design of uav systems

Design of UAV Systems

Putting it all together 24-1

Lesson objective - to show how to

Put it all togetherWith a focus on …

• The air vehicle

Objectives

Expectations - You will better understand how to approach air vehicle design

2002 LM Corporation

24-2


Putting it all together 2002 LM Corporation

How do we start - review

• Analyze the problem - What does the air vehicle have to do?- Is any information missing?

• Look at some potential solutions- What are the overall design drivers?

- Payload weight and volume- Range and endurance- Speed and propulsion type

• Pick a starting baseline • Analyze starting baseline

- Size and weight; range and endurance• Analyze the other approaches

- Compare results and select preferred baseline• Define preferred overall system

- Reasonable balance of cost, risk and effectiveness• Document results

Today

24-3



What kind of air vehicle - review

• Operates from 3000 ft paved runway (defined reqmn’t)

• Loiters over an area of interest (defined reqmn’t)- At h = 10-17Kft, 158nm-255 nm from base (derived)

- Baseline loiter time = 12 hrs, do trade study on 6 and12 hr (system engineer, team decision)

- Fly circular pattern, 2 minute turns (derived)- Maximum coverage area = 200nm x 200 nm (defined)- WAS for 10 sqm moving targets in 2 minutes (defined)

• Dashes 141 nm to target in 30 min. (derived reqm’nt)- Once per hour (follow-up customer response)- Based on WAS sensor or other information

• Images targets from 10 Kft (derived reqm’nt)• Operates in “all weather”

- 60% good weather, 30% bad but flyable, 10% terrible weather (unflyable)- This conflicts with our 100% availability assumption

24-4



Our first decision- review

• It is a very important one - What is the best propulsion cycle for the mission?

- Internal combustion (IC), turboprop (TBProp) and turbo fan (TBFan) engines can all meet baseline speed (280 kt) and altitude (10-17Kft) requirements

• We bring our team together for the decision- Speed and altitude is at the upper end of IC capability, reliability required will be a challenge for an IC engine

- TBProp is good cycle for low-medium altitude operations- TBFan is best at altitudes > 36 Kft but has best reliability

• We select the TBProp as our starting baseline and agree to evaluate a TBFan as the primary alternative- IC alternative decision will be based on size required

• Conventional wing-body-tail configuration(s) selected- Evaluate innovative concepts during conceptual design

• We document our decisions as “derived requirements”

24-5



Next decision- review

• How many engines? • Generally determined by available engine size

• The smallest number of engines will always be the lightest and lowest drag

• How big will they be?• Engine size is determined by thrust or horsepower-to-weight required to meet performance requirements

• One sizing consideration is takeoff; others are speed, acceleration and maneuver

• Initially we size for takeoff • We design for balanced field length (BFL) = 3000 ft

• Approximate BFL = 1500 ft ground roll to lift off speed, 1500 ft to stop if engine fails at liftoff

• Later we will calculate performance over the entire mission and ensure that all requirements can be met• This is what we will do today

24-6



Review – reqmn’t disconnect

• Initial system assessment assumed 100% air vehicle availability, weather now limits availability to 90%- This will affect SAR sizing (primarily)

- We assumed SAR operation 100% of the time, therefore, the SAR only needed 80% area coverage

- At 90% availability, the SAR needs to provide 89% area coverage (range increase to 102km) to achieve overall 80% (threshold) target coverage

• We decided to leave the baseline alone and finish the first design cycle before making the change?- During any design cycle, there will always be design and requirement disconnects

- If we change baseline every time we find a disconnect, we would never complete even one analysis cycle

• Orderly changes occur at the end of an analysis cycle

- Our methodology sizes the fuselage as a cylindrical center section with elliptical fore and aft bodies - The fuselage is defined in absolute and relative terms

-Fuselage equivalent diameter (Df-eq) is absolute but is iterated to assure volume required = available

-Relative variables are length to equivalent diameter ratio (Lf/Df-eq), and forebody and aftbody length ratios

- At a maximum speed of 282kts, a relatively low fineness ratio (Lf/Df-eq) can be used with minimum drag impact- We select a nominal value of 7.0 (cigar shape) to minimize wetted area (a weight and drag driver)

-If we assume the fuselage forebody length = 1Df-eq and the aftbody = 2Df-eq, center section length (Lc) ratio will be 4/7 or Lc/Lf = 0.571

Review - fuselage considerations

24-7



To get started we put payload in the fuselage center section, close to the vehicle center of gravity- It accommodates a payload weight of 720 lbm and a volume of 26.55 cuft (density = 27.1 pcf)

- It also carries some fuel (amount TBD, density = 50 pcf at “packing factor” PF = 0.8 or installed density = 40 pcf)

- And it carries airframe structure and some systems (landing gear, etc., nominal installed density = 25 pcf*)

- We assume other systems are in the fore & aftbodiesWe assume center section volume (Vc) is allocated entirely to payload at a packing factor (PF) = 0.7- Therefore, Vc required = 26.55/0.7 = 37.9 cuft

Later the spreadsheet will size for actual volume required

Review - fuselage volume

24-8



* 25 pcf is a reasonable estimate for installed electrical, mechanical systems including avionics, landing gear and engines

Review - fuselage geometry

24-9



- From simple geometry, we express fuselage center section volume in terms of fuselage center section equivalent diameter (De) and Length (Lc) or

Vc = (/4)LcDe^2 = (/4)(Lc/Lf)(Lf/Df)De^3

Vc = 37.9, Lc/Lf = 0.571 and Lf/De = 7

De = 2.29 ft, Lf = 16 ft and Lc = 9.16 ft- Assuming forebody length = De and aftbody length

= 2*De (k1 = 0.143, k2 = 0.286), L/D = 7 and w/h = 1- Fuselage wetted area (SwetF) would be = 106.3 sqft

- Knowing wetted area, we could calculate a fuselage drag coefficient = SwetFCfe/Sref (RayAD 12.5) and fuselage weight Wfuse) = SwetF*UWF - However, we will let the spreadsheet model do this for us later as part of an integrated analysis

or

where

24-10



Review - engine installation

• Simple engine installations are always best unless there are over-riding considerations• Such as high speed, stealth, thrust vectoring, etc.• Otherwise, complexity reduces overall performance

• Nacelle geometry is driven by engine installation• TBProp nacelles should be low drag, minimum length

• Our methodology models nacelles like mini-fuselages• Cylindrical center section, elliptical fore and aft bodies

• Nacelle type is defined by an input wetted area fraction (vs. a typical podded nacelle)

- 1.0 = typical podded commercial jet transport nacelle- 0.5 = nacelle attached to fuselage (e.g. Global Hawk)- 0.0 = engine buried in the fuselage (e.g. DarkStar)

• We assume a single, attached, aft mounted engine L/Dnac = 4; k1 = 0.2; k2 = 0.4; Dnac/Deng = 1.25;

nacelle Swet fraction = 0.5 Change from lesson 20

24-11



- Our design methodology sizes the wing separate from the fuselage - We have 4 primary decisions to make: size (planform area or Sref), shape (Aspect ratio or AR and taper ratio or ), sweep () and thickness ratio (t/c)

- Planform area will be determined by wing loading (W0/Sref), a primary design variable- A reasonable value for a turboprop is 30-60 psf

(PredatorB & RayAD Table 5.5)- We pick a value of 30 and later will refine the estimate

to ensure takeoff/cruise/loiter requirements are met- AR is a primary wing design variable determined by speed, maneuverability and lift-to-drag (L/D or LoD) ratio- High AR generally means high LoD (>20), low

maneuverability (a few g’s) and low speed (<350 kts)- For long endurance we select a starting value of 20

Wing considerations (expanded)

24-12



- Taper ratio () is a secondary wing design variable that drives wing drag due to lift achieved vs. a theoretical minimum (see RayAD Fig. 4-23)- A nominal value is 0.5 selected and needs no further

pre-concept design trade- Wing sweep is driven by speed, at a maximum speed of 282 kts we have no need for wing sweep

- Wing t/c has a major impact on wing weight, the higher the t/c, the lighter the wing weight- High t/c increases drag but trades favorably against

wing weight at low speed - At 282 kts we select a nominal maximum value (t/c = 0.13), it needs no further pre-concept design trades

- Of the wing design variables selected, only W0/Sref and AR need to be traded for our speed range

Wing - continued

24-13



- Another concept design wing consideration is volume available for fuel - Wing fuel volume is defined in terms of percent wing chord and span available for tankage

- Typically wing tanks start at the wing root or fuselage attachment and can extend to or near the wing tip

- For our UAV application we assume the wing tank starts at 10% span and extends to 90% span (1 = 0.1, 2 = 0.9). We estimate tank chord at 50% wing chord (Kc = 0.5) and fuel packing factor at 0.8- These initial estimates are not upper limit values- The tanks could extend from fuselage centerline to wing tip if required (1 = 0, 2 = 1) but it is unlikely that tank chord will exceed the assumed 50%

- Fuel density again is estimated at 50 pcf at PF = 0.8

Review - wing volume

Anotherchange

24-14



• During pre-concept design, our primary concern is tail type and size• We use parametric (historical) data to estimate both

horizontal and vertical tail size required• For “V-tails” we size using projected areas

• During conceptual design we will resize to ensure adequate stability and control and handling qualities

• Our geometry model defines horizontal tail area (Sht) and vertical tail area (Svt) as fractions of Sref or…

Sht = KhtSref and Svt = KvtSrefWhere for an average air vehicle

Kht ≈ .25 and Kvt ≈ .15“Average” V-tail area would be 0.39Sref

• Our UAV will use an average V-tail area fraction

Tail considerations

Anotherchange

Our aerodynamic model estimates lift and drag from geometry and input values of equivalent skin friction coefficient (Cfe) and “Oswald” wing efficiency (e)• We will assume a state-of-the art Cfe value of 0.0035

to reflect our assumption of good surface smoothness (See RayAD Table 12.3)

• Wing efficiency (e) is estimated at a value of 0.8 using parametric data for an unswept wing at AR = 20

The model uses these inputs to calculate minimum and induced drag coefficients (Cd0 and Cdi)

Lift coefficients are calculated from weight (W), Sref and flight dynamic pressure (q) where

• Cl = W/(qSref)Loiter and climb q are assumed to be at max L/D

Review - aerodynamic model

24-15



“Bottoms-up” weight estimates are based on a combination of methods • Airframe weight estimates use input unit weights and

calculated wetted or planform areas• Propulsion weight is based on T0/Weng or Bhp0/Weng• Landing gear weight (Wlg) is based on an input gross

weight (W0) fraction where Wlg = KwlgW0• “Other system” weights (Wsys) use another input weight

fraction where Wsys = KsysW0We will use nominal values from RayAD Table 15.2 adjusted for a typical turboprop UAV where• Wing unit weight (Uww) = 3.25 psf• Tail unit weight (Utw) = 2.6 psf• Fuselage/nacelle unit weight (Ufpnw)= 1.8 psf • Klg = 0.05 and Ksys (or “all-else empty” ) = 0.12

We also include an empty weight margin (5%)

Review - weight model

24-16



Anotherchange

Volume requirements are calculated while iterating bottoms-up weight and geometry • Fuel, payload, system and landing gear weights are used to

estimate fuselage and pod (if any) volume required• Fuel volume = fuel weight/( fuel densityPF ) • Payload volume = 26.55 cuft (chart 11-61)• Landing gear volume = gear weight/25 pcf• Other systems volume = other systems weight/25 pcf

Volume available is calculated by the geometry model using input estimates of useable volume per component • Nominal value = 0.7 for fuselage and pods (if any)• Nominal value for nacelles is a configuration variable

• In our baseline, we assume the nacelle is unavailable for anything except the engine, inlet and nozzle

Df-eq is adjusted to equate volume available and volume required plus 30% margin (or PF = 0.7/1.3 = 0.54)

Review - volume model

24-17



Finalchange

Our propulsion model is a simplified “cycle deck” used to represent both turboprops (TBP) and turbofans (TBF) • Engines are sized at sea level static conditions (h=0,

V=0) based on input values of thrust or power to gross weight required (T0/W0 or Bhp0/W0)

• The models predict performance at other values of altitude and speed by assuming that power or thrust vary primarily with airflow (WdotA)

Differences between TBFs and TBPs are determined by input values of bypass ratio (BPR), fan specific thrust (T0-fan/W0dotA-fan) and a reference speed (V0)• Our UAV studies will use the TBP and TBF values in

Lesson 18, chart 18.33

Review - propulsion model

24-18



Air vehicle performance is estimated using calculated values of gross weight (W0), empty weight (We or EW) and fuel weight (Wf) • The mission is calculated forward and backward

• Forward calculations use simplified performance models to estimate fuel required for engine start-taxi-takeoff, climb and cruise out to initial loiter location

• Another calculation works backward from empty weight and calculates fuel required for landing reserves and loiter, cruise back, dash from target, combat over the target (including payload drop) and dash to target

• The sum of the two subtracted from the starting fuel weight is the amount of fuel available for loiter

A Breguet endurance calculation using the pre and post loiter weights then predicts operational endurance

Review - air vehicle performance

24-19



We will define our mission to meet maximum distance requirements for each of the two mission types• WAS cruise out = 255nm at 27.4Kft

• Baseline operational endurance is 12 hr, with trade study options for 6 hr and 24 hr endurance

• Positive ID mission cruise out = 200 nm @ TBD Kft• We will size for 12 hrs over the surveillance area, including loiter and ingress/egress

The positive ID mission requires a 282 kt dash (out and back)• Based on requirement for 1 target ID per hour

3000ft balanced field lengthtakeoff and landing requirementsare assumed- Clto = 1.49, Bhp0/W0 = 0.092

Review - mission description

24-20



100 nm

158 nm

200 nm x 200 nm

255 nm

WAS mission definition

24-21



0 Engine start1 Start taxi2 Start takeoff3 Initial climb4 Initial cruise5 Start pre-strike refuel6 End pre-strike refuel

Start cruise 7 End cruise, start loiter8 End loiter9 Start ingress

10 End egress,combat11 Weapon release 12 Turn13 Start egress14 End egress, start cruise15 Start post-strike refuel16 End post-strike refuel17 End cruise18 Start hold19 End hold

NotationWAS MISSIONEngine start + taxi time = 30 minStart + taxi thrust level = 10%Takeoff (max thrust time) = 1 minClimb + cruise out distance = 255nmCruise altitude = 27.4KftCruise speed = TBDIngress/egress altitude = n/aIngress/egress speed = n/aIngress/egress dist. = 0Cruise back distance = 255 nmLanding loiter time = 1 hrLanding fuel reserves = 5%

See Lesson 21 – performance

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ID mission definition

24-23



0 Engine start1 Start taxi2 Start takeoff3 Initial climb4 Initial cruise5 Start pre-strike refuel6 End pre-strike refuel

Start cruise 7 End cruise, start loiter8 End loiter9 Start ingress

10 End egress,combat 11 Weapon release 12 Turn13 Start egress14 End egress, start cruise15 Start post-strike refuel16 End post-strike refuel17 End cruise18 Start hold19 End hold

Notation

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POSITIVE ID MISSIONEngine start + taxi time = 30 minStart + taxi thrust level = 10%Takeoff (max thrust time) = 1 minClimb + cruise out distance = 200nmCruise altitude = 10KftCruise speed = TBDIngress/egress altitude = 10KftIngress/egress speed = 282 ktsIngress/egress dist. = N282 nm

where N = number of searchesCruise back distance = 200 nmLanding loiter time = 1 hrLanding fuel reserves = 5%

Review - spreadsheet model

24-24



Configurations are defined in absolute and relative terms- Payload weight, volume and number of engines are

described in absolute terms (forebody, aftbody and length are relative to diameter)

- Fuselage diameter can be input as an absolute value or as a variable to meet volume requirements

- Aero and propulsion parameters (Cfe, e, Fsp0, f/a, etc.) are defined as absolute values

- Everything else (wing, tails area, engines, nacelles,etc.) is defined in relative terms (AR, W0/Sref, BHp0/W0, Sht/Sref, BHp0/Weng, Waf/Sref, UWW, etc.)

Missions are described in absolute terms- Takeoff times, operating radius, speed, altitude, etc

Most variables are input via worksheet Overall, some are input via worksheet Mperf- Mperf inputs are used to converge the overall solution

Overall worksheet inputs

24-25



Row DescriptionValue

08 Volume margin1.3

09 Headwind (kt) 010 Climb V/Vstall

1.2511 Loiter V/Vstall

1.113 Idle time (min)

3014 Idle power (%)

1015 Takeoff time (min) 116 Takeoff param

22017 Takeoff CL

1.518 Takeoff altitude 031 Landing loiter (min)

6032 Landing reserve

.0534 # of fuselages 135 Fuse. offset/(b/2) 036 Df (starting value)

2.2937 Lf/Df-equiv 738 Fuselage k1

.14339 Fuselage k2

.28640 Fuselage w/h 141 Forebody PF

0.742 Centerbody PF

0.743 Aftbody PF

0.746 Ln/Dn-eq 5 47 Dn-eq/Dengine

1.25

48 Nacelle k10.2

49 Nacelle k20.4

50 Nacelle w/h1.0

51 Nacelle Swet fract.0.5

52 Nac. Non-prop PF0.0

54 Number of pods 055 Pod offset/(b/2)

n/a56 Pod D-eq/Df-eq

n/a57 Pod L/D-eq

n/a58 Pod k1

n/a59 Pod k2

n/a60 Pod w/h

n/a61 Pod PF

0.065 Taper ratio

0.566 Thickness ratio

0.1367 Tank chord ratio

0.568 Tank span ratio 1


0.972 Horiz tail area

0.3973 Vert tail area 075 Skin frict coef

.003576 Oswold efficiency

0.877 Fuse drag factor

1.078 Wing drag factor

1.081 # of engines 1

82 Model Bhp0default

83 Eng Fsp90

84 Fan (prop) Fsp 585 Ref speed (kts)

5086 Bypass ratio

13387 Prop efficiency

0.888 Fuel/air ratio

default89 Engine L/D-eq

2.592 Starting W0

default93 Engine Hp0/Weng

2.2594 Eng. inst. wt. factor

1.395 Land gear fraction

.0596 System + wt.fract.

0.1297 Fuse+nac unit wt.

1.898 Wing unit wt.

3.2599 Horiz tail unit wt.

2.6100 Vert. Tail unit wt.

2.6101 Empty wt. margin

.05102 Misc. wt. Fraction

.02104 Fuel density

50105 Fuel PF

0.8106 Engine rho (unst’l)

22107 LG rho (instal)

25108 System + rho (inst’l)25109 Payload rho (inst’l)

27.12

Mperf worksheet inputs - WAS

24-26



Row Description Value4 h4 (kft) 27.4

5 h7-cruise (kft) 27.46 h7-loiter (kft) 27.47 h8-loiter (kft) 27.48 h9-10,13-14 (kft) 27.49 h11-12 (kft) 27.410 h14 (kft) 27.411 h17 (kft) 27.4 13. V-cruise 18014 V-ingress (& egress) 28215 Op dist (nm) 255 16 Ingress/egress (nm) 017 Combat (min) 019 Max climb M 0.4820 T factor (cruise&clmb) 121 T factor (op loiter) 122 T factor (ingress/combat) 123 SFC factor (cruise&clmb) 124 SFC factor (op loiter) 125 SFC factor (ingress/cmbt) 126 Drag factor 127 Airframe weight factor 128 Fus+nac Swet factor 1

Row Description Value

52 Df-equiv 0 – to iterate

2.29 – fixed Df

56 W0/Sref 3057 Fuel fraction TBD58 Additional fuel 0

61 Bhp0/W0 TBD

64 Payload retained (lbm) 72065 Payload dropped (lbm) 0

75 Aspect ratio 2076 Wing efficiency (e) 0.8

Design mission definition

Mperf worksheet inputs - ID

24-27



Row Description Value4 h4 (kft) 10

5 h7-cruise (kft) 106 h7-loiter (kft) 107 h8-loiter (kft) 108 h9-10,13-14 (kft) 109 h11-12 (kft) 1010 h14 (kft) 1011 h17 (kft) 10 13. V-cruise 18014. V-ingress (& egress) 28215. Op dist (nm) 200 16. Ingress dist.(nm) 14117. Combat (min) 0

Row Description Va

58 Additional fuel 0.0

64 Payload retained (lbm) 72065 Payload dropped (lbm) 0

75 Aspect ratio 2076 Wing efficiency (e) 0.877 Lamda 0.5

Secondary mission definition

24-28



The spreadsheet iterates the air vehicle to meet input weight, geometry,volume and propulsion requirements • Bottoms-up weights must be iterated by definition• Geometry is adjusted with each weight iteration to maintain proper fuselage-wing-tail relationships

• Engine and nacelle size is adjusted as requiredWaf/Sref and volume required/available are the variables used to converge weight and geometry during iteration• Waf/Sref is used as an input to the weight model and an output from the geometry model

• Fuselage diameter is adjusted to meet volume requiredWhen the values converge, mission model performance estimates will be valid, even though…

- Mission range may be short (or long)- Climb rate may be inadequate (even negative)- Cl may be too high (exceeding stall margins)

Initial sizing

24-28a



Civil/military certification requirements and good operating practice specify that certain speed and performance be mainatined. Typical values• Takeoff (V/Vstall 1.1)• Climb (V/Vstall 1.20)• Cruise (V/Vstall not defined)• Landing approach (V/Vstall 1.2-1.3)• Service ceiling = 100 fpm

UAVs have not yet established criteria but safety and good practice will dictate something similar• One difference will be operational loiter speed margin, to get high LoD we need to operate at V/Stall 1.1

For design project purposes, we will apply the above margins except we require enough thrust margin for 300 fpm (Ps = 5 fps)

Speed and performance margins

24-29



Performance convergence

Worksheet Mperf accepts new inputs to improve or adjust performance- Fuel fraction (FF) is adjusted to meet range and/or

endurance requirements- Bhp0/W0 or T0/W0 is adjusted to meet takeoff or rate of

climb requirements or achieve consistency (see below)- W0/Sref is adjusted to improve LoD or takeoff distance- AR and wing efficiency (e) can also be traded to

improve overall performanceThe values are adjusted by hand until a satisfactory solution is achieved• This includes ensuring adequate (and consistent, if

configurations are being compared) margins such as residual ROC, T-D and stall margin - Bhp0/W0 or T0/W0 is further iterated to achieve the

desired level of consistency

24-30



Spreadsheet demonstration

Df-equiv 3.04 3.04Waf/Sref - geom 13.25 13.25Waf/Sref 13.25 13.25b 40.6 40.6

W0 3304 3304W0/Sref 40.00 40.00

FF 0.1989 0.199

Wfuel (total) 657 657Wing fuel @ 50 ppcf 185 185Remaining fuel volume req'd (cuft) 9Fuselage center section vol (cuft) 88Pod volume (cuft) 0Wpay 707 707Wpay (dropped) 0 0

WE 1912 1913Bhp0/W0 0.121 0.121TOP Bhp0/W0 req'd 0.121 0.121Bhp0 400 400Sref 83 83Swetfpppn 210 210Swet 422 422

AR 20.00 20.00

Neng 1 1Bhp0 req'd (ea) 400 400TBP model Bhp0 (ea) 2111 2111ESF req'd 0.189 0.189Vol-eng(ea) 8.1 8.1D-eng 1.60 1.60D-nac 2.00 2.00

Column B Column C

Notional v

alues

24-31



Spreadsheet results

Engine size mismatch for WAS and ID mission - Negative Ps at 10 Kft, 282 Kt - requires Bhp0/W0

increase to 0.10Cruise speeds near LoDmax yielded best performance- 161 kts for WAS @ 17 Kft, 144 kts for ID at 10Kft

Positive ID was the driving mission- Baseline 12 hour operational endurance WAS air

vehicle sized to W0 = 3304 lbm, EW = 1912 lbm- For 12 IDs in 12 hrs, W0 = 16534 lbm, EW = 7996lbm

- Also required increased diameter fuselage (to 4.5 ft) to accommodate additional fuel required

Changing wing loading (W0/Sref) yielded little benefit- Higher: loiter and cruise speeds offset smaller wing- Lower: increased wing size offset smaller engine

Changing aspect ratio (AR) was of little benefit- Increased AR (25) yielded small weight improvement

Earlier e

xample problem

24-32



WAS concept

W0 = 3080 lbmEW = 1744 lbmAR = 20Sref = 77sqftSwet = 381 sqftPayload = 707 lbmFuel = 603 lbmPower = 373 Bhp TBPropMax endurance = 15.3 hrsMax speed = 350+ kts

39.2’

D Side

19.7’

D Side2.82’

This air vehicle can stay on station at 17Kft for 12 hours at an operating radius of 255 nm

Note – not to scale

Earlier e

xample problem

24-33



ID concept

W0 = 16534 lbmEW = 7996 lbmAR = 20Sref = 413 sqftSwet = 366 sqftPayload = 707 lbmFuel = 7660 lbmPower = 2000 Bhp TBPropMax endurance = 57.3 hrsMax speed = 350+ kts

90.9’

D Side

33.5’

D Side4.78’

This air vehicle can perform 12 IDs in 12 hours at 10Kft at an operating radius of 200 nm


Earlier e

xample problem

24-34



Parametric comparisons

During every step of the PCD process, we always test our performance estimates vs. data on known aircraft- This is essential to ensure our results make sense

Critical comparisons for our concept are defined by Breguet range and endurance equation variables- LoD, SFC and weights (airframe, propulsion and EW)

LoD comparison - We would compare our estimates to RayAD Fig 3.5 but

our vehicle is beyond Raymer’s parametric range- For AR=20, Sref = 77, Swet = 381 and Sref = 413,

Swet = 1614 our “wetted ARs” (A/[Swet/Sref]) = 4.0 and 5.1 vs. Raymer’s maximum value of 2.4

- But from the trend of the data our assessed values of LoDmax = 27-30 look slightly optimistic

- We can also compare to Global Hawk with a reported LoDmax of 33-34 at an estimated “wetted AR” 7

24-35



LoD comparison

This data shows that our model LoDmax estimate may be optimistic by about 5%- We will put a 10% multiplier on our Cdmin estimate

to correct for it (Cell B25 = 1.1)-Why do you suppose we corrected a 5% high LoD estimate by increasing minimum drag by 10%?- Could we have done it another way?

Maximum L/D trends

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35

0 2 4 6 8

Wetted AR = b^2/Swet

(L/D

)ma

x

Manned aircraftGlobal Hawk (est)

Manned aircraft data source: LM Aero data handbook

Model estimateCorrected value

24-36



SFC comparisons

SFC at 100-200 kts)

0.40

0.45

0.50

0.55

0.60

10 20 30 40 50

Altitude (Kft)

SFC

TPE331-14 (100 kts)PT6A-41 (100 kts)TPE331-14 (200 kts)PT6A-41 (200 kts)

Data source : Roskam A&P

SFC at 250 kts)

0.40

0.45

0.50

0.55

0.60

10 20 30 40 50

Altitude (Kft)

SFC

TPE331-14PT6A-41Other

Data source : Roskam A&P

* Note – turboprop SFC is defined in terms of horsepower. Our turboprop model converts horsepower to thrust and uses TSFC for performance calculations

This data shows that our model SFC estimates are within performance bounds for 2 typical turboprops*- Although 250 kts at 10Kft looks a little optimistic

This data shows that our calculated weights are low compared to regional turboprops but high compared to U-2 and Global Hawk - But the results are close enough for now

Another weight related issue is operating a high AR wing at 280 kts at low altitude (flutter and gust potential)

24-37



Weight comparisons

Airframe Weight Comparisons - (data from Roskam and Janes)

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GTOW/Sref (psf)

Waf/

Sre

f (p

sf)

Biz JetSE Piston PropME Piston PropReg TurboJet TransJet fightersMil Train

TR-1

Empty Weight Comparisons (data from Roskam and Janes)

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GTOW/Sref (psf)

EW

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sf)

Biz JetSE Piston PropME Piston PropReg TurboJet TransJet fightersMil TrainProp UAVsJet UAVs

Global Hawk

TR-1

24-38



Final comparison

GA Altair (Predator B variant)W0 = 7000 lbmEW = ?Sref = 315 sqftAR = 23.5Payload = 750 lbmFuel = 3000 lbmPower = 700Hp TPE-331-10TEndurance = 32 hrsMax speed = 210 kts

With these inputs our concept would have a 49 hour endurance at 50 Kft but require a 45% airframe weight reduction

Data comparison shows that our model estimates are reasonable, although some are probably optimistic• We have already decided to put a factor on our drag

estimates to reduce LoDmax to the data trend line• We will also should put a 10% multiplier on ingress-

egress SFC to put it in the middle of the parametric range• But we will have to wait for conceptual design to see if our

weights are optimistic or pessimisticSome people, however, will put on additional margins to ensure early estimates will be achievable• Typically 5-10% on SFC and drag and 10-20% on weight

Putting additional margins on our estimates, however, should not be necessary since our parametric data already shows they should be generally achievable- Adding more margin would be overly conservative and

negate otherwise valid design solutions

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Overall conclusions

24-40



Adjusted baseline

39.9’

D Side

19.95’

2.0’

D Side

7.9’

2.85’’


This air vehicle has 10% drag and 280 Kt SFC multipliers and can stay on station for 12 hours at 17Kft or perform 2.8 ID missions at 10Kft in 2.8 hours


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Balancing mission requirements

Since size requirements for vehicles to do the WAS and ID missions are so different, we will do a study to determine which size vehicle can do both missions at the lowest cost using the following approach:• Size WAS concepts for 6, 12, 24 and 48 hours of loiter

• ID mission performance will be a fallout• We will then calculate the number of aircraft required for

24/7 surveillance for 30 days for both missions • We will do simple weight based cost estimates

• Air frame and systems less installed propulsion: $ 200/lbm EW-Weng for ICProp (Lesson 8-45), $400 for TBProp, $800/lb for TBFan

• Payload: $5000 per pound • Engine : $150/lbm for ICProp, $700/lbm for TBProp,

$1000/lb for TBFan • Finally we will do a simple cost effectives

comparison to select our preferred size concept

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WAS sortie rate elements

In order to estimate number of aircraft required we have to perform a preliminary sortie rate analysis• See Lesson 7 (Sortie Rate) chart 10,

• We will use the maintenance and planning times in chart SRR-10 as representative values

The nominal mission ground times required are:• Maintenance and flight preparations – 180 minutes• Preflight checks - 6 minutes• Post landing checks and taxi - 25 minutes

The remaining elements of the sortie are• Engine start-taxi-takeoff - 31 minutes• Time to climb to 17 Kft – 6.7 minutes• Outbound and return cruise time - 184 minutes• WAS - 6, 12, 24 or 48 hours• Landing loiter – 60 minutes• Land – 3 minutes

Use RAND data and adjust

for UCAV vs. U

AV

• Include maintenance

time = f(f

light h

rs)

24-42a



Sortie rate elements

* http://www.rand.org/publications/MR/MR1028/

Therefore:SR(UAV) 24hours/[1.68FT + 4.9]SR(UCAV) 24hours/[1.68FT + 5.9]

TAT 3 hrs

UCAV unique

Include in flight time

24-43



ID sortie rate elements

The ID mission sortie is identical to the WAS mission except for the flight times where the spreadsheet values are:

• Time to climb to 10 Kft – 3.5 minutes• Outbound and return cruise time - 163 minutes• ID’s = 1 hour each

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Still valid

Still valid

Time required to fly a WAS sortie are:• 6 hr loiter - 496 min. + 6 hrs = 14.26 hrs• 12 hr loiter - 20.26 hrs• 24 hr loiter - 32.26 hrs• 48 hr loiter - 56.26 hrs

The number of missions an air vehicle can fly in 30 days vs. the number required, therefore, are:• 6 hr loiter - able to fly 50.5 missions vs. 480 required • 12 hr loiter - can fly 35.5 missions vs. 240 required• 24 hr loiter - can fly 22.3 missions vs. 120 required • 48 hr loiter - can fly 12.8 missions vs. 60 required

The number of flight vehicles required, therefore, are:• 6 hr loiter - 480/50.5 = 9.5 10• 12 hr loiter - 240/35.5 = 6.8 7• 24 hr loiter - 120/22.3 = 5.4 6 • 48 hr loiter - 60/12.8 = 4.7 5

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WAS coverage requirements

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problem

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WAS air vehicles required

The total number of air vehicles required are greater than the number required to meet flight requirements• We assume one air vehicle is always on standby in

case one of the flight vehicles has a problem• And we assume all vehicles vehicle under go

maintenance at rate of 3.4hrs + 0.68*Flight TimeThe total number of air vehicles required for continuous WAS mission coverage, therefore, are:• 6 hr loiter - 10 + 1 = 11• 12 hr loiter - 7 + 1 = 8• 24 hr loiter - 6 + 1 = 7• 48 hr loiter - 5 + 1 = 6 Earlier example problem

24-46



WAS air vehicle cost

At a nominal air vehicle cost of $400 per pound of empty weight and a nominal payload cost of $5000 per pound, we can calculate WAS costs as follows:

• 6 hr loiter - 12 air vehicles = $6.6M, Payloads = $38.9MTotal cost = $45.4M




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xample problem

• You should include

engines as separate

cost element

24-47



ID mission requirements

Assuming one target identification per hour, the times required to fly an ID sortie are:• 1 ID – 471.5 minutes + 1 hrs = 8.86 hrs• 2 IDs - 9.86 hrs• 4 IDs - 11.86 hrs and 8 IDs - 15.86 hrs

The number of missions an air vehicle can fly in 30 days vs. the number of IDs required are:• 1 ID – able to fly 81.3 missions vs. 720 required • 2 IDs - can fly 73.0 missions vs. 360 required• 4 IDs - can fly 60.7 missions vs. 180 required • 8 IDs - can fly 45.4 missions vs. 90 required

The number of flight vehicles required, therefore, are:• 1 ID - 720/81.3 = 8.85 9• 2 IDs - 360/73.0 = 4.93 5• 4 IDs - 180/60.7 = 2.96 3 • 8 IDs - 90/45.4 = 1.98 2

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Equivalent WAS coverage

WAS sortie equivalent IDs are:• 6 hr loiter = 1.5 IDs• 12 hr loiter = 2.8 IDs• 24 hr loiter = 5.1 IDs• 48 hr loiter = 9.5 IDs

The number of ID missions a WAS air vehicle can fly in 30 days vs. the number required, therefore, are:• 1.5 IDs - can fly 76.8 missions vs. 478.9 required • 2.8 IDs - can fly 67.7 missions vs. 261 required• 5.1 IDs - can fly 55.4 missions vs. 141.2 required • 9.5 IDs - can fly 41.4 missions vs. 75.9 required

Total number of ID vehicles required, therefore, are:• 6 hr loiter or 1.5 IDs - 478.9/76.8 = 6.2 8* • 12 hr loiter or 2.8 IDs - 261/67.7 = 3.9 5* • 24 hr loiter or 5.1 IDs - 141.2/55.4 = 2.5 4* • 48 hr loiter or 9.5 IDs – 75.9/41.4 = 1.8 3*

* If WAS and ID vehicles are identical, a 2nd back up is not required

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ID air vehicle cost

At a nominal air vehicle cost of $400 per pound of empty weight and a nominal payload cost of $5000 per pound, ID costs are:

• 1.5 IDs – 8 air vehicles = $4.2M, Payloads = $24.7MTotal cost = $29.0M

• 2.8 IDs - 5 air vehicles = $2.9M, Payloads = $14.1MTotal cost = $17.0M



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Total cost

The most cost effective single vehicle solution for both missions is an 18 hour WAS vehicle that can also perform 4 IDS• Therefore ID air vehicles launch once every 4 hours

while WAS air vehicles launch once every 18 hours for an average of 7.3 missions per day

Air Vehicles Required(initial baseline)

0

2

4

6

8

10

6 12 18 24 30 36 42 48

Operational endurance capability (hrs)

Nu

mb

er

WAS missionID mission

Air Vehicle + Payload Cost(initial baseline)

0

20

40

60

80

6 12 18 24 30 36 42 48

Operational endurance capability (hrs)

Pro

cure

men

t c

ost

($M

)

WAS missionID missionCombined

24-51



Resulting configuration


44.2’

D Side

21.2’

2.1’

D Side

8.5’

3.0’

This air vehicle can stay on station for 18 hours at 17Kft or perform 4 ID missions at 10Kft in 4 hours


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24-52



What it really looks like


This air vehicle can stay on station for 18 hours at 17Kft or perform 4 ID missions at 10Kft in 4 hours

Approximately to scale

44.2’

2.1’

8.5’

21.2’

3.0’

Looks like a ½ scale TBProp Global Hawk

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24-52



TBProp status

We have completed our first pre-concept “design cycle”• We have explored the basic concept and found that one

4000 lbm class vehicle can meet both WAS and ID mission requirements at minimum cost

• The vehicle size is reasonable and the internal volume available should accommodate the required payloads, propulsion, systems and fuel

• We have shown that the required weight, aerodynamic and propulsion performance levels are consistent with the state-of the art and should be achievable

However, we have not completed pre-concept design• We still have a requirement problem resulting from the

assumption of 100% availability vs. 90% flyable days• We also need to conduct goal vs. threshold and and

explore alternative TBProp architectures (Charts 8-59/63)And we need to evaluate alternate propulsion concepts

24-53



Alternative propulsion concepts

One of our early decisions was to compare TBFan and IC engine concepts against our TBProp baseline• But only if an IC engine of appropriate size is available

• However, the minimum size TBProp required to perform the ID mission is 420 Hp

• This minimum power required exceeds the size of the largest available IC engine

• Therefore, we can drop IC the engine from our study on the basis of size incompatibility

TBFan concept evaluation will be straight-forward with few decisions required• At the relatively low speeds and altitudes associated

with our mission, there is only one viable option• A fuel efficient high bypass ratio (BPR) engine

We select a nominal BPR=5 as being representative of high efficiency engines of this type

24-4



TBF alternative

We develop a spreadsheet model nearly identical to a TBProp, the major differences being engine definition• From PCD Review Part 1.5, PRR-14, nominal T0/Weng

= 5.5; installed thrust loss 10% (for good installation)• From PRR-26 TBFan parametric data we select a fan

specific thrust value of 25 sec for BPR = 5• From PRR-22 we select the remaining model inputs

Geometrically, the only difference will be the nacelle• TBFan nacelles are modeled as open-ended cylinders

where by definition k1n = k2n = 0• We assume nominal values of Lnac/Dnac = 4 and

Dnac/Deng = 1.25Takeoff performance will also be different, a 3000 ft balanced field length for a jet (ground roll of 1500 ft) requires a thrust based takeoff parameter of 100

24-55



Overall TBFan inputs

Col DescriptionValue

03 R-start (nm)default

17 Headwind (kt) 018 Clmax 1.219 V/Vstall

1.2525 LoD start

default26 SFC start

default27 EWF start

default28 Kttoc start

default31 Idle time (min)

3032 Idle power (%)

1033 Takeoff time (min) 134 Takeoff param

10035 Takeoff CL

1.540 Takeoff altitude 043 Landing loiter (min)

6044 Landing reserve

.0547 # of fuselages 148 Fuse. offset/(b/2) 050 Lf/Df-equiv 751 Fuselage k1

.14352 Fuselage k2

.28653 Fuselage w/h 155 Nacelle De start

default56 Ln/Dn-eq 4


57 Dn-eq/Dengine1.25

58 Nacelle k1 059 Nacelle k2 060 Nacelle Swet fract.

0.561 Nacelle w/h

1.063 Number of pods 064 Pod offset/(b/2)

n/a65 Pod D-eq/Df-eq

n/a66 Pod L/D-eq

n/a67 Pod k1

n/a68 Pod k2

n/a69 Pod w/h

n/a73 Taper ratio

0.374 Thickness ratio

0.1375 Tank chord ratio



0.978 Tank pack factor

0.878 Horiz tail area

0.3979 Vert tail area 084 Skin frict coef

.003585 Oswold efficiency

0.886 Fuse drag factor

1.087 Wing drag factor

1.0

Col Description Value90 # of engines 191 Model Bhp0 default92 Eng Fsp 9093 Fan Fsp 2594 Ref speed (kts) 10095 Bypass ratio 596 Installed T0 0.997 Fuel/air ratio default98 Engine L/D-eq n/a99 Engine density n/a101 Starting W0 default103 Engine T0/Weng 5.5104 Eng. inst. wt. factor 1.3105 Land gear fraction .05106 System wt.fraction 0.1107 Fuse+nac unit wt. 3108 Wing unit wt. 5109 Horiz tail unit wt. 3110 Vert. Tail unit wt. 3111 Empty wt. margin .05113 Misc. wt. Fraction .02

Changes from TBProp shown in red

Earlier S

preadsheet

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Mperf TBFan inputs

Col Description Value3 h4 (kft) 17

4 h7-cruise (kft) 175 h7-loiter (kft) 176 h8-loiter (kft) 177 h9-10,13-14 (kft) 108 h11-12 (kft) 109 h14 (kft) 1710 h17 (kft) 1713 Vcruise 20014 V-ingress/egress 28015 WAS op dist (nm) 255 15 ID op dist (nm) 20016 WAS dash (nm) 016 ID dash (nm) 14117 Combat (min) 019 Max climb M 0.4820 T factor (cruise) 121 T factor (loiter) 122 T factor (combat) 123 TSFC factor (cruise) 124 TSFC factor (loiter) 125 TSFC factor (combat) 126 Drag factor 1


27 Airframe weight factor 128 Fuse+nac Swet factor 1

59 Df-equiv3.04

63 Waf/Sref (input)TBD

67 W0/Sref 4068 Fuel fraction

TBD

74 Payload retained (lbm)707

75 Payload dropped (lbm) 0

77 T0/W0 TBD

78 Aspect ratio 20

Changes from TBProp shown in red

Earlier s

preadsheet

24-57



TBFan WAS concept

W0 = 4865 lbmEW = 2454 lbmAR = 20Sref = 122 sqftSwet = 517 sqftPayload = 707 lbmFuel = 1656 lbmEngine = 1299 Lbf TBFanMax endurance = 15.4 hrsMax speed = 280 kts

• This air vehicle can stay on station at 17Kft for 12 hours at an operating radius of 255 nm

• It is 41% heavier than a TBProp with the same performance

49.4’

D Side

21.3’

D Side3.04’

1.8’’

7.2’


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24-58



TBFan ID concept

W0 = 5660 lbmEW = 2761 lbmAR = 20Sref = 141 sqftSwet = 573 sqftPayload = 707 lbmFuel = 2133 lbmEngine = 1511 Lbf TBFanMax endurance = 18.8 hrsMax speed = 280 kts

49.4’

D Side

21.3’

D Side3.04’

1.8’’

7.2’

• This air vehicle can perform one ID at 10Kft at an operating radius of 200 nm

• It is 39% heavier than a TBProp with the same performance


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24-59



TBFan conclusions

The TBFan alternative is bigger and about 40% heavier than the TBProp baseline for both design missions• The relatively low-speeds and altitudes required

really are optimum for TBProp operations • TBFan cycles are better suited for higher speeds and

altitudesWe can now confidently drop the TBFan concept from further consideration• And document the results of our alternative concept

study as rationale for our future exclusive focus on TBProp engines

We will also document the rationale for selecting an 18 hour WAS capability for our preferred baseline to meet both WAS and ID mission requirements

24-60



TBProp continued

Even though we have concluded that the TBProp is the best overall solution to meet mission requirements, we still need to address some unresolved issues• The impact of 10% of the weather being unflyable vs. our

assumption of a 100% flight rate vs. the threshold requirement for 80% target coverage

• The cost effectiveness of designing for threshold vs. goal performance

• The effectiveness of alternative • See Lesson 3, charts 13-15

• The support concept required• Overall system life cycle cost

24-61



Homework

1. Using spreadsheet TBProp.AE261Example.xls and total mission procurement cost as the figure of merit, for the TBProp example, do the following trades (one trade for each team member, individual grades):• Aspect ratios (AR) of 10-20-25-30 at W0/Sref = 30• W0/Sref of 15-30-45-60 at AR = 10 • Aspect ratios (AR) of 10-20-25-30 at W0/Sref = 60• W0/Sref of 15-30-45-60 at AR = 30

2. Select best combination of W0/Sref and AR and use TBProp.AE261Example.xls to trade 12-24-48 hr WAS loiter times (team grade). Select the best loiter time and explain why it turned out that way

3. Use TBProp.AE261Example.xls to determine best WAS and ID cruise speeds. Explain why (team grade)

4. Discuss ABET issues #5 and #6 and document your conclusions (one paragraph each – team grade)

24-61



Intermission

design of uav systems

Documents