp and SFC = WdotF/Hp • - This makes us work the problem backwards • - In a traditional propeller aircraft analysis, we first calculate Bhp available and then multiply by p to determine thrust horsepower (Thp) available • - The Breguet range equation includes p in the numerator and SFC is based on uninstalled Bhp • - In our model we calculate thrust and fuel flow directly • - We, therefore, have to calculate Thp from the definition Thp = T*V(fps)/550 = T*V(kts)/325.6 and then divide by p to get Shp • - Then we calculate SFC from fuel flow and Shp 18-41

  43. Design of UAV Systems • An important issue for any engine model or data is installation losses • - All installations degrade power or thrust available compared to engine company test stand-type data • - The differences are large enough to effect even pre-concept design estimates • Jet engine losses derive from multiple factors • - Inlet and nozzle losses • - Power extraction • Turboprops also have to deal with prop efficiency • IC engine installations have similar losses (air induction system, mufflers, generators and props) • - Bleed air • - Etc. c 2002 LM Corporation Propulsion Installation losses • We will capture these effects using simple installed performance knock down factors • - We will use 0.8 - 0.95 for TBJ and TBF installations • - p will capture all losses for ICs and TBPs • During conceptual design, actual performance losses should be calculated for the specific designs studied 18-42

  44. Design of UAV Systems c 2002 LM Corporation Propulsion Example - TBProp • 1. Our TBProp UAV weighs 1918 lbm, has a balanced field length requirement of 3000 ft (ground roll = 1500 ft) and Clto = 1.49 and wing loading of W0/Sref = 30 psf (qto = 20.2 psf , takeoff speed = 77.2 kts). We assumed a nominal cruise of 180 kts at 27.4Kft, an initial cruise weight (w4) = 1726 and a cruise lift-to-drag ratio (LoDcr) of 23. What size engine is required for takeoff and will it meet cruise requirements? • 2. From RayAD Figure 5.4, the required prop aircraft ground roll takeoff parameter for is 220where • TOP =[W0/Sref]/[CLt/o*(Bhp0/W0)] = • = qt/o/(Bhp0/W0) • Bhp0/W0 = qto/220 = 0.092 • - Engine size, therefore, is • BHP0 = 1918Klb*0.092 = 176.5 Bhp or 18-43

  45. Design of UAV Systems c 2002 LM Corporation Propulsion TBProp takeoff estimate • 3. Because propeller models have singularities at V=0, the TBProp is sized at V = V0 = 50 kts (M = 0.076) • - At an assumed p = 0.8*, takeoff thrust (T0) can be calculated directly by definition of BHP or • T0  Bhp0*550*p/KTAS*1.689 = 919.6 lbf • - Equation 18.8 is solved for total airflow using chart 18.33 TBP model values or Wdota = 919.6/[90/134 +5*(50/50)*(133/134)] = 163.2 pps • - Knowing WdotA0, the TBP model can now predict thrust, airflow and fuel flow at cruise conditions • - For simplicity, this will be done only once at Vcr = 180 kts and an altitude of 27.4 Kft (M=0.3) • - Then it will be programmed in a spreadsheet * p is assumed to account for all installation losses 18-44

  46. Design of UAV Systems c 2002 LM Corporation Propulsion TBProp cruise estimate • 4.Equations 18.9-11 provide estimates of total airflow (WdotA) at cruise (M=0.302), where •  = (a@27.4Kft)*(1+.2M^2)^3.5 = 0.3556 •  = (a@27.4Kft)*(1+.2M^2) = 0.8264 • WdotA = WdotA0*/sqrt() = 64 pps • - Core airflow by definition of BPR is Wdota/(BPR+1) or for BPR = 133, Wdota-gg = 0.48 pps • - Fuel flow is calculated using the model fuel-to-air ratio value f/a = 0.0292 corrected for , or …. • WdotF = WdotA-gg*(f/a)* ^.75 = 0.012 pps or 43.4 pph • - Equation 5.8 is used to calculate Ta where • Ta = WdotA*(Fsp-gg/(1+BPR)+Fsp-fn*(V0/V) *(BPR/(1+BPR)) • = 64*(90/134+5*(50/180)*(133/134) = 130.9 lbf 18-45

  47. Design of UAV Systems c 2002 LM Corporation Propulsion TBProp cruise - cont’d • - Thp is calculated using Eq 18.2 or • Thp =130.9*180/325.6 = 72.4 Hp • while • Shp = 72.4/0.8 = 90.4 Bhp • - Finally SFCcr  WdotF/Shp is calculated and found to be • SFCcr = 43.4pph/90.4Bhp = 0.48 pph/Bhp • - Next we need to compare thrust “available” against thrust “required”  drag • - We get this by dividing weight by LoDcr or • D ≈ 1726lbm/23 = 75 lbf which is 57% of Ta and shows that the TBProp meets cruise thrust requirements at 180 kts 18-46

  48. Design of UAV Systems c 2002 LM Corporation Propulsion TBProp summary TBP sizing 1. Select takeoff speed (Vto), calculate qto  [W/S]/Clto 2. Estimate takeoff Bhp0 required (RadAD Fig 5.4) 3. Select takeoff “sizing” speed = V0 (i.e. 50 kts) 4. Calculate Tavail at sizing speed (Ta0 = 325.6pBhp0/V0) 5. Calculate WdotA0 from Fsp and Ta0 at V0 (Eq 18-8, BPR = 133) TBP performance 1. Select speed (KTAS) & and altitude (h) 2. Calculate ,  and M at h (atmosphere spreadsheet) 3. “Correct”  and  for M (Eqs 18-10 and 18-11) 4. Calculate total (prop+engine) WdotA (Eq 18-9) 5. Calculate engine airflow (WdotA-gg = WdotA/[BPR+1]) 6. Calculate corrected fuel-to-air ratio (Eq 18.7) 7. Calculate fuel flow (WdotF= WdotA-ggcorrected fuel-to-air ratio) 8. Calculate thrust available (Eq 18-8) 9. Calculate uninstalled Bhp (= TaKTAS/[325.6 p) 10. Calculate uninstalled SFC [SFC = WdotF/Bhp(uninst)] 11. Check that Ta > D = W/LoD 18-47

  49. Design of UAV Systems c 2002 LM Corporation Propulsion Typical example - TBFan • Our TBFan alternative weighs 2914 lbm, has a balanced field length requirement of 3000 ft (ground roll = 1500 ft) and Clto = 1.49 and wing loading of W0/Sref = 40 psf (qto = 26.9 psf , takeoff speed = 89.1 kts). We assumed nominal cruise at 300 kts at 27.4Kft, an initial cruise weight (w4) = 2645lbm and a cruise lift-to-drag ratio (LoDcr) of 22.5. • 2. From RayAD Figure 5.4, the required jet aircraft ground roll takeoff parameter for is 100where • TOP =[W0/Sref]/[CLt/o*(T0/W0)] = • = qt/o/(Bhp0/W0) • T0/W0 = qto/100 = 0.269 • - Engine size, therefore, is • T0 = 2914*0.269 = 784 lbf 18-48

  50. Design of UAV Systems c 2002 LM Corporation Propulsion TBFan performance where • 3. The TBF model also has a singularity at V=0 and is sized at V = V0 by solving for WdotA0 • Fsp0 = Fsp-gg/(1+BPR)+ Fsp-fn*(BPR/(1+BPR) • T0 = WdotA0*Fsp0 • 4. Therefore for BPR = 5.0, Fsp-gg = 90, Fsp-fn = 30 (vs. 25 at BPR = 8) and T0 = 784 lbf • WdotA0 = 784/(90/6+30*5/6) = 19.6 pps • 5. Performance at other conditions is determined using Equations 18.9-11 and V0 = 100 kts. For example, at h = 27.4 Kft, V = 300 Kts (M = 0.503) : •  = (@27.4Kft)*(1+.2M^2)^3.5 = 0.37969 • = (@27.4ft)*(1+.2M^2) = 0.8527 WdotA = WdotA0*delta/sqrt() = 8.4pps where and and 18-49

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