1. Saint Louis University
Parks College of Engineering, Aviation, and Technology
Humanitarian Service Aircraft
Zenghui Liu
Frank Meyer
Matthew Satcher
Keegan Smith
December 14, 2012
2. 2
Table of Contents
I. Introduction ................................................................................................................................7
A. Mission Objective....................................................................................................................7
B. Mission Requirements..............................................................................................................7
C. Mission Profile.........................................................................................................................8
II. Design......................................................................................................................................10
A. Fuel Fractions........................................................................................................................10
B. Weight Sizing........................................................................................................................12
C. Fuel Sensitivities ...................................................................................................................14
D. Wing and Power Sizing..........................................................................................................14
E. Wing Configuration...............................................................................................................20
F. Airfoil Selection ....................................................................................................................21
G. Fuselage & Interior Layout......................................................................................................23
H. Empennage Sizing and Configuration.......................................................................................24
I. Propeller Sizing......................................................................................................................26
J. Propulsion.............................................................................................................................27
K. Weight Distribution & C.G. Location........................................................................................28
L. Structures..............................................................................................................................30
M. Drag Polar.............................................................................................................................30
N. Performance................................................................................Error! Bookmark not defined.
O. Stability And Control ..............................................................................................................32
P. Risk Analysis...........................................................................................................................33
Q. Cost Estimation......................................................................................................................35
R. Competitive Comparison.........................................................................................................36
III. Appendix ..............................................................................................................................37
A. Fuel Sensitivities....................................................................................................................37
B. Wing and Power Sizing...........................................................................................................39
C. Airfoil Selection.....................................................................................................................41
D. Fuselage and Interior Layout...................................................................................................45
E. Empennage Sizing and Configuration:.....................................................................................46
F. Propeller Sizing......................................................................................................................47
3. 3
G. Weight Distribution & C.G. Location .......................................................................................48
H. Drag Polar.............................................................................................................................53
I. Performance ...............................................................................Error! Bookmark not defined.
J. Stability And Control..............................................................................................................57
K. Cost Estimation .....................................................................................................................59
I. References:...........................................................................................................................60
4. 4
Nomenclature
AR = aspect ratio
𝑏ℎ = horizontal tail span
𝑏 𝑊 = wing span
c = chord
𝑐 𝐻𝑇 = chord of horizontal tail
𝑐 𝑉𝑇 = chord of vertical tail
𝑐 𝑚 = pitching moment coefficient
𝑐 𝑛 = yawing moment coefficient
𝐶̅ 𝑊 = wing mean chord
𝐶 𝐷0
= zero drag coefficient
𝐶 𝐿 𝛼
= lift curve slope
𝐶 𝐿 𝛽
= rolling moment with sideslip
𝐶 𝑓𝑒
= equivalent skin fraction coefficient
𝐶 𝑚 𝛼
= pitching moment curve slope
𝐶 𝑛 𝛽
= yawing moment derivative with sideslip angle
𝐶ℓ = rolling moment coefficient
𝐶 𝐿 = lift coefficient
𝐶 𝐿
= lift coefficient
𝐶 𝑎𝑣𝑖 𝑜 𝑛𝑖𝑐𝑠 = avionics cost
𝐶 𝑓 𝑙𝑎𝑚 = laminar fraction drag
𝐶 𝑓 𝑡𝑢𝑟𝑏 = turbulent fraction drag
𝐶 𝑓 = chord length of flap
D = diameter
D = drag
D = fuselage structural depth
De = engine diameter
E = endurance
e = Oswald efficiency factor
FF = form factor
Fp = vertical force produced by propeller disk or inlet front face
FTA = number of flight-test aircraft
Fw = fuselage width at horizontal tail intersection
Ht = horizontal tail height above fuselage
Hv = vertical tail height above fuselage
I = moment of inertia
Iyaw = yawing moment of inertia
Kd = duct constant
Kfus = empirical pitching moment factor
Ky = aircraft pitching radius of gyration
Kz = aircraft yawing radius of gyration
L = fuselage structural length
La = electrical routing distance, generators to avionics to cockpit
Ld = duct length
Lec = length from engine front to cockpit
Lf = maximum length of fuselage or nacelle
Lf = total fuselage length
5. 5
Lm = extended length of main landing gear
Ln = extended nose gear length
Ls = single duct length
Lsh = length of engine shroud
Lt = tail length; wing quarter-MAC to tail quarter-MAC
Ltp = length of tailpipe
M = Mach number
Mcg = moment at center of gravity
Mmax = engine maximum Mach number
N = rotation rate obtained from engine
Nc = number of crew
Nci = number of crew equivalent
Nen = number of engines
Neng = total productivity times number of engines per aircraft
Nf = number of functions performed by controls
Ngen = number of generators
Nl = ultimate lending load factor
Nlt = nacelle length
Nm = number of mechanical functions
Nmss = number of man gear shock struts
Nmw = number of main wheels
Nnw = number of nosewheels
Np = number of personnel on board
Ns = number of flight control systems.
Nt = number of fuel tanks
Nu = Number of hydraulic utility functions
Nw = nacelle width
Nz = ultimate load factor
P = power
q = dynamic pressure
Q = five years production quantity
R = range
R = Reynolds number
Rkva = system electrical rating
𝑆 𝐻𝑇 = area of horizontal tail
𝑆 𝑉𝑇 = area of vertical tail
𝑆 𝑊 = wing area
Scs = total area of control surface
Scsw = control surface area
Se = elevator area
Sf = fuselage wetted area
SFC = engine specific fuel consumption
Sfw = firewall surface area
Sht = horizontal tail area
SM = static margin
Sn = nacelle wetted area
Sr = rubber area
Sstall = stall speed
Svt = vertical tail area
Sw = trapezoidal wing area
t/c = thickness to chord ratio
6. 6
𝑇𝑡𝑢𝑟𝑏𝑖𝑛𝑒 𝑖𝑛𝑙𝑒𝑟 = turbine inlet temperature
T = thrust
TDPF = tail damping power factor
TDR = tail damping ratio
Te = thrust per engine
Tmax = engine maximum thrust
URVC = unshielded rudder volume coefficient
V = velocity
Vi = integral tanks volume
Vp = self-sealing “protected” tanks volume
Vpr = volume of pressurized section
Vt = total fuel volume
𝑊𝑒
= empty weight
W = total fuselage structural width
Wc = maximum cargo weight
Wdg = flight design gross weight
Wec = weight of engine and contents
Wen = engine weight
Wf = final weight
Wf = maximum width of fuselage or nacelle
Wfw = weight of fuel in wing
Wi = initial weight
Wl = landing design gross weight
Wpress = weight penalty due to pressurization
Wuav = uninstalled avionics weight
( 𝑥 𝑐⁄ ) 𝑚 = counterclockwise location of the airfoil maximum thickness location
X = location
Δ𝐶 𝐷0 𝑓𝑙𝑎𝑝
= flap drag
Λ = wing sweep at 25% MAC
𝛼 = angel of attack
Δ𝛼0𝐿 = the reduction of zero lift angle due to flap
𝛿𝑓𝑙𝑎𝑝 = flap angle in degree
𝛾 = climb angle
𝜆 = taper ratio
𝜇 = relative density parameter
𝜂 𝑝 = propeller efficiency
𝜌 = density of air
𝜖 = tail angle of attack
𝜙 = bank angle
7. 7
I. Introduction
Throughout the world, there are many flourishing societieswhich have all ofthe supplies and
amenities that one can dream of. However, along with these societies are those that have little and
the people who live in these places struggle every day to get by. Not only are the se societies present,
but they are ever abundant throughout the world and do not seem to be vanishing anytime in the
near future. From Asia to South America, anonymous villages and cities are seeking for needs such
as food, shelter, and healthcare. The only reliefseen by these third-world families arrives through
the efforts made by humanitarian societies and missionaries.
However, these villagesand cities are usually not known among the simplest places to reach.
Because ofthe lack offunds to build proper airports and roads, supplies rarely reach the places of
greatest need. This is one ofthe biggestobstaclesmissionaries have to overcome in their quest to
help the aforementioned peoples. From these difficulties,the need for a rugged, reliable transport
vehicle arises.
Nowadays, the market is stocked full ofoff-road trucks and other ground transportation vehicles
that can provide a means to deliver cargo across tough to reach areas. Although there are indeed
many ground vehicles to accomplish this mission, the field ofreliable air transportation is
beginning to become a necessity. These are machines that can not only take offfrom paved
runways, but also from fields and minimally maintained airstrips. This is where the Humanitarian
Service Aircraft or HS Aircraft enters the picture.
A. Mission Objective
The mission objective is to build an airplane which could be used by one of the mission
outreach organizations in order to support a community in need. The aircraft will be focused on
cargo and passenger transportation. Due to the likelihood that the community being supported
will have rough terrain and unpaved roads, the airplane will need to be able to land and take-off
in uncommon conditions. Range will also be an important consideration, as the pilots will likely
need to travel long distances in order to reach the communities. Finally, the aircraft will need to
accommodate multiple passengers in order to both afford technical and medical experts the
opportunity to reach these communities and to allow the opportunity for non-emergency medical
transports.
B. Mission Requirements
Because the aircraft will need to take off and land in uncommon conditions, the most
important requirement of the HS aircraft is the ability to meet STOL (short take-off and landing)
distances. To meet this requirement, the HS aircraft will be designed to take off and land on a
1500 ft. runway, while also clearing a 50 foot obstacle located at the end of said runway.
Because the aircraft is meant to carry supplies or passengers to and from remote locations, the
HS aircraft will be able to support a crew member of 175 lbs. and a total payload of 1600 lbs.
8. 8
The 1600 lbs. may be divided between the maximum passenger capacity, set at 9 people
weighing 175 lbs. each, or the payload may be used for cargo.
C. Mission Profile
Though designing a humanitarian aircraft may seem like a new idea, there are actually some
companies out there who are building aircraft for this exact purpose. One of these companies is
the Quest Aircraft Company. From one aircraft, in particular, the inspiration for the HS aircraft
was drawn. The Quest Kodiak is an aircraft used for humanitarian missions and also as a small
business cargo transport. It was because of this aircraft that the Quest Aircraft Company was
sought out to join the HS aircraft team in an attempt to design a new model. Most of the initial
estimates of the HS aircraft stemmed from the data sharing between the Quest Company and the
HS team. One of the largest contributions of data from Quest came in the form of a mission
profile.
Figure I-1) Quest Kodiak
Though Quest did not set the mission profile in stone, the company managed to provide basic
range and altitude requirements from data arranged from their previous experience with other
mission groups. From this data, a mission area was established. Using Bogotá, Colombia as a
possible mission base to determine a reasonable mission profile, the HS aircraft will have a range
of 800 NM (about 920 miles), 400 NM to the mission destination and 400 NM back to Bogotá.
As seen from Figure I-2 below, this range will provide the HS aircraft with plenty of
opportunities to reach the outermost parts of Colombia.
9. 9
Figure I-2) 400 NM Radius about Bogotá, Colombia
The rest of the mission profile includes both travel distances, to and from the mission
destination. As Figure I-3 shows below, the HS aircraft will cruise at an altitude of 15,000 ft.
while maintaining a maximum service ceiling of 25,000 ft. As research has shown, the Andes
Mountains average a height around 14,000 feet, while the highest peak in the Andes is 22,000
feet. The Andes Mountains can be seen in the topographic portion of the map in Figure I-2.
These mountains stretch across a large portion of the western coast of Southern America, making
the altitude chosen for the mission profile acceptable for a large number of missions which could
occur in this part of the world.
It should also be noted that the mission destination prominent in the mission profile is the
location where the aircraft is flying to in order to drop off cargo or bring professional teams of
engineers, doctors, or other qualified specialists. It was assumed that there would be no
opportunity to refill at these sites, which is why the arrival at the mission destination is only half
of the entire mission profile.
10. 10
Figure I-3) Humanitarian Service Aircraft Mission Profile
II. Design
A. Fuel Fractions
One of the first steps was to find the fuel fractions for the aircraft. This was done for both the
single engine and the twin engine case. The leg fuel fraction values shown in Table II-1 are shown for
each leg, with each fraction being the amount of fuel used during that leg in comparison to the total
amount of fuel. The values for the Engine Warm-up, Taxi, Take Off, Climb, Descent,and Landing legs
were all found in Roskam2
Table 2.1.
Table II-1) Leg Fuel Fractions (Leg vs. Total)
Single Engine Twin Engine
Phase
Leg Fuel
Fraction Phase
Leg Fuel
Fraction
Engine Warm-up 0.995 Engine Warm-up 0.992
Taxi 0.997 Taxi 0.996
Take Off 0.998 Take Off 0.996
Climb 0.992 Climb 0.990
Cruise 0.915 Cruise 0.917
Loiter 0.984 Loiter 0.983
Descent 0.993 Descent 0.992
Landing 0.993 Landing 0.992
11. 11
The leg fuel fractions for the Cruise and Loiter legs were found using the Bregeut Range and
Endurance equations, as well as Roskam2
equations 2.9 and 2.11. Some initial assumptions were
necessary for these calculations, as found in Roskam2
Table 2.2 for single and twin engine propeller
driven aircraft. They can be seen below in Table II-2 and Table II-3. Note that the ranges for cruise in
Table II-2 are assumed to be 400 NM, half of the actualrange. This is because the mission profile dictates
that the aircraft lands after the first 400 NM, and then repeats the trip from warm-up to landing for the last
400 NM.
Table II-2) Range Assumptions
Single Engine Twin Engine
cp (lb/hr) ηp L/D Rcruise (NM) cp (lb/hr) ηp L/D Rcruise (NM)
0.6 0.8 9 400 0.6 0.82 9 400
Table II-3) Endurance Assumptions
Single Engine Twin Engine
cp (lb/hr) ηp L/D Eltr (hr) Vltr (mph) cp (lb/hr) ηp L/D Eltr (hr) Vltr (mph)
0.6 0.7 11 0.5 157 0.6 0.72 10 0.5 157
The Cruise and Loiter fuel fractions were then calculated with the following equations:
𝐿𝑒𝑔 𝐹𝑢𝑒𝑙 𝐹𝑟𝑎𝑐𝑡𝑖𝑜𝑛 𝑐𝑟𝑢𝑖 𝑠 𝑒 = 𝑒
−𝑅 𝑐𝑟𝑢𝑖𝑠𝑒
375∗
𝐿
𝐷
∗
𝜂 𝑝
𝑐 𝑝 [II. 𝐴.1]
𝐿𝑒𝑔 𝐹𝑢𝑒𝑙 𝐹𝑟𝑎𝑐𝑡𝑖𝑜𝑛 𝑙 𝑜𝑖𝑡𝑒𝑟 = 𝑒
−𝐸𝑙𝑡𝑟
375∗
𝐿
𝐷
∗
𝜂 𝑝
𝑐 𝑝
∗𝑉𝑙𝑡𝑟
[II. 𝐴.2]
Finally, the total fuel fractions were found and are shown in Table II-4. These were found by
multiplying the leg fuel fraction by the total fuel fraction of the previous leg. The Mission Fuel Fraction
was thus the landing fuel fraction.
12. 12
Table II-4) Total Fuel Fractions
Single Engine Twin Engine
Phase
Total Fuel
Fraction
Phase
Total Fuel
Fraction
Engine Warm-up 0.995 Engine Warm-up 0.992
Taxi 0.992 Taxi 0.988
Take Off 0.990 Take Off 0.984
Climb 0.982 Climb 0.974
Cruise 0.899 Cruise 0.893
Loiter 0.884 Loiter 0.878
Descent 0.878 Descent 0.871
Landing 0.872 Landing 0.864
Mission Fuel
Fraction
0.872
Mission Fuel
Fraction
0.864
B. Weight Sizing
Having calculated the fuel fractions, the weight sizing process was next,following Roskam’s
method. The take-off weight of the aircraft can be calculated as shown in Equation II.B.1.
𝑊𝑇𝑂 = 𝑊𝑓𝑢𝑒𝑙 + 𝑊𝑒𝑚𝑝𝑡𝑦 + 𝑊𝑡𝑓𝑜 + 𝑊𝑐𝑟𝑒𝑤 + 𝑊𝑝𝑎𝑦𝑙𝑜𝑎𝑑 [II. 𝐵. 1]
The weight of the crew and the payload weight were already known and the unusable fuel weight
was assumed to be negligible. The weight of the fuel was determined using the mission fuel fraction
through the following equations:
𝑊𝑓𝑡𝑜_𝑠𝑖𝑡𝑒
= 𝑊𝑇𝑂 ∗ (1 − 𝑀 𝑓𝑓) [II. 𝐵. 2]
𝑊𝑓𝑓𝑟𝑜𝑚_𝑠𝑖𝑡𝑒
= (𝑊𝑇𝑂 − 𝑊𝑓𝑡𝑜_𝑠𝑖𝑡𝑒
−
1
2
𝑊𝑝𝑎𝑦𝑙 𝑜 𝑎𝑑)∗ (1 − 𝑀 𝑓𝑓) [II. 𝐵. 3]
𝑊𝑓𝑟𝑒𝑠𝑒𝑟𝑣𝑒
= 0.1 ∗ (𝑊𝑓𝑡𝑜 𝑠𝑖𝑡𝑒
+ 𝑊𝑓𝑓𝑟𝑜𝑚 𝑠𝑖𝑡𝑒
) [𝐼𝐼. 𝐵. 4]
𝑊𝑓 = 𝑊𝑓𝑡𝑜 _𝑠𝑖𝑡𝑒
+ 𝑊𝑓 𝑓𝑟𝑜𝑚_𝑠𝑖𝑡𝑒
+ 𝑊𝑓𝑟𝑒𝑠𝑒𝑟𝑣𝑒
[II. 𝐵. 5]
The weight of the fuel from the site is calculated with the assumption that half of the payload has
been left at the site. The reserve fuelis assumed to be ten percent of the fuel used during the mission.
Equation II.B.5 shows that the total fuel is the combination of the fuel used during the entire mission and
the reserve fuel.
13. 13
In order to account for the empty weight, the below equation was used from the Roskam2
book
(Equation 2.16) which makes an approximation for the empty weight based on previous aircraft of similar
configuration.
𝑊𝑒𝑚𝑝𝑡𝑦 = 10
(
log( 𝑊𝑇𝑂 )−𝐴
𝐵
)
[II. 𝐵. 6]
Where A and B were given empty weight parameters from Table 2.15 in Roskam2
and are shown
below in Table II-5.
Table II-5) Empty Weight Parameters
A B
Single Engine -0.144 1.1162
Twin Engine 0.0966 1.0298
In order to find the appropriate take-off weight for both the single engine and twin engine airplanes,
the take-off weight was assumed to be a certain weight. The empty and fuel weights were then found with
this initial guess through equations II.B.2 and II.B.3. The used fuel weight needed to be calculated for two
legs, the flight to the site and the flight back from the site.
Then the fuel, empty, and crew weights were subtracted from the take-off weight. If the resulting
value, which can be seen from equation II.B.1 to be the cargo weight, was not very close to 1600 lbs, the
take-off weight was changed until this relationship was satisfied. From this iterative process,the
following final weight values were found:
Table II-6) Engine Weights
Single Engine Twin Engine
Weight (lbs) Weight (lbs)
WTO 8000 WTO 14940
Wf,used 1820 Wf,used 3680
Wf,reserve 182 Wf,reserve 368
Wf 2000 Wf 4050
Wempty 4225 Wempty 9115
Wcrew 175 Wcrew 175
Wpayload 1600 Wpayload 1600
14. 14
C. Fuel Sensitivities
A study was done on the sensitivity of the weight in comparison to changes in different parameters.
The equations for these studies can be found in Appendix A, but the results of the sensitivity studies are
shown below.
Table II-7) Single Engine Fuel Sensitivities
To Site From Site
∆W
(lbs)
∆W
(lbs)
∆R = 1 NM 5.91 ∆R = 1 NM 5.70
∆E = 10 min 145 ∆E = 10 min 139
∆cp = 0.05 lbs/hp/hr 197 ∆cp = 0.05 lbs/hp/hr 190
∆ηp = 1% --- -29.6 ∆ηp = 1% --- -28.5
∆L/D = 1 --- -263 ∆L/D = 1 --- -253
Table II-8) Twin Engine Fuel Sensitivities
To Site From Site
∆W
(lbs)
∆W
(lbs)
∆R = 1 NM 22.0 ∆R = 1 NM 24.3
∆E = 10 min 590 ∆E = 10 min 651
∆cp = 0.05 lbs/hp/hr 733 ∆cp = 0.05 lbs/hp/hr 808
∆ηp = 1% --- -107 ∆ηp = 1% --- -118
∆L/D = 1 --- -977 ∆L/D = 1 --- -1078
D. Wing and Power Sizing
With the take-off weight having been determined it was now time to begin calculating initial
values for the wing area, take-off power, and maximum lift coefficients. Aircraft are designed
around the following performance objectives: stall speed, take-off field length, landing field
length, cruise speed, FAR climb rates, and time to climb. From the resultant data the
combination of the highest possible wing loading and lowest possible power loading that meet
the performance requirements was selected. This was done in order to achieve the lowest
possible weight and cost.
15. 15
The first step in sizing the wing was to select a stall speed and approximate lift coefficients
for clean, take-off, and landing configurations. FAR 23 states that single engine aircraft may not
have a stall speed of greater than 61 knots at take-off weight. A stall speed of 60 knots was
selected in order to ensure the requirement was met. Lift coefficients were then selected based
on Roskam2 Table 3.1. With a stall speed and lift coefficient now selected the take-off wing
loading was calculated using the below equation.
𝑉𝑠 = {2( 𝑊/𝑆)/𝜌𝐶 𝐿 𝑚𝑎𝑥
}
1/2
[II.D.1]
Take-off performance requirements were decided based on the various situations that are
likely to be encountered while preforming humanitarian missions. The aircraft must take-off
from an altitude of 5,000 feet, from field length of 1,500 feet while clearing a 50 foot tall
obstacle at the end of the runway. Using Eqs. [III.B.1] through [III.B.3], located in Appendix B
the following take-off parameters were calculated:
Table II-9) FAR 23 Take-off Parameters
WTO (lbs) 8000
sTO (ft) 1500
sTOG (ft) 904
(W/S)TO (SL, flaps up) (psf) 36.2
(W/S)TO (SL, flaps 1/2) (psf) 25.6
(W/S)TO (5,000 ft, flaps up) (psf) 31.2
(W/S)TO (5,000 ft, flaps 1/2) (psf) 22.1
TOP23 (SL) (lbs2
/ft2
hp) 145.6
TOP23 (5,000 ft) (lbs2
/ft2
hp) 125.4
S (SL, flaps up) (ft2
) 220
S (SL, flaps 1/2) (ft2
) 310
S (5,000 ft, flaps up) (ft2
) 255
S (5,000 ft, flaps 1/2) (ft2
) 360
P (flaps up) (hp) 1045
P (flaps 1/2) (hp) 670
b (ft) 53.9
16. 16
Next, the aircraft was sized for landing. The two main variables that determine the landing
distance required are the landing weight and approach speed. Landing weight was determined
based on a ratio of landing to take-off weight. Using Roskam2 Table 3.3 a ratio of 0.92 was
established. Equations [III.B.5] through [III.B.8] were used to calculate the below landing
parameters. Correlation between landing wing loading and maximum landing lift coefficient was
then established using Eq. [II.D.1].
Table II-10) FAR 23 Landing Parameters
WL (lbs) 7360
sL (ft) 1500
sLG (ft) 774
VsL (flaps down) (kts) 54
VA (kts) 70.3
CLmaxL (flaps down) 2.7
(W/S)L (flaps down) (psf) 38.2
S (ft2
) 193
b (ft) 39.3
In order to size the aircraft to climb requirements an initial estimation of the wetted
surface area must first be made. This was done using Eq. [III.B.9] in conjunction with Roskam2
Table 3.5. An equivalent skin friction coefficient was then estimated using Roskam2 Fig. 3.21a.
The equivalent skin friction, seen in Table II-11 below, can be found from a combination of Eq.
[III.B.10] and Roskam2 Table 3.4.
Table II-11) FAR 23 Climb Parameters
Swet (ft2
) 1330
f (ft2
) 154
CDo 0.12
Assuming a parabolic drag polar, a drag coefficient of:
𝐶 𝐷 = 0.116 + 0.047𝐶 𝐿
2
was found, for a clean configuration, by assuming an aspect ratio of 10, a Oswald’s efficiency
factor of 0.85, and using Eqs. [III.B.11] and [III.B.12].
Now, FAR 23 climb requirements need to be taken into account. There are two FAR 23
climb requirements that apply to a single engine propeller driven aircraft. The first is FAR 23.65
17. 17
which states that all aircraft must have a minimum climb rate at sea level of 300 fpm and a
steady climb angle of at least 1:12 with the flaps in the take-off position. It was assumed that
flaps in the take-off position will increase the zero lift drag coefficient by 0.015. Using Eqs.
[III.B.13] through [III.B.17] yielded the following:
Table II-12) FAR 23.65 Climb Parameters
RCP (HP/lbs) 0.009
CDo (TO flaps) 0.13
(CL
3/2
/CD)max 8.99
CL (RC max) 2.61
CD (RC max) 0.52
(L/D)max climb 5.56
CGRP 0.16
(W/P)(W/S)1/2
86.4
P (hp) 435
The second FAR 23 climb requirement that applies is FAR 23.77 which states that the steady
climb angle of an aircraft must be at least 1:30 with the flaps in the landing position and the landing gear
extended. Flaps in the take-off position were assumed to increase the zero lift drag coefficient by 0.60.
Once again, Eqs. [III.B.13] through [III.B.17] were used; When combined with Eq. [III.B.18] they
resulted in the climb parameters below.
Table II-13) FAR 23.77 Climb Parameters
CDo (L flaps) 0.17
(CL
3/2
/CD)max 7.96
CL (RC max) 3.1
CD (RC max) 0.7
(L/D)max climb 4.4
CGRP 0.15
(W/P)(W/S)1/2
25.7
P (hp) 1577
Next, the aircraft needed to be sized to its service and absolute ceiling. An absolute ceiling of
30,000 feet along with a service ceiling of 25,000 feet was chosen based on mission profile demands. A
rate of climb at service ceiling was taken to be 250 fpm in order to ensure the min required climb rate
would be surpassed. Equations [III.B19] through [III.B.21],along with those used previously to size
climb, gave the following:
18. 18
Table II-14) Ceiling Parameters
(L/D)max ceiling 6.79
RCo (ft/min) 1500
tcl (min) 11.2
CL (max Ceiling) 2.73
CD (max ceiling) 0.46
CGRP 0.15
(W/P)(W/S)1/2
16.6
S (ft2
) 4166
P (hp) 1947
b (ft) 182.56
The last set of performance requirements to calculate is sizing to cruise speed. A cruise speed of
200 knots at an altitude of 15,000 feet was selected in order to use the least fuel while still traveling with
prudence as humanitarian service aircraft are commonly used to provide emergency services such as
medical evacuations. Power index was estimated using Roskam2
Fig. 3.30. By assuming a propeller
efficiency of 0.85 Eqs. [III.B.22] through [III.B.24] yielded the following cruise parameters.
Table II-15) Cruise Speed Parameters
Vcruise (kts) 200
Ip (from fig 3.29) 1.6
CD (cruise) 0.05
CDo (cruise) 0.09
W/S (psf) 67
S (ft2
) 120
P (hp) 308
b (ft) 30.9
All of the performance parameters that effect wing and power size are accounted for. The final
step was matching all of the sizing requirements. This was done by writing a MATLAB program, shown
in Appendix D, to take all of the different power loading parameters and graphing them against wing
loading. The program was used to generate Fig. II.1 shown below.
19. 19
Figure II-1) Performance Parameter Matching Results
Looking at Fig. II-1 above the design point was selected in order to maximise wing loading
and minimise power loading while still meeting the performance requirements. This will yeild a
design that is as light as possible and as inexpensive as possible. Table II-16 below shows the
resultant design.
Table II-16) Matching Results
WTO (lbs) 8000
CLmax (flaps up) 1.9
CLmax TO (flaps 1/2) 2.1
CLmax L (flaps down) 3.3
AR 8
(W/S)TO (psf) 28
S (ft2
) 286
(W/P)TO (lbs/ft2
) 9
PTO 889
20. 20
E. Wing Configuration
After completion of the wing sizing, the next step in the design process was to decide upon a
suitable wing configuration to meet the mission constraints. This was no easy task, for the
majority of the initial concerns when choosing a wing configuration was bogged down with trade
studies. However, most of the information found in these studies was more explanatory than
numerical, and thus most of the positives and negatives of each configuration will be presented
accordingly.
The overall configuration was the first aspect to be investigated. There were 6 configurations
to choose from: conventional (tail aft of wing), flying wing (no horizontal tail or canard), tandem
wing, canard, three surface, and joined wing. There are advantages and disadvantages to each
design, but due to the rarity of, the limited available information on, and the improbable like of
use for a conventional aircraft of the tandem wing, canard, three surface and joined wing, only
the conventional wing and the flying wing will be discussed as possible choices. The
conventional wing has a variety of designs that can impact any variable from lift and drag, to
maneuverability, while the flying wing is more limited in its advantages and disadvantages. The
biggest advantage of the flying wing design is its ability to reduce drag over the entirety of the
aircraft while maintaining a high level of maneuverability. However, the biggest disadvantage
stems from this maneuverability. Because of the lack of vertical stabilizing control surfaces,
there is a decrease in lateral stability. When designing a STOL aircraft, which operates as low
velocities, stability was more important to the HS aircraft than maneuverability. The other issue
is that the flying wing design does not allow for much room for cargo and crew, which is a
majorly limiting factor for a mission where the main purpose is to deliver cargo and transport
passengers. This is the main reason the conventional design was chosen over the flying wing
design.
Figure II-2) Strutted High Wing Configuration
Step two of the wing configuration process was a bit more difficult due to the
comparisons that were necessary. The decision had to be made whether to structure the wing
with a braced (or strutted) support or simply construct the configuration as a cantilever wing.
According to Roskam3, the braced wing is more efficient for cruise velocities below 200kts,
21. 21
which is right in the range of our presumed cruise speed. Because of this, along with the fact
that braced wing structures tend to be lighter, stronger and cheaper than the cantilever wing, we
decided to choose the braced wing method. However, there are trade-offs to choosing this
configuration. Because of the struts holding up the wings, the aircraft experiences an increase in
drag.
The third and final wing configuration variable was vertical wing placement. The
available choices of high, mid, and low wing designs each had their own advantages and
disadvantages, though none more significant than the other. Roskam Part II3 explains the
advantages and disadvantages through a tabulated ranking system, 1 being the best and 3 being
the worst.
Table II-18) Roskam Wing Placement Comparison
High Wing Mid Wing Low Wing
Interference Drag 2 1 3
Lateral Stability 1 2 3
Visibility from Cabin 1 2 3
Landing Gear Weight 3 2 1
Total 7 7 10
Though the high wing design experiences more drag, it also increases the lift factor due to
the increase in wing area on the top of the fuselage. Landing gear weight is another non-factor
due to the need for a rugged, fixed system which could survive landing in rough climates.
Therefore, the landing gear was already predetermined as a weight liability regardless of wing
placement. Stability and visibility are of great necessity when flying in remote areas and thus are
two incredibly important factors to consider. Otherwise, the advantages and disadvantages of the
vertical wing placement change with varying speeds and altitudes. Because of all of these
factors, and the ease that came with applying the strutted wing configuration, the HS aircraft was
chosen to be a high wing design.
F. Airfoil Selection
Since the wing sizing had been completed as well as the wing configuration, all that was left
to arrive at an image of the wing was to apply an airfoil to the design. Following Roskam’s
Method, however, proved to be quite difficult when deciding on this design parameter. Because
this method did not give a specific way to choose an airfoil, outside sources were needed. Using
22. 22
the previously calculated CL Max from table II-16, along with Abbott and Doenhoff’s Theory of
Wing Sections1 book, the wing t/c was able to be found. Another result of using this book was
the ability to reduce the amount of applicable airfoils to four: the NACA 1410, NACA 2410,
NACA 23021, and NACA 4418.
Using the XFLR5 software, each airfoil was tested at two different conditions. The first test
condition was intended to simulate take-off altitude and velocity, while the second condition was
to emulate cruise condition. These velocities and altitudes for take-off and cruise were 72 knots
at 5000 feet and 190 knots at 15,000 feet, respectively. The variables M and Re were then
calculated as required inputs for the software using Appendix Table III-3 and Table III-4. The
four airfoils were then tested over an α range of -5 degrees to 20 degrees. Four output graphs
were given from the testing (Appendix Figures III-2 through III-5): Cl vs. α, Cl vs. Cd, Cm vs. α,
and Cl/Cd vs. α.
Each one of these graphs showed important aspects of each airfoil. One of the most
important conclusions from the four output graphs was the fact that the NACA 4418 airfoil
achieved the highest (Cl/Cd) MAX, and not only did that, but it achieved this value at the lowest α.
The NACA 4418 airfoil also had the highest Cl,0. The only deterring point of this airfoil that was
determined from the graphs was the Cm vs. α. When placed against the other three airfoils, the
NACA 4418 had the highest pitching moment about the wing. This would require an increase in
the St or the it. However, the positives of this airfoil far outweighed the negatives and thus, the
NACA 4418 was chosen as the HS aircraft’s airfoil.
23. 23
Figure II-3) NACA 4418 2D and 3D Airfoil
G. Fuselage & Interior Layout
The initial design criteria stated the HSA aircraft was to be able to seat 10 people (9
passengers, plus 1 pilot) with a total payload weight (passengers, cargo, and crew) of 1775 lbs.
These specifications lead to a common two seat cockpit and a passenger seating arrangement
featuring a single 16 in wide isle dividing four rows of passenger seats with a seat on each side
of the isle as shown in figure II-4 below. This layout allows for either 8 passengers with a pilot
and copilot or 9 passengers with a single pilot. Each of the 8 cabin passengers has a 17 in wide
seat with a 2 in wide outer armrest, 42 inches of head room, and 10 inches of forward leg room.
There is also 51 cubic feet of cargo room allocated in the aft of the aircraft. The 8 cabin seats are
all removable to increase the cargo room to a maximum of 280 cubic feet.
24. 24
Figure II-4) Empennage Configuration
Figure II-4) Interior Layout
The payload constraints of the aircraft ultimately created the final shape of the fuselage; a
complete layout with basic dimensions can be seen in figure III-6 in the appendix. The fuselage
is 33.1 ft in total length and features an elliptical shaped cabin in order to maximize passenger
ergonomics and cargo room. HSA has a total of four doors; one on each side of the cockpit for
the pilot and copilot/passenger, a large door on the left aft of the cabin for boarding passengers
and loading cargo, and finally a large rear cargo door that doubles as a ramp to ease the loading
of large heavy cargo.
H. Empennage Sizing and Configuration
Upon completion of the main wing, the empennage was the next portion of the aircraft to
be sized. Because Roskam’s Method became tedious and
rather confusing around this portion of the design
process, the Raymer book began to be heavily favored.
Though these two books employ different methods to
design an aircraft, both methods share many similarities
that output mirroring answers. Thus, it was decided that
the methods may actually work better in tandem with one
another, as to compare results.
Comparing both Roskam and Raymer pointed the
empennage design into a narrow field for tail
considerations. Though there are many configurations
that had to be considered, such as the T-tail and V-tail,
25. 25
Sv (ft2
) 26.3
Chord0 (ft) 4.3
Chordt (ft) 2.4
Span (ft) 6.5
Λ.25 0.35
the simplicity and weight saving properties of the conventional tail led the design down that path.
Simply put, the conventional tail would provide adequate stability and control at a light weight,
while saving the team time and resources (due to the myriad of previous aircraft employing the
design for comparison) researching methods to optimize the design. Also, the purpose of the
aircraft did not require superior control or maneuverability which immediately negated any
advantages that some of the tail configurations presented. Also, the conventional tail was stated
to be among the cheapest configurations, which was important for a smaller aircraft. A dorsal fin
was also added to the configuration to increase the stall angle and help prevent rudder lock.
Using initial sizing chapter of Raymer and the established fuselage length from the
interior layout, the moment arm between the main wing aerodynamic center and the horizontal
tail aerodynamic center were then sized and estimated to be 60% of the fuselage length, or
roughly 21ft. Though this number would later change slightly, this was a good start to find
where the horizontal and vertical tail would lie aft of the main wing. The next step in the process
would simply be finding the corresponding areas of the tail surfaces. Continuing on with
Raymer’s method, the tail volume coefficients were estimated using table 6.4 in the book.
Plugging these two values into corresponding equations III.D.I – III.D.4 of Raymer (found in the
appendix), the horizontal and vertical tail wetted areas were found to be 65.33 ft2 and 26.11 ft2,
respectively. The rest of the dimensions of the vertical and horizontal tail are shown in the tables
below.
After establishing the taper and sweep of each of the tail surfaces from suggestions in
Raymer, tables with such values can be found in the appendix, the airfoils of each of the tail
sections had to be determined. With the assistance of professors as well as comparison data, the
airfoils were narrowed down to a handful of symmetric NACA foils. After further research, the
two most prominent choices among STOL aircraft emerged: the NACA 0006 and the NACA
0009. Applying each airfoil to the 3-D model in the Pro-Engineer software, it became apparent
that the NACA 0006 seemed much too thin to support the sheer size of the horizontal tail.
Therefore to keep the empennage standardized, the NACA 0009 was chosen as the airfoil for
both surfaces.
St (ft2
) 50.1
Chord (ft) 2.9
Span (ft) 17.3
Table II-19) Dimensions of the Vertical and Horizontal Stabilizers
26. 26
I. Propeller Sizing
Continuing with the sizing trend established through the last few design sections, the
propeller sizing was the next aspect of the aircraft that had to be determined to help during the
calculation of the drag polar in the proceeding steps. Instead of continuing with the Raymer
equations, a comparison between both Raymer and Roskam was used in this process. Both
methods laid out ways of finding both the number of propellers needed as well as the blade
diameter, however, Roskam describes equations to find the blade loading after a diameter is
chosen. Here are the results of both methods in table form as well as extra calculations can be
found in the appendices.
Due to noise issues as well as flow issues around the propeller tip, the tip velocity was the
main constraint when deciding on the number of propellers to use as well as their size. Because
the tip velocity should remain less than the critical mach number of the propeller airfoil, the rule
of thumb established in the Raymer book states that the tip speed should be less than 850 ft/s.
Also, noise is usually a factor when determining propeller diameter, therefore the tip speed was
limited even farther to 750 ft/s. Considering this tip speed, the equations were implemented for
two, three and four propellers. The output provided only one reasonable choice for the number
of propellers, four. Using the diameter that was found for four propellers, the blade loading was
then determined through equations in the Roskam book. As one can see from the tables above,
the blade loading ended around 4.6 lbs/ft2.
np 4
Pbl (lbs/ft2)
Dp (ft)
4.0 9.14
4.1 9.03
4.2 8.92
4.3 8.82
4.4 8.72
4.5 8.62
4.6 8.52
4.7 8.43
4.8 8.34
4.9 8.26
5.0 8.18
np 4
Kp 1.5
Dp (ft) 8.54
(Vtip)static 711
(Vtip)helical 722
Table II-20) Raymer (left) and Roskam (right) Properties of the Propeller
27. 27
J. Propulsion
Using the matching results, shown in table II-20 below, the power required for take-off
was found to be 889 HP. A propeller efficiency of 85% was then taken into account resulting in
an actual shaft horsepower requirement of 1,000 for take-off. The high power required along
with a cruising altitude of 25,000 led to the decision of using a turboprop engine. Turboprop
engines offer a greater power to weight ratio, increased reliability, and greater efficiency at
higher altitudes than piston-prop engines. Twin-engines were also ruled out early in the design
phase due to the massive increase in weight from 4225 to 9115 lbs empty.
Table II-20) Propulsion Parameters
Matching Results
WTO (lbs) 8000
CLmax (flaps up) 1.9
CLmax TO 2.1
CLmax L 3.3
A 8
(W/S)TO (psf) 28
S (ft2
) 286
(W/P)TO (lbs/ft2
) 9
PTO 889
In order to select the best engine for our aircraft 67 turboprop engines from five different
manufactures were compared. A comparison table can be seen by looking at table 1.2 located in
appendix A. The best engine for our aircraft was chosen based on shaft horsepower, specific fuel
consumption, OPR, weight, physical dimensions, cost, and maintenance. Honeywell’s TPE 331-
12 was the engine of choice for the aircraft. It offers the required 1,000 SHP with a SFC of only
0.553 and a weighing in at a mere 415 lbs. Honeywell’s turboprop engine offers the best power
to weight ratio and lowest specific fuel consumption of any engine in its class. The initial cost of
the engine was also very competitive. The only down side to this high performance engine is the
maintenance cost. The cost to maintain and inspect the TPE 331 is greater than the Pratt Whitney
PT6 due to the increased complexity of its internal components. However, this increased cost
was deemed an acceptable trade-off for the increase in performance as the ability to take-off
from very short airfields is the main purpose of the aircraft. The engine parameters for the TPE
331-12 can be seen below in table II-21.
Table II-21) Engine Parameters
28. 28
Engine Parameters
Weight (lbs) 405
Length (ft) 2.27
Width (ft) 1.81
Height (ft) 1.58
rps 26.52
Power (hp) 1000
ESHP (hp) 1050
SFC 0.553
ESFC 0.523
The initial estimated engine parameters from Raymer resulted in table II-22 shown below.
Engine Parameters
Weight (lbs) 445
Length (ft) 4.69
Diameter (ft) 1.84
Table II-22) Final Engine Parameters
K. Weight Distribution & C.G. Location
Having completely sized almost every aspect of the aircraft, the next step in the design
process was to calculate and tabulate the weight of each individual component and its location.
Continuing with the methods laid out by Raymer, each weight was found using a series of
equations (shown in the appendix) that took into account the size and shape of the components
along with their configurations. Here is an example of those equations, this one estimating the
weight of the main wing.
𝑊 𝑤𝑖𝑛𝑔 = 0.0103𝐾𝑑𝑤 𝐾𝑣𝑠( 𝑊𝑑𝑔 𝑁𝑧)
0.5
𝑆 𝑤
0.622 𝐴0.785 (
𝑡
𝑐
)
𝑟𝑜𝑜𝑡
−0.4
∗ (1 + 𝜆)0.05(𝑐𝑜𝑠Λ)−0.1 𝑆𝑐𝑠𝑤
0.04
Finding the location of each component was much more difficult considering the effect
on the center of gravity (C.G.) and because of this, the effect on the static stability of the aircraft.
Though a few basic approximations were given for initial placement, most of the components
had to either pass the eye test, which simply says to put the parts where they look correct, or had
to be tinkered with using other aircrafts’ weight charts for comparison. However, placing the
29. 29
wing was the toughest challenge of any of the components due to the fact that it is the most
important way to change the static margin by changing the moment arm between the wing and
C.G. location. Here is the complete component breakdown of weights and locations when the
aircraft is at max cargo at the farthest aft C.G. position.
Component Weight (lbs) Location (ft) Moment (in-lb)
Wing 737 11.4 8395
Horizontal Tail 96 30.9 2971
Vertical Tail 83 31.9 2643
Fuselage 1048 16.5 17333
Main Landing Ger 408 14.0 5716
Nose Landing Gear 116 3.1 360
Installed Engine 653 3.1 2026
Fuel System 123 5.6 688
Flight Controls 153 16.5 2530
Hydraulics 0.20 16.5 3
Electrical 316 16.5 5224
Avionics 433 5.3 2297
Pilot/Crew 350 8.6 3004
Cargo/Passengers 1600 19.2 30667
Fuel 1986 9 17870
Furnishings 64 8.6 549
Totals 8165 12.5 102278
As one can see from the table, the final wing location chosen for the design ended up
right around 11.4 ft. aft of the nose of the aircraft. This location provided acceptable stability
(shown later in Section K), while maintaining visibility from the pilot’s cabin and clearance for
the proposed location of the driver door.
Table II-20) Weight, Location and Moment of each Component
30. 30
L. Structures
The structure for HSA, like most aircraft, needs to be lightweight, strong, and affordable.
Constructing the any major components from composite was ruled out early in the structural
design phase. Composites while offering high strength with low weight are often expensive,
difficult to work with, and harder to repair than conventional metal alloys. The obvious metal
alloy of choice for the majority of the aircrafts structure was 6061 aluminium; it offers a high
strength to weight ratio, is relatively inexpensive, readily available, easy to work with, and easy
to repair with conventional tools. Ease of repair is important given the remote operating locations
of the aircraft; if the airplanes’ structure was to become damaged landing in a remote area it
needs to be able to be repaired using conventional tools.
Figure II-5) Wing Loading
With 6061 aluminum chosen as the metal of choice for the structure of the aircraft a
bending moment analysis was performed on the wing of the aircraft. The wing having a half span
of 286 in can be seen in figure II-5 above. The resulting bending moments when the wing is
subjected to 14,000 lb/in are shown in table III-8 and figure III-9. A shear force diagram, along
with the slope and deflection of the wing can be seen in figures III-8, figure III-10, and figure II-
11 respectively. Upon examination of the bending moment data it was determined that the wing
spars needed to be placed at approximately 44% of the half span.
M. Performance
While the component weights were being finalized, the drag polar for the entire aircraft
was being compiled. Initially, the CD0 was approximated through a simple equation in Raymer:
𝐶 𝐷0
= 𝐶 𝑓𝑒
𝑆 𝑤𝑒𝑡
𝑆 𝑟𝑒𝑓
31. 31
However, this was only used as a comparison when compiling the entire component drag
breakdown. This initial CD0 was found to be .0227 at cruise conditions. Because the various
surfaces of the aircraft used different equations to find each component drag contribution, each
surface had to be analyzed separately and then compiled later. The different components broke
down as follows: wing, fuselage, horizontal tail, vertical tail, strut 1, strut 2, landing gear back
tires, and landing gear front tire.
Using various equations that can be found in the appendix, as well as the Reynold’s
number, wetted area, and coefficient of friction, the entire aircraft CD0 for cruise, take-off and
landing were found to be .024426, .05442, .08426, respectively. As one can see, the CD0 of the
take-off and landing phases are much higher than that of cruise due to the addition of drag
caused by the deflection of the flaps and an increase of lift. At take-off the flaps are only
deflected about 60% while during landing the flaps are fully deflected.
After finding these CD0 for each stage of flight, the entire drag build-up for the aircraft
was found using a basic aerodynamics equation adding together the parasite and induced drag:
𝐶 𝐷 = 𝐶 𝐷0 +
𝐶 𝐿
2
𝜋𝐴𝑅𝑒
Using a range of CL values from -1 to 5, the CD of the entire aircraft was plotted as seen below.
Figure II-6) Drag Polars at Cruise, Take-Off and Landing
-2
-1
0
1
2
3
4
5
6
0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6
CL
CD
Drag Polar
Cruise
Take-Off
Landing
32. 32
After finding the drag polar for the entire aircraft, the take-off distance had to be
calculated. Because the aircraft was considered a STOL aircraft, take-off was a major factor in
considering the aircraft design a success. However, there was more than just an estabilished
take-off distance to be considered when calculating whether or not the aircraft could clear the
obstacle at the end of the runway. The roll distance, rotation distance and transition distance all
needed to be considered before the obstacle could be considered cleared. Therefore, using
equations in the performance section of the appendix, the total runway distances were found and
placed in the table below. The total take-off distance when clearing a 50 foot obstacle at the end
of the runway was found to be 1146 ft. This distance clearly surpasses the initial design
requirement of take-off within 1500 feet.
Range (mi) 830.1
SG (ft) 531.6
SR (ft) 121.5
Stransition (ft) 493.1
Total (ft) 1146.2
Table II-21) Range, Roll, Rotation and Transition Distances
O. Stability and Control
Having all of the components sized, weighed, and locations finalized, the next step in the
process was to determine the Static and Dynamic stability of the aircraft. Using the most
forward and aft C.G. locations, the C.G. travel was easily determined. As one can see from the
figure below, the C.G. travel was calculated to be 13.4 inches. At each of these forward and aft
C.G. locations, the static margin was also calculated. At the forward most C.G. location (empty
weight with 2 pilots) the static margin was found to be 33%. Now this may seem very stable for
an aircraft, but due to the fact that the aircraft being designed was primarily being used for cargo
transport, this static margin was acceptable as the maximum. However, the most important static
margin percentage was the aft C.G. location when the aircraft would be flying at maximum
cargo. This percentage was found to be 13%. Due to the fact that the aircraft would be flying
most of the time at max cargo, or somewhere close to it, a static margin of 13% was almost ideal
for maintaining a level and yet, maneuverable flight experience.
33. 33
Figure II-7) C.G. Locations with respect to the entire aircraft
From here, the phugoid and short period mode were to be found. Using a matlab
program to find the stability derivatives, the parameters of our aircraft were plugged in
P. Risk Analysis
A general risk analysis, plotting possible failures verses the severity of their consequence
on a scale of 1 to 5, was performed on the aircraft. A plot of the risk analysis is shown below in
figure 1.1. Engine failure was the first scenario to be considered due to its inherent presence on
all aircraft. This scenario ranked a 5 on severity due to the remote locations the aircraft will be
operating in. If the aircrafts engine were to fail it could take rescue teams days to find and rescue
the occupants if the pilot was to be forced to perform an emergency landing in a remote area.
However, the likelihood of a engine failure occurring ranked a mere 2 as turboprop engines are
incredibly reliable especially when compared to their piston-prop counterparts.
34. 34
Figure 1.1
Landing gear failure was the next scenario to be examined. The HSA aircraft is designed
to land in fields that may be of only fair condition: rough, unmaintained, and likely consisting of
grass, dirt, and gravel. The poor and unknown condition of the remote airfields is the reason why
landing gear failure ranked a 3 on the likelihood of occurrence. This scenario also scored a 4 for
severity of consequence as the airfields poor condition could cause a large amount of damage to
the aircraft if the landing gear were to fail. The remote location of the airfield also increased the
consequence as it would likely take an extended period of time before parts and tools to fix the
aircraft could be acquired.
Hydraulic system failure was another scenario that was considered. The likelihood of
having hydraulic system failure is rather low, only a 2. The severity of this occurring ranked a 3.
If some sort of hydraulic failure was to occur the pilot would have several different options at his
disposal to counter the issue. It would be most likely that a single seal would develop a leak in
which case the pilot would still maintain partial control of the particular control surface
associated with the faulty hydraulic. If a control surface’s hydraulics were to fail completely the
pilot would still be able to compensate for the stuck control surface by actuating the remaining
controls to oppose the force exerted by the rogue surface. The only case where the pilot would be
unable to compensate would be during a complete hydraulic system failure; in such a case the
35. 35
pilot would lose total control of all control surfaces. The likelihood of all the hydraulic systems
totally failing simultaneously is so remote that it would rank less than a 1 on a scale of 1 to 5.
The possibility of underestimating the drag when calculating the performance of the
aircraft was another scenario we considered. If the drag was in face underestimated the
performance of the aircraft would suffer; range would decrease, fuel consumption would
increase, cruise speed would decrease, and take-off distance would increase slightly. The
severity of this happening is quite low at 2. This is due to the HAS aircraft not being designed to
travel at high speeds or travel great distances. The requirements of the aircraft were to travel 800
miles without refueling and take-off in a distance of 1500 ft, while clearing a 50 ft obstacle at an
elevation of 5,000 ft. This criterion was met with room to spare. The Breguet range equation
gave a max range of 830 miles along with a take-off distance of 1146 ft using Roskam’s
equations. The only real repercussion of underestimating the drag would be an increased take-off
distance; with calculations showing 354 ft to spare the consequences are minimal. However, the
likelihood of this occurring was given to be 3. The reason for the fair likelihood is because the
drag data for the aircraft was obtained using XFLR5. There was insufficient time and manpower
to build an entire 3D model and verify the drag data with wind tunnel testing.
Finally, the last scenario considered was that the stability margins were underestimated.
This would pose a fairly serious consequence of 3. If the stability margin turned out to be much
lower than calculated the aircraft could become difficult to maneuver and overly sensitive to the
pilots inputs. On the other hand, if the stability margins were in fact much greater than what was
calculated the airplane would be overly stable and lack maneuverability. The HSA aircraft is
designed to be used as a transport for people and supplies. Therefore, the stability margins were
designed conservatively with a static margin range of 13-33%. It is highly unlikely that the
aircrafts stability margins would turn out to be overly low and the aircraft become unstable; the
likelihood of this scenario occurring was determined to be 1.
Q. Cost Estimation
In order to estimate the cost of the HSA aircraft equations 18.1-18.9 from Raymer were
used. These calculations give a cost breakdown for the aircraft based on the following
parameters: aircraft empty weight, maximum velocity, number of aircraft to be produced in five
years, number of flight tests, number of engines, maximum engine thrust, maximum Mach
number, turbine inlet temperature, and complexity of avionics. A table containing the cost break
down can be seen below in table 1.6; the cost estimation was based on a total quantity of 500
aircraft being produced in 5 years.
Table 1.6
36. 36
EngineeringHours $1,002,020.65
ToolingHours $805,304.80
ManufacturingLabor Hours $4,833,919.17
QualityControl Hours $642,911.25
DevelopmentSupport $17,219,313.58
FlightTesting $6,786,496.91
ManufacturingMaterials $185,823,508.28
EngineeringProduction $1,213,942.05
RDT&E + flyaway $1,587,113,179.69
Total CostPerAircraft $3,174,226.36
R. Competitive Comparison
The HSA is a STOL aircraft that is designed specifically to take-off and land on rough
unpaved fields in order to deliver its payload to villages in remote locations. There are currently
two aircraft in use for this mission profile the Quest Kodiak, and Cessna 208. The Kodiak like
the HSA was designed specifically for these sort of humanitarian missions; Cessna’s 208 while
not designed specifically for this mission has become commonly used due to them being reliable
and readily available. HSA has been designed to take the customer into even more remote of
areas then the Kodiak or C208 are able to. A comparison of the three aircraft can be seen below
in table 1.8.
Take-off distance is an important factor, the shorter the take-off distance, the tighter and
more secluded of an area the aircraft can land in. The C208 manages a take-off distance of 2,420
ft and the Kodiak 1,181 ft, both planes at sea level and max take-off weight. The HSA takes off
in a mere 1,146 ft at max take-off weight and while at an elevation of 5,000 ft. The payload
capacity of the HSA is also 2.5 times that of the Kodiak and nearly as much as the C208 which is
a much larger aircraft incapable of taking off from sub 1500 ft airfields with payload. Finally, the
HSA is also very competitively priced at $250k less than the Cessna 208 and a mere $70k more
than the Kodiak.
Table 1.8
HSA QuestKodiak Cessna208
PayloadCapicityw/max fuel (lbs) 1775 733 2324
Take-off Distance atmax weight(ft) 1146 1181 2420
Pricing(USD) $1,770,000 $1,700,000 $2,020,000
37. 37
III. Appendix
A. Fuel Sensitivities
The fuel sensitivities were calculated using data from the fuel fractions and weight sizing portion,
such as the Mission Fuel Fraction, the A and B empty weight parameters,the cruise and endurance ratios
38. 38
(Table II-2 and Table II-3), and the take off-weight. First, two more mission parameters were calculated,
the Mission Reserve FuelFraction and the Mission Unusable Fuel Fraction with the following equations:
𝑀𝑟𝑒𝑠 =
𝑊𝑓𝑟𝑒𝑠𝑒𝑟𝑣𝑒
𝑊𝑓
[𝐼𝐼𝐼. 𝐴.1]
𝑀𝑡𝑓𝑜 =
𝑊𝑡𝑓𝑜
𝑊𝑓
[𝐼𝐼𝐼. 𝐴.2]
Two more empty weight parameters,C and D were calculated using the following equations. It
should be noted that the payload weight in D is different for the differing legs of the mission, seeing as
half of the payload is not present during the second leg or the return trip.
𝐶 = 𝑀𝑓𝑓 ∗ (1 + 𝑀𝑟𝑒𝑠)− 𝑀𝑡𝑓𝑜 − 𝑀𝑟𝑒𝑠 [𝐼𝐼𝐼. 𝐴.3]
𝐷 = 𝑊𝑝𝑎𝑦𝑙𝑜𝑎𝑑 + 𝑊𝑐𝑟𝑒𝑤 [𝐼𝐼𝐼. 𝐴.4]
After this, three more parameters were calculated. They are as follows:
𝑅̅ = 𝑅 𝑐𝑟𝑢𝑖𝑠𝑒 𝑐 𝑝,𝑐𝑟𝑢𝑖𝑠𝑒 (
1
375𝜂 𝑝,𝑐𝑟𝑢𝑖𝑠𝑒
𝐿
𝐷 𝑐𝑟𝑢𝑖𝑠𝑒
) [𝐼𝐼𝐼. 𝐴.5]
𝐸̅ = 𝐸𝑙𝑡𝑟 𝑉𝑙𝑡𝑟 𝑐 𝑝,𝑙𝑡𝑟 (
1
375𝜂 𝑝,𝑙𝑡𝑟
𝐿
𝐷 𝑙𝑡𝑟
) [𝐼𝐼𝐼. 𝐴.6]
𝐹 =
−𝐵𝑊𝑇𝑂
2
𝑀𝑓𝑓(1+ 𝑀𝑟𝑒𝑠)
𝐶 𝑊𝑇𝑂(1− 𝐵) − 𝐷
[𝐼𝐼𝐼. 𝐴.7]
Values from all of these calculations for the single engine, to site case are shown below in Table III-1.
Table III-1) Sensitivity Parameters: Single Engine, To Site
Mff 0.872 A -0.144
Mres 0.100 B 1.1162
Mtfo 0.000 C 0.8589
D (lbs) 1775.30
39. 39
WTO (lbs) 8000
Rbar 0.089
Ebar 0.016
F (lbs) 26614.98
These parameters were then used in conjunction with the following equations in order to find the
appropriate fuel sensitivities. The fuel sensitivities for the single engine, to site case (before corrected to
normal change values as shown in part II.C of the report) are shown in Table III-3.
𝛿𝑊𝑇𝑂
𝛿𝑅
=
𝐹𝑐 𝑝,𝑐𝑟𝑢𝑖 𝑠 𝑒
375𝜂 𝑝,𝑐𝑟𝑢𝑖𝑠𝑒
𝐿
𝐷 𝑐𝑟𝑢𝑖𝑠𝑒
[𝐼𝐼𝐼. 𝐴.8]
𝛿𝑊𝑇𝑂
𝛿𝐸
=
𝐹𝐸𝑙𝑡𝑟 𝑐 𝑝,𝑙𝑡𝑟
375𝜂 𝑝,𝑙𝑡𝑟
𝐿
𝐷 𝑙𝑡𝑟
[𝐼𝐼𝐼. 𝐴.9]
𝛿𝑊𝑇𝑂
𝛿𝑐 𝑝
=
𝐹 ∗ 𝑅 𝑐𝑟𝑢𝑖𝑠𝑒
375𝜂 𝑝,𝑐𝑟𝑢𝑖 𝑠 𝑒
𝐿
𝐷 𝑐𝑟𝑢𝑖𝑠𝑒
[𝐼𝐼𝐼. 𝐴.10]
𝛿𝑊𝑇𝑂
𝛿𝜂 𝑝
=
−𝐹 ∗ 𝑅 ∗ 𝑐 𝑝,𝑐𝑟𝑢𝑖 𝑠 𝑒
375( 𝜂 𝑝,𝑐𝑟𝑢𝑖𝑠𝑒)
2 𝐿
𝐷 𝑐𝑟𝑢𝑖𝑠𝑒
[𝐼𝐼𝐼. 𝐴.11]
𝛿𝑊𝑇𝑂
𝛿 (
𝐿
𝐷
)
=
−𝐹 ∗ 𝑅 ∗ 𝑐 𝑝,𝑐𝑟𝑢𝑖𝑠𝑒
375𝜂 𝑝,𝑐𝑟𝑢𝑖𝑠𝑒 (
𝐿
𝐷 𝑐𝑟𝑢𝑖𝑠𝑒
)
2 [𝐼𝐼𝐼. 𝐴.12]
Table III-3) Fuel Sensitivities: Single Engine, To Site
δWTO/δR 5.91 (lbs/NM)
δWTO/δE 868 (lbs/hr)
δWTO/δcp 3940 (lbs/lbs/hp/hr)
δWTO/δηp -2960 (lbs)
δWTO/δ(L/D) -263 (lbs)
B. Wing and Power Sizing
𝑆 𝑇𝑂 = 8.134 ∙ 𝑇𝑂𝑃23 + 0.0149 ∙ 𝑇𝑂𝑃23
2
[𝐼𝐼𝐼. 𝐵.1]
60. 60
I. References:
1) Abbott, I. and Doenhoff, A. Theory of Wing Sections.Mineola, NY: Dover Publications, Inc,
1959. Print.
2) Roskam, J. Airplane Design Part I: Preliminary Sizing of Airplanes. Ottawa,KA:Roskam
Aviation and Engineering Corporation, 1989. Print.
3) Roskam, J. Airplane Design Part II: Preliminary Configuration,Design and Integration of the
Propulsion System. Ottawa,KA:Roskam Aviation and Engineering Corporation, 1989. Print.
4) Roskam, J. Airplane Design Part III: Layout Design of Cockpit, Fuselage, Wing and Empennage:
Cutaways and Inboard Profiles.Ottawa,KA:Roskam Aviation and Engineering Corporation,
1989. Print.
5) Roskam, J. Airplane Design Part IV: Layout of Landing Gear and Systems. Ottawa,KA:Roskam
Aviation and Engineering Corporation, 1989. Print.
6) Raymer, D. Aircraft Design: A Conceptual Approach, Playa delRey, CA: Conceptual Research
Corporation, 2012. Print
