ME 408 Aircraft Design

Final Report for Team FSLAP

“Four-Seat Light Airplane”

Matt Mayo

Chris Hayes

Bryant Ramon

Designed in 1956 more Cessna 172 Skyhawk’s have been built than any other aircraft in

history, this is greatly due to their legendary formula of affordability, reliability and ease of

flying. Despite their legacy the design is aging and competitors such as the Cirrus SR22 boast

faster cruise speeds sleek designs. The main objective of this aircraft design is to breathe new life

into the Cessna 172 Skyhawk, updating the aircraft while retaining the ubiquitous brand.

Our hard design goals were simple, a range of 800 nautical miles with one pilot and 3

passengers in a 192 ft2 cabin; other secondary performance goals such as improved climb rate

and take off /landing field length were also considered during the design process.

The most significant design changes are the removal of wing braces, the enlargement of

the wing and the introduction on a turboprop engine. These changes allowed us to decrease the

drag and increase the thrust of the aircraft, bringing it more in line with the performance

characteristics of modern four seaters.

Overall we are content with the final design of the airplane. Our suggestion to

management is not to replace the Skyhawk but to dub this aircraft the Super Skyhawk and offer

alongside the Skyhawk as an option for buyers who are interested in more performance at an

increased cost.

Performance Quote

Requirements/ Proposed

Performance Item Targets Design Delta% Cessna 172 Cirrus SR 22 -Design Payload (Non-Expendable) 1 Pilot / 3 pax 4 n/a 4 4

-Passenger Allowance 250 lbs/pax 250lbs/pax n/a 250 250

-Cabin Length/Width/Height 12’/4’/4’ (192 ft^3) 194 ft^3 1% 158 ft^3 184 ft^3 -Design Payload (Expendable) 0 0 n/a 0 0

-Design Range w/Max Payload 800 nm 800 nm 0% 640 nm 400 nm -Design Time-on-Station w/ Payload 0 0 n/a 0 0

-Stall Speed 160 nm/hr 158.5 ktas -0.9% 124 ktas 183 ktas

-AEO Takeoff Field Length 2.5 deg/s 7.27 deg/s 190.8% 13.8 deg/s 11.11 deg/s -Service Ceiling >20,000 ft 36,900 ft 84.5% 14,000 ft 17,500 ft

-Unit Cost

Chapter 2 – Take-off Weight Estimate

Mission Analysis Summary (W/Wo) Weight Parameter Symbol Value Fraction Empty Weight (lb) We 2,094.6 0.5280

Payload (lb) Wp 1,000.0 0.2521

-Expendable Wpe 0.0 0.0000

-Non-expendable Wpne 1,000.0 0.2521

Fuel Load (lb) Wf 872.4 0.2199

-Mission Fuel Burned Wfb 823.0 0.2075

-Reserves Fuel Wr 41.2 0.0104

-Trapped Fuel Wtf 8.2 0.0021

Design Takeoff Gross Weight (lb) Wo 3,967.0 1.0000

Surplus Empty Wt. (lbs) 0.00 Table 1: Mission Analysis Summary for Chapter 2

The spreadsheet itertow.xls was used for the analysis of the takeoff weight. For our

Cessna redesign, the design drivers that affected the spreadsheet the most were Operating Range,

Cruise Altitude, Cruise Mach number, Specific Fuel Consumption, and Aspect Ratio. One of our

hard requirements was our design range with max payload. The specified range was 800 nm

(400nm operating radius). Our initial desire was to come up with a design that would make

extensive use of composite materials which would thus give us a structure factor, SFACT, of

around 0.5. Our design SFACT resulted in a value of 0.5280 after many iterations. Also resulting

from countless iteration of our spreadsheets is our L/D ratio. The L/D ratio used in our design is

14 which is smaller than the initial value of 17. This L/D is obtained from the “aero” sheet in

chapter 7.

Another hard requirement for our design was our passenger allowance (i.e. our non-

expendable payload). The specified payload included a pilot and three passengers weighing 250

lbs each.

The cruise altitude (although not an explicitly specified design driver) was determined

from our max cruise speed requirement of >160

. A decided cruise altitude of 10,000 feet

results in a cruise Mach number of 0.25 (subsonic). This Mach number results in a cruise speed

of 267.5

or 158.5

. Although the soft requirement was not met entirely it was close enough

to proceed with our other calculations. Our goal in this design is to meet as many of our soft

requirements as possible or come relatively close to each.

Since there is no combat time for our design, the TSFC and engine thrust are not used in

the takeoff weight estimate. The engine SFC is 0.7713, which comes from our engine analysis in

chapter 7. Our team allotted a 15 minute loiter time to account for the necessary time for less-

experienced pilots to position the aircraft prior to the landing approach.

For the estimate of the fuel-weight fraction used in cruise and loiter, the governing

equations were modified to account for the propeller efficiency. Our propeller efficiency

calculated from the engine spreadsheet in Chapter 7 was 0.72 (for our turboprop design). The

0.72 propeller efficiency value replaced the original typical value of 0.8.

Given all of these values and assumptions our resulting spreadsheet is shown in our

appendix under Figure A.1. Our final takeoff weight was 3,967 lbs. Of this, 1000 lbs is payload,

872.4 lbs is fuel, and 2,094.6 lbs is the empty or structure weight.

Chapter 3 – Wing Loading Selection

Wing Loading Selection Summary

Design Wing Loading W/S (lb/f^2) 17.81

No. Flight Regime Parameter Value Target Del%

1 AEO Take-off S_TO (f) 945.8 1800.0 -47%

2 Landing S_L (f) 1,119.9 1500.0 -25%

3 Cruise Start S (f^2) 222.7

4 Cruise End H (f) 10,000.0

5 AEO Climb dH/dt (f/min) 731.1 900 -19%

6 Acceleration n 7.149

7 Turn - Instantaneous psi_dot (deg/s) 7.270 2.5 191%

8 Turn - Sustained psi_dot_act (deg/s) 2.500 2.0 25%

9 Ceiling H (f) 51,240.3 20000 156%

10 Glide Gamma (deg) 2.490 3 -17%

11 Stall Speed Vstall (ktas) 47.6 45 6% Table 2: Wing Loading Selection Summary for Chapter 3

The spreadsheet wingld.xls was used for the analysis of the wing loading selection. The

wing loading spreadsheet was used to obtain many of the soft requirements for our design.

Although design range is one of our hard requirements the optimum wing loading in cruise was

not used as the design wing loading. The reasoning for this was that a smaller wing loading was

needed in order to meet or come close to a lot of our other soft requirements. Furthermore, the

specified operating radius would allow the aircraft to make cross-country flights without the

frequent refueling.

By using a smaller wing loading we were able to meet our landing and takeoff field

lengths, instantaneous and sustained turn rates, and AEO climb rate requirements. Our resulting

takeoff and landing field lengths are 945.8 ft and 1119.9 ft respectively. The resulting

instantaneous and sustained turn rates are 7.27

and 2.5

respectively. Our AEO climb rate

was 731.3

.

The value used in the takeoff and landing portion of the spreadsheet came from

our chapter 9 analysis for plain flaps. The parasite drag value is the clean parasite drag which

comes from our key outputs in the aero sheet from chapter 7. The Oswald’s efficiency value used

is one supplied by Professor Geiger which is an accepted value for e.

Lastly our calculation of thrust stemmed from our desire to reach our soft requirement for

climb. In order to get the desired climb rate we had to iterate through a number of T/W values

which caused our thrust value to change until it reached a value that allowed our climb rate to be

met.

Other soft requirements met from the wing loading spreadsheet are glide slope, service

ceiling, and stall speed. All the analysis is shown in our chapter 3 spreadsheet in our appendix

under Figure A.2

Chapter 4 – Main Wing Design

Calculations

b 40.87 ft

Meff 0.250

cr 5.45 ft

ct 5.45 ft

m.a.c. 5.4 ft

b 0.9682

CLa 0.0863 1/deg

CLo 0.1726

atrim 0.54 deg

CLtrim 0.2192

k 0.0531

CD 0.0133

L/D 16.48

Total Drag 185.9 lbf Table 3: Key outputs values from Chapter 4

The spreadsheet wingd.xls was used for the analysis of the main wing design. Because

our aircraft is a subsonic design with a low Mach number our main wing did not require a

leading-edge sweep. Our aircraft is designed to be homebuilt therefore our taper ratio input is 1.

This results in a conventional rectangular design for our wing. With an aspect ratio of 7.5 (taken

from chapter 2) and a surface area of 222.7 ft^2 (from chapter 3) we end up with a wing span of

40.87 feet.

The wing section selected for the design is the NACA 2412 which is the same airfoil the

Cessna 172S uses. The reasoning behind that airfoil was mainly our attempt to remain under the

Cessna brand by simply redesigning the Skyhawk and not making a completely new design.

Furthermore, the current model Cessna uses an interference factor of above 1 to account for wing

braces. Our design will have an interference factor equal to 1 by getting rid of the wing bracers.

The large aspect ratio and no leading edge sweep makes our wing act almost two-

dimensional. The following figures are outputs from the chapter 4 spreadsheet as well as the

airfoil data from Theory of Wing Sections. The chapter 4 figures show the wing half span vs.

axial position in one graph and the lift coefficient vs. angle of attack on the other. The airfoil data

figures are taken straight from the Theory of Wing sections book.

Figure 1: Wing Lift Curve from Chapter 4 Figure 2: Wing planform from Chapter 4

Figure 3.a and 3.b: Wing Section data for the NACA 2412

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

-4 -2 0 2 4 6 8 10 12 14 16 18

Lif

t C

oeff

icie

nt,

CL

, (-

)

Angle of Attack, a, (deg)

Wing Lift Curve 2D Lift3D Lift

0

5

10

15

20

25

0 5 10 15Win

g H

alf

Sp

an

, b

/2,

(ft)

Axial Position, (ft)

Wing Planform

All the analysis is shown in our chapter 4 spreadsheet in our appendix under Figure A.3

Chapter 5 – Fuselage Design

Viscous Drag Calculations: Elliptic Cylinder Fuselage Shape

x/L x (ft) H (ft) W (ft) P (ft) Sw(ft^2) Rex CF Drag (lbf) Volume (ft3)

0.00 0.0 1.00 1.00 3.1

0.10 3.2 2.50 2.50 7.9 25.1 4.5E+06 3.40E-03 6.4 8.2

0.20 6.4 4.50 4.50 14.1 45.2 9.0E+06 3.04E-03 10.2 31.6

0.30 9.6 4.50 4.50 14.1 45.2 1.4E+07 2.85E-03 9.6 50.9

0.40 12.8 4.50 4.50 14.1 45.2 1.8E+07 2.72E-03 9.2 50.9

0.50 16.0 4.50 4.50 14.1 45.2 2.3E+07 2.63E-03 8.8 50.9

0.60 19.2 4.00 4.00 12.6 40.2 2.7E+07 2.56E-03 7.6 45.4

0.70 22.4 2.40 2.40 7.5 24.1 3.2E+07 2.50E-03 4.5 26.3

0.80 25.6 1.40 1.40 4.4 14.1 3.6E+07 2.45E-03 2.6 9.3

0.90 28.8 0.90 0.90 2.8 9.0 4.1E+07 2.41E-03 1.6 3.4

1.00 32.0 0.50 0.50 1.6 5.0 4.5E+07 2.37E-03 0.9 1.3

Totals: 298.6 61.3 278.1

Table 4: Fuselage output calculations for Chapter 5

The fuselage shape is constructed as a series of ellipses that are blended and tapered to

allow enough room for the crew of 4 and to provide space for the engine. A fineness ratio of

0.136 was selected to minimize the total drag on the fuselage. The overall fuselage length is 32

ft. with a diameter of 4.35 ft.

Because our design is subsonic our aircraft does not experience wave drag. Therefore,

based on our inputs our calculated viscous drag for our aircraft is 61.3 lbf. The equivalent drag

coefficient, normalized by the wing area, is 0.004331. The following figure shows the fuselage

design of our aircraft.

Figure 4: Fuselage perimeter graph from Chapter 5

0

5

10

15

20

25

30

0 10 20 30 40

Fu

sela

ge

Peri

mete

r, (

ft)

Axial Position, (ft)

Fuselage Perimeter

All the analysis is shown in our chapter 5 spreadsheet in our appendix under Figure A.4

Chapter 6 – Tail Design

Tail Design Summary

Total VT TE

Drag (lbf) Del% Sweep (deg)

Conventional 56.9 Base 10.6

T-Tail 54.3 -4.6% 10.6

Cruciform 56.1 -1.5% 10.6

H-Tail 58.1 2.1% 10.6

V-Tail 53.6 -5.9% 10.6

Inverted V-Tail 53.6 -5.9% 10.6

Y-Tail 55.8 -1.9% 10.6

Twin Tail 59.8 5.1% 10.6

Control Canard 56.9 0.0% 10.6

Lifting Canard 56.9 0.0% 10.6

Main Vertical Horizontal

Wing Tail Tail

Airfoil Section NACA 63-006 NACA 63-006

Max Thickness, % 0.120 0.060 0.060

LE Sweep, deg 0.0 35.0 15.0

Aspect Ratio, - 1.300 3.000

dCL/da, 1/deg 0.0320 0.0594

Table 5: Tail Design summary from Chapter 6

The airfoil used for our horizontal and vertical tail is the same as the one by the Cessna

172 Skyhawk (NACA 63-006). With this airfoil we have a thickness-to-chord ratio of 0.06. The

CVT and CHT values for the vertical and horizontal tail were 0.040 and 0.700 respectively.

These values are based on historic data for aircraft of this type. For LHT and LVT we took

~60% of the fuselage length based on Professor Geiger’s advice and Corke page 126. That

results in LHT and LVT values of 18.0 ft. Our taper ratios for the horizontal and vertical tail

were the same at 0.5. The aspect ratios for the vertical and horizontal tail were 1.3 and 3

respectively based on Corke Table 6.5. Our aspect ratios fall within the range supplied by the

table. The leading edge sweepback angles for the vertical and horizontal tail were 40 degs and 10

degs respectively. With all these values established and inputted into the spreadsheet we are able

to calculate our tail drags. For the vertical tail the viscous drag was 11.04 lbf. The viscous drag

on the horizontal tail is 25.90 lbf. The tail planforms for both the vertical and horizontal tail are

shown in the following figures for chapter 6. All the analysis is shown in our chapter 6

spreadsheet in our appendix under Figure A.5

Figure 5.a and 5.b: V-Tail and H-Tail planform plots from Chapter 6

With the main wing, fuselage, and tail analyses done we can create a basic external shape

of the aircraft. The following figure depicts just that.

Figure 6: Basic External Shape based on our main wing, fuselage, and tail analysis

0

1

2

3

4

5

0 2 4

H-T

ail S

pa

n,

(ft)

Axial Position, (ft)

H-Tail Planform

Chapter 7 – Propulsion System Design

Engine Selection Summary Value Units

Number of Engines 1 -

Uninstalled Engine Power, SLS ISA Max 945.0 shp/eng

Reference Engine PT6A-50 -

Engine Scale Factor 1.000 -

Type of Engine Turboprop -

Ave. SFC, @ Design Cruise 0.7494 lbm/hr/shp

Engine Weight 215.6 lbf/eng

Engine Length 39.6 in

Engine Max Diameter 19.6 in

Propeller Diameter 7 ft

# of Blades 2 -

Table 6: Engine Selection Summary from Chapter 7

The engine we selected was the PWC PT6A-50 turboprop engine. The PT6A turboprop

engine is a powerhouse that offers unmatched performance, reliability and value in its class of

500 – 2,000 shaft horsepower for a wide range of applications. From our engine selection

summary you can see that the uninstalled engine power is 945 shp. Our aircraft incorporates a 7-

foot diameter, two-bladed propeller system based on the airframe size and comparison aircraft.

With these inputs we have a propeller tip Mach of 0.631 which is below the threshold of 0.85.

This chapter 7 spreadsheet is used to determine the amount of thrust tour engine-propeller system

can produce.

The total drag on our aircraft is a combination of the drag forcers on the wing, fuselage,

and tail. Based on our previous spreadsheets our total drag is 281.15 lbs. Our engine-propeller

system can produce 906.92 pounds of thrust with a power scale factor of 1.

Initially we decided on a reciprocating propeller system without a super charger but that

failed to give us the values we seek. The reciprocating propeller system with the super charger

also proved to be inefficient. Ultimately we settled on the turboprop as the happy medium. The

following figure from Chapter 7 shows the relationship between the various propeller systems

and drag as a function of Mach number.

Figure 6: Relationship between the various propeller systems and drag as a function of Mach number

The corrected static thrust value that will be used in the next chapter is 2749 lbs. All the

analysis is shown in our chapter 7 spreadsheet in our appendix under Figure A.6

Chapter 8 – Takeoff and Landing Analysis

Takeoff and Landing Summary

Design Gross Weight lbf 3967

Altitude ft 0

Engine Thrust lbf/eng 906

Stall Speed nm/hr 47.7

Field Lengths

AEO Takeoff ft 2,021

OEI Takeoff ft #NAME?

Landing ft 2,231

Regulatory Compliance

AEO, Gear Up Vy ft/min 704

OEI, Gear Up Vy ft/min 0

OEI, Gear Up G % 0.0%

OEI, Gear Down G ft/min 0.0%

Table 7: Takeoff and Landing Summary from Chapter 8

0

200

400

600

800

1000

1200

1400

1600

1800

0.00 0.10 0.20 0.30 0.40 0.50

Th

rust

an

d D

rag

(lb

f)

Mach Number, (-)

Cruise Thrust and Drag Drag

Recip w/o S/C

Recip with S/C

Turboprop

The purpose of the Chapter 8 spreadsheet is to get accurate updated values for takeoff

and landing field lengths. Apart from just updating the values obtained from Chapter 3, the

chapter 8 spreadsheet introduces new parasite and lift coefficients due to flaps. A further

understanding of the flap calculation will be done in the next section. The inputs necessary for

this spreadsheet are wing aspect ratio, wing area, takeoff and landing weights, and static engine

thrust. The overall drag coefficient was taken as the sum of the coefficients for wing, fuselage,

and tail obtained in previous spreadsheets. The increase in drags due to flaps was specified to be

0.08625. The rolling friction coefficient corresponds to a grass field which is consistent with our

recreational aircraft. For takeoff, a climb angle of 5 degrees was used with an obstacle of 50

feet to meet FAR Part 25 regulatory compliances. Our calculated takeoff field length was 2145.5

ft which doesn’t meet our soft requirement but it pretty close. We would have to consider a

higher climb angle of around 8 degrees in order to reach our desired field length. The change in

climb angle will affect our thrust which we will need to adjust.

For landing, a descent angle of 8 degrees was used with all other parameters

remaining the same except for the updated weight for landing. Our final landing field length

resulted in 1819 ft. although our field length for landing is larger than what we expected it is still

lower than our take off length which we like. The following figures show the breakdown for both

takeoff and landing. All analysis was done using the Chapter 8 spreadsheet which can be found

in our appendix under Figure A.7

Figure 7: Take-off and Landing Breakdowns from Chapter 8

1 60% 2

14%

3 12%

4 14%

Take-off Breakdown

1 19%

2 13%

3 18%

4 50%

Landing Breakdown

Chapter 9 – Enhanced Lift Design

Flap Design Summary

Design Units

Type of TE Flaps plain -

LE Flaps No -

Flap Area / Wing Area, Swf/Sw 0.50 -

Flap Deflection Angle, df 40.00 deg

Flap Chord / Wing Chord, cf/c 0.40 -

Flap Span / Wing Span, bf/b 0.50 -

CL,max 2.31 -

DCDo, flaps 0.0863 -

Table 8: Flap design summary for Chapter 9

For our aircraft, plain tailing-edge flaps were used. Our flaps have an angle of deflection

of 40 degrees with a length of 40% the wing chord covering 50% of the wing span. Most of the

vital inputs in this spreadsheet were taken from the wing design spreadsheet (chapter 4). The

aspect criterion designated for the wing to be in the “high” category. That meant the low aspect

ratio basic wing results are to be ignored.

All relevant analysis can be seen in the spreadsheet corresponding to this chapter found in

our appendix under Figure A.8. The following figures are taken from that spreadsheet and show

the wing planform and wing lift curve for the flaps.

Figure 8: Wing Planform for flaps Figure 9: Wing Lift Curve of main wing w/ and w/o flaps

0

3

6

9

12

15

0 3 6 9 12 15

Win

g H

alf

Sp

an

, b

/2,

(ft)

Axial Position, (ft)

Wing Planform

0.0

0.5

1.0

1.5

2.0

2.5

-40 -20 0 20 40

Lif

t C

oeff

icie

nt,

CL

, (-

)

Angle of Attack, a, (deg)

Wing Lift Curve w/oFlaps

Chapter 10 – Material Selection

Results

Parameter Symbol Units Value

Max Maneuver Load Factor nmax maneuver - 2.161

Max Gust Load factor nmax gusts - 3.754

Design Load Factor ndesign - 5.631 Table 9: Load Factor summary for Chapter 10

The material selection spreadsheet mainly focused on load factor, shear forces, and

bending moment distribution on the wing and fuselage. The calculation of the load factor for the

turning phase (both instantaneous and sustained) rely on the input Mach number and turn rate.

Because our turn rates were a soft requirement we focused on that load factor calculation. The

acceleration and intercept phase load calculations don't correspond to our aircraft. We added a

safety factor of 1.5 to our load calculation and used that as our design load. We set our ranges for

min and max load factors to -2 and 4 based on the Corke table. In the following figure you will

see how our maneuver loads and gust loads fall within those ranges.

Figure 10: V-n diagram for maneuvering and gust for Chapter 10

The inputs for the wing loads and fuselage loads sheets were taken directly from previous

worksheets in order to calculate the moments and shear on the wing and fuselage. The following

figures show the loads, shear, and moment graphs for both the wing and fuselage.

Figure 11a, 11b, 11c: Wing Load, Wing Shear, and Wing Moment for Chapter 10

Figure 12a, 12b, 12c: Fuselage Load, Fuselage Shear, and Fuselage Moment for Chapter 10

For the material selection of the wing spar and fuselage and longeron, the width, height,

thickness, and material used were inputs. After much iteration, our design will incorporate a

wing spar of width 2.06 in, height 7.00 in, and 0.51 in thickness. Our fuselage longeron will be 4

in. in width, 2 in. in height, and 1 in thickness. The skin material for both will be 2024-T3

aluminum alloy. The core material for the spar and longeron is 4130 normalized steel alloy and

2024-T3 aluminum alloy respectively. The following figures show the cross sectional area of

our spar and longeron.

Figure 13 a, 13 b: Cross section for the Wing Spar and Fuselage Longeron

Chapter 11 – Stability and Control

Weight Summary General

Component Symbol Fighter Transport Aviation

Wing Wwing 631 512 487

Horizontal Tail Wh-stab 67 18 26

Vertical Tail Wv-stab 20 41 26

Fuselage Wfuse 794 888 303

Main Gear Wmain lg 112 77 180

Nose Gear Wnose lg 59 30 32

Engine(s) Weng 553 553 595

Remaining Components Wrem 674 674 555 Empty Weight We 2,910 2,793 2,204

Design Gross Weight Wo 3,967 3,967 3,967 Empty Weight Fraction We/Wo 0.734 0.704 0.556

Table 10: Weignt summary for Chapter 11

For all of the component weights, general aviation was considered the appropriate

category for the weight estimate. All relevant inputs such as design load factor were all taken

from previous spreadsheets 2-10. In our initial refined weight estimate we saw that our refine

weight was too big. The main components of the weight were the wing, fuselage, and landing

gear. We decided to go back and change our wing loading which is why we were unable to meet

our vstall requirement. By taking into consideration the use of composite materials in our wing

we were able to reach a target weight that we were okay with. The weight summary above shows

the breakdown of all the weight estimates for general aviation for our aircraft.

Static Stability & Control Summary Static Margin Value Comments

-Center of Lift 0.3553 xcl/L

-Center of Gravity 0.3401 xcg/L

-Static Margin @ Wcr, start 9.0% stable

-Dtrim / Dtotal 0.146 Dtrim high, See Corke Page 279

Stability Coefficients

-Longitudinal, Cm,a -0.0019 stable Corke: -1.5

Chapter 12 – Cost Estimate

Cost Estimate Summary

Year 2013 Number of Development Aircraft 2

Number of Production Aircraft 5300

Production Rate (per month) 50

Amortization Period (# of ac) 4000

Initial Unit Cost (1986 Model) $689,852

Final Unit Cost (1986 Model) $661,368

Initial Price Markup 4%

Profit (%) 10

Chapter 13 – Trade Summary

0.0

200.0

400.0

600.0

800.0

1000.0

1200.0

1400.0

0 0.5 1 1.5 2

Ra

ng

e (n

au

tica

l M

iles

)

Specific Fuel Consumption (lb/(lbf·h))

Specific Fuel Consumption vs Range (nm)

Range (nm)

760

770

780

790

800

810

820

830

840

850

860

870

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8

Ran

ge (

nau

tica

l mile

s)

Cruise Mach

Cruise Mach vs Range (nm)

Jane’s Style Datasheet

Overall Length 32 ft

Height 7 ft

Wings Span 40.5 ft

Root Cord 5.5 ft

Tip Cord 5.5 ft

Aspect Ratio 7.5

Engines

Prop

Diameter 7 ft

Wheels Wheel Base 3.12 ft

Internal Dimensions Cabin Length 12.5 ft

Width 4 ft

Height 4 ft

Areas Wing area 222.7 ft^2

Flaps Area 44.5 ft^2

V-Tail 15.2 ft^2

Rudder Area 2.3 ft^2

H-Tail 20.2 ft^2

Weights Empty 2096.4 lb

Take Off 3967 lb

Landing 3143.9 lb

Fuel 872.4 lb

Payload 1000 lb

Wing Loading Wing Loading 17.81 lb/ft^2

Performance S_TO 2021.1 ft

S_L 1394.6 ft

Rate of Climb 704 ft/min

Sevice Celing 36,900 ft

Cruise Speed 158.5 ktas

Stall Speed 47.6 ktas

Range 800 nm

3-D Solid Model

Sources for Competitor Data

Cessna.com

Cirrus.com

Jane’s All the World’s Aircrafts

Wikipedia

Cirrus SR22 Pilot’s Manual (cirrus glide slope)

Airliners.net (cirrus glide slope)

Excel Sheets (Development Costs)