NASA_Medevac_ Report-2016_Maldonado

NASA Glenn Research Center 1
Conceptual Design of a Medical Evacuation Air Vehicle with Distributed Turboelectric
Propulsion
Victor Maldonado1
University of Texas at San Antonio, San Antonio, TX, 78249
Faculty Fellowship Program Summer 2016
Don Simon2
; Sanjay Garg3
NASA Glenn Research Center, Cleveland, OH 44135
Abstract
The need to rapidly transport injured victims to a trauma center from remote emergency locations is critical
to a society’s medical response and infrastructure. In this preliminary study, a conceptual design for a
medical evacuation (Medevac) air vehicle was completed in order to explore more fundamental questions
related to what propulsion technology is most suitable for these types of missions. The specific mission
considered in this study assumes an autonomous vehicle with one onboard paramedic required to perform
basic navigation tasks (e.g. landing), load the victim into the vehicle, and administer first aid en-route to the
trauma center. Based on the distribution of trauma centers around the country and the “golden hour” for
medical treatment, the range and cruise speed for a Medevac air vehicle concept was determined at 805 km
and 217 knots, respectively. The cruise speed requirement indicated a departure from traditional rotorcraft
concepts, and the need to land in confined spaces made it clear that VTOL (Vertical Takeoff and Landing)
capability would be necessary. A conceptual design approach was implemented to estimate the gross takeoff
weight (3,250 lb) and configure the aircraft. Hybrid turboelectric propulsion was determined as the most
viable propulsion concept due to its distributed power nature and ability to more easily vector thrust for
VTOL and forward flight. Specifically, a propulsion system that consists of a turboshaft engine, electric
generator, and 4 AC motors with ducted fans is considered. Initial blade element theory calculations suggest
that such an aircraft would require approximately 1 MW of power for VTOL, which is feasible with
emerging motor technology. An analytical turboelectric engine model is also constructed which models the
turboshaft engine, generator, and electric motor to guide more advanced analysis and turboelectric
technology going forward.
I. Introduction
HIS paper addresses a challenge question posed by NASA (Convergent Aeronautics Solutions) regarding
suitable propulsion technologies for medical evacuation air vehicles. Medical evacuation (medevac) is the
practice of movement and en-route care provided by medical personnel to injured patients being evacuated from
the scene of an accident to a trauma center using medically equipped ground vehicles or aircraft. In order to
successfully study this problem, a medevac mission was defined by considering the following scenario; a person is
seriously injured somewhere in the U.S. and requires medical treatment and transportation to the nearest trauma
center. It was determined based on studying the distribution of trauma centers on a map of the country, that there
exists a maximum distance of approximately 250 miles from a given point to the nearest trauma center. Such an
example is illustrated in the regional map of California shown on Figure 1. A 250-mile radius circle is drawn
around a level-1 trauma center in Fresno, showing the service land coverage of a hypothetical aircraft deployed
from that trauma center with a range of 250 miles. The mission was defined in terms of a round-trip (from the
trauma center to a 250 mile distance location and return to the facility) hence a total range of 500 miles (plus 5%)
was specified as the range requirement.
1
Assistant Professor, Department of Mechanical Engineering, University of Texas at San Antonio.
2
AST Control Systems, Intelligent Control and Autonomy Branch, MS 77-1, 21000 Brookpark Road.
3
Chief, Intelligent Control and Autonomy Branch, MS 77-1, 21000 Brookpark Road.
T

The cruise speed was established based on the 10-1-2 rule formed by NATO in which injured victims should
ideally receive first-aid medical attention within 1-hour of injury, and arrive at a trauma center within 2 hours. This
guideline is based upon serious injuries where the patient will likely require surgery. In the unlikely scenario where
the patient is located at a maximum 250 miles away from the nearest trauma center, then the aircraft would require
a cruise speed of 250 miles/hr to arrive at the emergency location in one hour (neglecting the time required for
dispatch of the aircraft). Moreover, the mission considers an onboard paramedic (EMT) in order to load the patient
onto the aircraft upon arrival, and administer first-aid en-route to the medical facility.
Figure 1. Representative map with trauma centers to determine MedEvac range
Three basic aircraft configurations or concepts are initially envisioned to carry out MedEvac missions. These
aircraft concepts are: (1) a conventional helicopter (one main/ tail rotor), (2) a fan-in-vehicle “flying car” type
concept, and (3) a hybrid helicopter (or multi-copter) and fixed-wing aircraft. The conventional helicopter is the
most common type of VTOL aircraft, and offers the highest flight efficiency and lowest power loading of any
rotorcraft configuration. Manned conventional helicopters are currently in use for medical evacuation, however due
to the constrained forward flight speed (at an average of about 241 kph), which is significantly below the specified
434 kph cruise speed MedEvac mission requirement, it is not a suitable concept for this study. Moreover, helicopter
large rotor strikes become a concern when landing in confined spaces or brush such as what may be encountered in
the wilderness or natural disaster scenarios. The second concept of a “flying car” configuration is typically
characterized as having a car or truck type airframe with or without ground driving capability. It also has a
combination of ducted fans and/or open rotors for propulsion. This concept has been proposed by the Department
of Defense, e.g. the DARPA road capable Transfomer (TX) UAS, which has a flight speed of about 241 kph. In
addition, the first operating UAV specifically designed for MedEvac applications, the Air Mule manufactured by
Urban Aeronautics of Israel can be classified as a “flying car” and has a speed of 185 kph. The major drawback of
this type of configuration is the low-speed flight due in part to the large amount of parasite drag of the airframe,
decreasing the average speed below that of conventional helicopters. The last concept configuration is the hybrid
helicopter fixed-wing aircraft that combines the VTOL capability of helicopters with the high-speed forward flight
efficiency of fixed-wing aircraft. Historically, hybrid aircraft such as the tilt-rotor V-22 Osprey sacrifice the
performance of both flight modes (compared to dedicated single flight mode helicopters or fixed-wing aircraft) in
order to perform both mode functions to a lesser performance degree. This is to be expected, however when the
flight requirements are fixed for high-speed forward flight (at the limit or beyond the capability for conventional
rotorcraft) with significant range and VTOL capability, a hybrid design such as this is practically the only option.
The disadvantages of hybrid aircraft, however, are that they generally require more installed power (during takeoff
and landing) and are more mechanically complex designs due to the flight transition mechanism. Innovation in the

propulsion systems and control method can minimize the adverse impact of these factors. After an evaluation of the
3 aircraft concepts, the selected concept for the medevac investigation is the hybrid rotorcraft fixed-wing aircraft.
II. Aircraft Conceptual Design
A preliminary takeoff weight calculation was then carried out for this concept based on fixed-wing formulations to
estimate the empty weight, fuel weight, and payload as follows,
𝑊!" = 𝑊!"#$% + 𝑊!"#$ + 𝑊!"#$%"& (1)
The empty weight makes up a certain percentage of the take-off weight which in general depends on the type of
aircraft. The ratio between the empty weight to the take-off weight is the structure coefficient and generally
decreases with increasing take-off weight. The structure coefficient is given as,
𝑠 =
!!"#$%
!!"
(2)
A structure coefficient of s = 0.6 was utilized based on historical figures for a general aviation aircraft. The total
fuel weight used during the mission is calculated by estimating and summing the fuel weight for each individual
flight phase individually. We must first establish a mission profile or flight plan for the MedEvac UAV, which
consists of the following phases:
1. Engine start-up and take-off from trauma center
2. Climb and acceleration to cruise conditions
3. Cruise out to emergency location
4. Loiter
5. Landing, load victim (10 minutes)
6. Engine start-up and take-off
7. Climb and acceleration to cruise conditions
8. Cruise out to trauma center
9. Landing, unload victim
These flight phases will be described further in the next section. The weight used in each of these phases is
governed by a combination of empirical fuel weight fractions and range equations which is related to the total fuel
weight 𝑊!"#$!!"" − 𝑊!"!"#!$ according to the following equation,
𝑊!"#$%#& 𝑊!"#$!!"" !"#$
=
!!",!"#
!!"
!!"#$%
!!",!"#
!!"#$%&
!!"#$%
!!"#$%&
!!"#$%&
!!"#$%#&
!!"#!"#
(3)
Additional fuel weight fractions were utilized for phases 6 to 9 for the return trip to the trauma center. A schematic
of the mission profile is shown on Figure 2.
Figure 2. Mission profile for medevac

A. Takeoff Weight Analysis
We will begin with officially stating some important design requirements and target performance specifications:
1. Range: 500 miles +5% 456 nm (845 km)
2. Cruise Speed 235 ktas (434 km/h, M = 0.35)
3. Cruise altitude 8,000 ft (2,438 m)
4. Cabin & Crew Unmanned, 1 paramedic + 1 victim
Next, we will state the parameters that are used as input to the take-off weight analysis. The values are based on
established flight requirements and conceived from a hybrid VTOL-fixed wing aircraft:
1. Cruise Mach number 0.35
2. Cruise altitude, ft 8,000
3. Range (nm) – one-way 228
4. Engine TSFC (cruise) 0.5
5. Loiter time (min) 15
6. Loiter altitude (ft) 1,000
7. Fuel reserve (%) 5
8. Trapped fuel (%) 1
9. Structure factor 0.6
10. Payload (lb) 700
11. Aspect ratio 5
A description of the parameter and an explanation for the selected values is given. The cruise Mach number is
calculated as 0.35 from a conversion of the required cruise speed of 434 km/h. The cruise altitude is set as 2,439 m
which is consistent with general aviation aircraft. A range of 228 nm one-way trip is required based on a map
analysis of the location of trauma centers (level 1 or 2) in the US. The thrust specific fuel consumption (TSFC) was
chosen as 0.5 lb-fuel/hr/lbf, which is an estimate based on a modern turbo-prop engine. Potential engines for the
medevac vehicle will be analyzed in more detail later, a hybrid turbofan/ ducted fan concept may seem attractive
given the multicopter fixed-wing concept. The loiter time is specified as 15 minutes, which is dedicated to finding a
proper landing spot. The fuel reserve and trapped fuel percentages are standard figures found in any aircraft design
textbook, values of 5% fuel reserve and 1% trapped fuel (in the fuel tank system) were used. The structure factor
was calculated according to the following simplified formula, which at this point of the design process can only use
the take-off weight as an input.
𝑠 = 𝐴𝑊!"
!
(4)
For a jet transport, the constants A and C have values of 1.02 and -0.06 respectively. If we substitute a value for the
take-off weight of 3,000 lbs (a crude initial estimate) we solve for a structure factor of s = 0.63. We will utilize a
value of 0.6 for the calculation of takeoff weight. A payload weight estimate of 700 lbs is used, which factors the
weight of two adults and the medical equipment that is contained onboard the aircraft. A more refined payload
estimate based on MedEvac treatment during evacuation may be performed. Finally, a wing aspect ratio of 4 is
utilized to calculate the cruise flight L/D of 14 which is a historically realistic value for small manned aircraft.
The take-off weight formulation developed by Dr. Thomas Corke and others was applied, with minor
modifications given the purposes of this aircraft design. For the first take-off weight iteration, we assume a take-off
weight estimate which is 5,000 lb. We calculate the weight of the aircraft following the start-up and take-off phase
by employing the take-off fuel weight fraction as follows,
!!",!"#
!!"
= 0.975 (5)
Using this fuel weight fraction, the final weight after the take-off phase is 4,875 lbs, where 125 lbs of fuel was used
for the takeoff. The next flight phase is climb and acceleration to cruise conditions. The fuel weight fraction is
calculated according to,

!!"#$%
!!",!"#
= (1 − 0.04)𝑀! (6)
Where Mc is the cruise Mach number, stated above as 0.35. The fuel weight fraction is fairly linear up to a fraction
and decreases exponentially with Mach number. Incorporating the above fuel weight fraction, we solve for the
aircraft weight after climb with the following,
𝑊!"#$% = 𝑊!"
!!",!"#
!!"
!!"#$
!!",!"#
= 4,807 𝑙𝑏𝑠 (7)
The cruise to destination phase fuel weight fraction is calculated next using what is known as the Brequet range
equation for turbo jet engines. Re-arranged to solve for the fuel weight fraction,
(8)
As shown above, the fuel weight fraction (and thus fuel expenditure) increases exponentially with the range R,
thrust specific fuel consumption over velocity C/V, and the drag over lift D/L. These values are calculated in the
Excel formulation using the flight Mach number (to calculate velocity in ft/s) and the lift to drag ratio roughly
estimated (for Mach numbers less than one) by the relation, , where AR is the wing aspect ratio.
Substituting the climb weight above and solving for the aircraft weight after cruise results in,
(9)
This corresponds to a cruise range of 228 nm. After cruise, the loiter phase is determined from the following,
(10)
Substituting the cruise weight and solving for the loiter weight yields the following,
(11)
The landing phase is similar to the takeoff phase in that the same empirical formula for the fuel weight fraction is
used for landing,
(12)
The aircraft weight at landing is solved below,
(13)
It is expected that upon landing, the injured victim will be loaded to the aircraft within 10 minutes, before takeoff is
again initiated. After takeoff, the following phases are repeated using the same formulation as above; climb and
acceleration to cruise, cruise to trauma center, and landing. After the fuel weight fractions are utilized, the final
landing weight at the trauma center is estimated to be 4,131 lb with a total mission fuel weight of 921 lb. It was
found that two additional iterations were required to converge to a constant and final takeoff weight. The procedure
was again repeated for iterations 2 and 3 using the same formulation procedure and the weight values for each
iteration. A final takeoff weight of 3,245 lb was calculated, with a required fuel weight of 598 lb.
lim
C D
R
c b V L
cruise
W
e
W
=
/ 10L D AR= +
Wcruise
=
Wclimb
e
R
C
V
D
L
= 4,635lbs
D
EC
cruise L
loiter
W
e
W
=
Wloiter
=
Wcruise
e
EC
D
L
= 4,593lbs
Wlanding
Wloiter
= 0.985
Wlanding
= 0.985Wloiter
= 4,524lbs

B. Wing Loading Analysis
The wing loading was estimated based on an analysis of the cruise phase (since it’s the most significant) and
historical trends. The analysis for the cruise phase begins by stating the input parameters, which are the following:
Aspect ratio, 7
Initial cruise altitude, (ft) 8,000
Cruise Mach number, 0.35
Initial cruise weight, (lb) 3,119
First we will calculate the base drag coefficient, using the following equations,
(14)
The skin friction coefficient, is based on flat plate measurements for turbulent flow and is calculated as follows,
(15)
The form factor is likewise calculated,

(16)
An estimate of the interference factor for this scenario would be about 1.4, based on engine interference on the
wing. Finally, the ratio of the wetted wing surface area to planform area, must be calculated, this value is
approximately 2.039. The base drag coefficient is now calculated according to Eq. 14 as 0.0140. The wing loading
which gives maximum range is found when the parasite drag is three times the induced drag as follows,
(17)
Where and lbf/ft2
. Given the aircraft weight at initial cruise of 3,119 lb, a wing area
of 74.31 ft2
is found. Historical trends for aircraft indicate a wing loading of about 40 lb/ft2
for twin turboprop
aircraft, suggesting that the calculated concept wing loading of approximately 42 lb/ft2
is representative for this
type of propulsion system. Given the wing loading and the weight of the aircraft at initial cruise, a wing area of
74.30 ft2
is found.
C. Aerodynamics and Flight Performance

The aerodynamics analysis will begin with the wing. In order to select a suitable airfoil, we consider the cruise
conditions. This includes the cruise Mach number and the fact that lift equals weight, and verify that the design lift
coefficient of the airfoil (average lift coefficient during cruise) is within the drag bucket of the candidate airfoil. In
addition if possible, the range of lift coefficients from the beginning of cruise to the end must also lie within the
drag bucket of the airfoil to achieve a high and cruise efficiency. We will compute the lift coefficient at the
beginning of cruise as follows,
(18)
A
iH
crM
ciW
oDC
CDo
= Cf
FQ
Swet
S
fC
C f =
0.455
log10 Remac( )
2.58
1+ 0.144Meff
2
( )
0.65
= 0.00357
F = 1+
0.6
x / c( )m
t
c
!
"
#
$
%
&+100
t
c
!
"
#
$
%
&
4!
"
#
#
$
%
&
&
1.34M 0.18
cos Λt/cmax
( )
0.28!
"#
$
%&=1.396
wetS S
W
S
!
"
#
$
%
&
cruise
= q
CDo
3k
= 41.97−lbs/ft2
k =1 π Ae = 0.047 q =133.38
L D
CLi
≅
1
q
W
S
!
"
#
$
%
&
cruise,i
≅ 0.31

At the end of cruise, the dynamic pressure remains the same (constant cruise velocity and altitude) but the wing
loading decreases, making the lift coefficient at the end of cruise less than in the beginning,
(19)
The design lift coefficient is defined as the average of these two values, shown below,
(20)
Therefore, the airfoil we will select must contain a range of within the drag bucket. Additional
consideration is that the airfoil should generate a relatively high maximum lift coefficient with a low base drag
coefficient. A survey of NACA 5 digit airfoils, particularly the 64-412 that may be good candidate airfoil for this
concept reveals that the drag bucket extends from a section lift coefficient of 0.2 to 0.6, and a minimum base drag
coefficient of 0.0045. Moreover, the maximum lift coefficient is about 1.6 at an angle of attack of 16 degrees, and a
Reynolds number of 6 million. These results indicate that the design lift coefficient of the MedEvac concept is
within the drag bucket of this airfoil. Another candidate airfoil is the NACA 65-412, which has very similar
aerodynamics characteristics as the NACA 64-412 and 65-415.
The wing sweep angle, is defined as the angle between a line perpendicular to the aircraft's
centerline and the leading edge. Wing sweep is necessary primarily for high subsonic/ transonic Mach number
flight aircraft to reduce the effects of transonic and supersonic flow by increasing the critical Mach number of the
wing. Effectively the wing only sees the flow velocity perpendicular to the wing and thus the effective Mach
number becomes as follows,
(21)
where is related to according to,
(22)
The wing sweep angle will be selected considering cruise flight conditions and the following factors: The wing
aspect ratio, (which is effectively 10.35 with winglets, however we will assume the base aspect ratio of 9 for just
the wing), historical data of as a function of maximum Mach number, and the condition of "Pitchup" which is
undesirable. First, taking a look at historical data for general aircraft with a Mach number around 0.35, they have
been designed with low amounts of leading edge sweep. As such, we will impose a leading edge sweep of 3
degrees; the sweep at the quarter chord becomes 2 degrees. The taper ratio, is defined as the ratio of the wing tip
chord to the root chord, . Taper affects the distribution of lift along the wing span.
Figure 3. Wing planform
CLf
≅
1
q
W
S
!
"
#
$
%
&
cruise, f
≅ 0.28
CLdesign
≅ 0.295
LC 0.28−0.31"
#
$
%
LEΛ
coseff LEM M∞= Λ
critM LEΛ
1
cos
critM ∝
Λ
A
ΛLE
λ
ct
cr
0
2
4
6
8
10
12
0 2 4 6 8 10 12
y(ft)
x (ft)
Wing Planform

The most optimum lift distribution (minimum induced drag, or drag due to creating lift) has been shown to be
elliptical, created with an elliptical wing planform. Elliptical wings however are not practical to manufacture and
are rare. We will select a rectangular planform with a taper ratio of 0.75. A planform view of the wing is shown in
Figure 3, where the span is along the y-axis and the chord is along the x-axis. A summary of the wing geometric
parameters is presented in Table 1.
Wing area, S = 74.31 ft2
Taper ratio, λ = 0.75 Max thickness, = 0.12
Wing span, b = 22.8 ft LE wing sweep, = 3° wing sweep, = 1.4°
Aspect ratio, A = 7 1/4c wing sweep, = 1.8° TE wing sweep, = -1.7°
Table 1. Wing geometric parameters
An aerodynamic analysis of the cruise phase will be performed to estimate the aerodynamic forces (lift and drag)
during cruise conditions. Since the aircraft design is optimized for cruise, these results are most important. The
result of 3-D flow effects on a 2D airfoil is to reduce the lift coefficient, and increase the drag coefficient for a
given angle of attack (compared to a 2-D wing) effectively reducing the ratio. The lift curve slope for a 3-D
wing airfoil is thus reduced from a value of 0.1/ deg for the 2-D airfoil, to 0.087/ deg as given,
(23)
where the following parameters,
(24)
(25)
The wing lift coefficient is given as follows,
(26)
where the last term is the value of the lift coefficient at which is given as follows,
(27)
the zero-lift angle of attack, has a value of -2.5° obtained from the airfoil properties of the NACA 64-412.
Using Eq. 27 we can now solve for the angle of attack required at the beginning of cruise, This angle of attack is
called the trim angle of attack, which is . A the end of cruise, the trim angle of attack is .
A summary of the trim angles of attack and lift coefficients are below, where "1" and "2" refer to beginning and
end of cruise respectively,
(28)
The cruise speed and altitude Reynolds number of the MedEvac aircraft (based on mean aerodynamic chord) is
calculated as 7.47x106
whereas the airfoil data is for a Reynolds number of 6x106
.
The drag and lift associated with the wing will be calculated for cruise in this section. First we will
express the total drag coefficient as the sum of three drag components as follows,
max( )t c
LEΛ max( )t c max( / )t cΛ
1/4CΛ TEΛ
L D
dCL
dα
=
2π A
2+ 4+ Aβ( )
2
1+
tan2
Λt/c( )
β2
"
#
$
$
%
&
'
'
= 0.087−deg−1
β = 1− Meff
2
= 0.94
F =1+
0.6
t c( )max,x/c
!
"
#
#
$
%
&
&
t c( )max
+100 t c( )
4
1.34M 0.18
cos Λt/c,max( )
0.28
=1.36
CL
=
dCL
dα
α +CLα=0
0α =
CLα=0
= −
dCL
dα
α0L
= 0.22
0L
α
αtrim
=1.1° αtrim
= 0.7°
αtrim,1
=1.1°;CLtrim,1
= 0.31
αtrim,2
= 0.7°;CLtrim,2
= 0.28

(29)
The first term on the right is the base drag coefficient, the second term is the induced drag coefficient, and the last
term is a loss term if the lift coefficient during cruise does not encompass the drag bucket of the airfoil. In this case,
it does because the airfoil was chosen to have minimum drag during cruise, and so the last term becomes zero.
Beginning with the base drag coefficient,
(30)
The skin friction coefficient, is based on flat plate measurements for turbulent flow and is calculated as follows
with the value indicated,
(31)
The form factor is likewise calculated according to the following,

(32)
The next parameter to be calculated is the wing interference factor, Q which takes into account interference
between the wing and any component. A slightly higher form factor of 1.4 was utilized based on the envisioned
propulsion system which may include ducted rotors interacting with the wing. Finally, the ratio of the wetted wing
surface area to planform area, must be calculated, this value is approximately 2.023. The base drag
coefficient is calculated below,
(33)
We will now move on to the second term of Eq. 3.11. The induced drag coefficient is given as follows,
(34)
where, (35)
The elliptical wing lift parameter can be calculated more accurately from the suggested value of 0.8 with the
following formula,
(36)
where e' has a value of 0.98 suggested for a wide range of taper ratios and sweep angles. The fraction is the
approximate maximum fuselage diameter to wing span ratio, again has the value of 0.197. The induced drag can
now be calculated (for a beginning of cruise trim lift coefficient of 0.310) as the following,
(37)
The total drag coefficient for the wing as currently designed, during the beginning of cruise is thus estimated as the
following,
(38)
The drag coefficient at the end of cruise is calculated in the same way using the trim lift coefficient at the end of
cruise. The drag generated by the wing during at the beginning and end of cruise ("1" and "2" respectively) can be
calculated according to,
CD
= CDo
+ kCL
2
+ k ' CL
−CLmin,D
( )
CDo
= Cf
FQ
Swet
S
fC
Cf
=
0.455
log10
Remac( )
2.58
1+0.144Meff
2
( )
0.65
= 0.00311
F = 1+
0.6
x / c( )m
t
c
!
"
#
$
%
&+100
t
c
!
"
#
$
%
&
4!
"
#
#
$
%
&
&
1.34M 0.18
cos Λt/cmax
( )
0.28!
"#
$
%&=1.36
Swet
S
CDo
= 0.0121
CDi
= kCL
2
k =
1
π Ae
= 0.048
e = e' 1−
d
b
"
#
$
%
&
'
2(
)
*
*
+
,
-
-
= 0.94
/d b
CDi
= 0.00464
CD1
= CDo
+CDi
= 0.017

(39)
Finally, the lift to drag ratios, can be calculated using the trim lift coefficient and the total drag coefficient for
beginning and end of cruise,
(40)
The ratios take into account only the drag contribution from the wing (which is most significant); the
fuselage, and tail would contribute additional drag such that it would lower the actual flight lift-to-drag ratio. The
values of will be re-calculated once drag estimates for these components are found. Due to the VTOL nature
of the aircraft where lift is derived from the propulsion system, high lift devices such as flaps or slats on the wing
are deemed not critical to the design, particularly at this early development stage.

D. Fuselage Configuration

The primary purpose of the fuselage is to allow mounting of the flight surfaces, propulsion system/ engines and
accommodate the crew, passengers, baggage, fuel, and other supporting flight systems. The MedEvac concept is
designed to transport 2 passengers; the paramedic (who may serve in a limited capacity as a pilot) and the injured
victim. It is also designed for range of 525 miles (845 km) and must house the fuel required as well as the medical
equipment necessary. Other considerations include adequate structural integrity across the flight envelope, ease of
access to the main cabin for casualty loading/ unloading, and proper layout and electronic equipment for unmanned
operations and communication to ground base/ trauma center. When sourcing the equipment there were many
inter-disciplinary issues that needed to be considered. The weight must be minimized so as to maintain the
maximum takeoff weight calculated and ensure flight performance is satisfactory. Additionally, the power
requirements of the medical devices must be low and should ideally function on battery power for at least 2 hours.
Another requirement is that the devices should be user friendly. The conditions imposed by the use of a MedEvac
UAV is such that a personnel trained in medical first aid (similar to the qualification of a paramedic) will be
responsible for loading the victim and administering first aid while en-route to the trauma center. As such, the
MedEvac aircraft should be autonomous as much as possible, requiring limited attention from the paramedic.
Below is a table of the medical equipment deemed required with approximate weight, which will be used to
configure the fuselage. Moreover, the maximum weight of the victim and paramedic is taken as 220 kg (110 kg
each) for a total payload of 397 kg (873 lb). The dimensions of the dock and stretcher are considered when creating
the fuselage shape and internal volume. A general idea of the shape of the fuselage is necessary in order to define
the location and length of the main sections of the fuselage: the nose, cockpit, cabin, and tail sections. In addition, a
fairly specific fuselage shape is required to perform volumetric calculations to verify that it will satisfy flight
requirements and an aerodynamic analysis to estimate the drag particularly during cruise.
Table 2. Medical equipment and weight requirement
The fuselage design begins with an analytically definable shape and as it progresses through preliminary and
detailed design, it is modified due to practical considerations. There are many general fuselage shapes, the shape
for a particular design should be chosen based on type and purpose of the aircraft. We will select the shape known
D1
= qSCD1
=165.9−lbf
D2
= qSCD2
=157.4−lbf
L D
L
D
!
"
#
$
%
&
cruise,1
=
CLtrim,1
CD1
=18.52
L
D
!
"
#
$
%
&
cruise,2
=
CLtrim,2
CD2
=17.63
L D
L D
Item Weight (kg)
Dock: 1.26 m L x 0.3 m W x 1.085 H 76
Pro XT stretcher: 2 m L x 0.6 m W x 0.81 H 60
Stretcher locking mechanism 8
Medical equipment on dock 20
Equipment accessories in dock drawers 13
Total 177

as a Sears-Haack, The shape is described analytically through the relation for the top and bottom walls of the
fuselage,
r(x)
r(0)
!
"
#
$
%
&
2
= 1−
2x
l
(
)
*
+
,
-
2!
"
#
#
$
%
&
&
P
; −l 2 ≤ x ≤ l 2( ) (41)
The fuselage centerline (FCL) lies along the z = 0 line and the fuselage radius is given as function of x normalized
by the maximum radius (height or width of the fuselage). With Eq. 4.2, a profile for the side and top views of the
fuselage can be defined. The power, P can be varied along x to give a different line curvature which is used to
shape the fuselage. When 0P = , the radius equals the maximum radius and the function describing ( )r x is a
horizontal line. For higher values of P, the distance from the FCL, ( )r x decreases for a given x value and
decreases the diameter of the fuselage which sharpens the nose or tail of the fuselage. Taking this into
consideration, we will assume a maximum fuselage height of 5.75 ft which can be changed latter based on volume/
aerodynamic analysis. Since the top and bottom half of the fuselage is asymmetric, Eq. 41 is used to describe each
half with a different P distribution. This distribution is iterative and determined by visualizing its effect on the
shape. After some design iteration for fuselage requirements based on payload needs and historical aircraft, the
following fuselage shape was determined shown as a side view and top view shown on Figure 4 and 5 respectively.
The fuselage length is determined as 19 ft (5.79 m), the maximum height is 5.25 ft (1.6 m), and maximum width is
4 ft (1.22 m). Now that the fuselage geometry has been quantified analytically, we can perform aerodynamics
analysis to calculate the drag and its effect on the lift to drag ratio.
Figure 4. Side profile of fuselage
The fuselage viscous drag is given as follows,
(42)
The above variables are the same as for the wing, where again q is the dynamic pressure during cruise, S is the
wetted surface area of the fuselgae, Cf is the skin friction coefficient, F is the form factor, and Q is the interference
factor. The skin friction can be calculated for cruise conditions by using the summation method at each local stream
wise location along x to update the Reynolds number. The formula is the same as for the wing skin friction based
on flat plate experiments for turbulent flow,
(43)
-10
-8
-6
-4
-2
0
2
4
6
8
10
0 5 10 15 20
Height (ft)
Length (ft)
Top
fuselage
Ff
= qSCf
FQ
Cf
=
0.455
(log10
Reeff
)2.58
(1+0.144MC
2
)0.65

Figure 5. Top profile of fuselage
The fuselage form factor is given as,
F =1+
60
f 3
+
f
400
=1.085
(44)
where f is the inverse fineness ratio, or the fuselage length to vaerage diameter ratio of 4. The interference factor
will be equal to 1.2 since there my be engine pylons attached to the fuselage. The viscous drag is calculated as the
sum of 100 fuselage drag elements with their own local value of surface area and skin friction coefficient. The
fuselage drag is calculated according to Eq. 45 with the following value,
Ff
= DF
=187.8 lbf (45)
The drag coefficient corresponding to this drag is given as follows,
CD
=
DF
qS
= 0.0060 (46)
The fuselage drag coefficient and total drag values are 0.0060 and 187.8 lbf, in comparison to the wing drag
coefficient of 0.017 and drag of 166 lbf at cruise conditions.
E. Tail Configuration

The horizontal and vertical tail design deals with important preliminary design considerations such as the type of
tail arrangement, its placement on the fuselage, its size, as well as its aerodynamic characteristics. The first
consideration to determine is the tail arrangement. For this, we will restrict ourselves to the main types of tail
arrangements, which are: conventional tail, t-tail, and cruciform tail. A t-tail design places the horizontal tail on top
of the vertical tail. This design has two main advantages compared to a conventional tail; it allows the horizontal
tail to be made smaller because it is placed up high away from the wake of the wing. It also allows the vertical tail
to be shorter because the horizontal tail acts like a winglet which increases the aspect ratio and decreases induced
drag. In addition, a t-tail design is optimal for stall control; if the main wing stalls the wake created by the wing
-10
-5
0
5
10
0 5 10 15 20
Width (ft)
Length (ft)
Right
fuselage

will most likely not interfere with the horizontal tail and elevator because it is placed up high on the vertical tail.
The only main advantage that a conventional tail offers is that it can be made slightly lighter due to placing the
horizontal tail on the bottom of the vertical tail (less reinforcement required) as opposed to the top. Due to these
considerations, a t-tail design will be used for the MedEvac concept.
The size of the horizontal and vertical tail are determined from historical information using coefficients
that correlate the size of the tail surfaces with the size of the main wing which is known. The area of vertical tail
can be found from the equation,
SVT
= CVT
bW
SW
LVT
(47)
where VTC is the vertical tail coefficient. The value of this coefficient depends on the type of aircraft, a value of
0.08 was utilized for this aircraft design. The parameters Wb , WS , and VTL are the wing span, wing area, and
distance between the quarter-chord locations of the mean aerodynamic chords of the main wing and vertical
stabilizer. Considering the length of the fuselage and placement of the leading edge of the wing at about x/L of 0.5
and vertical tail at the rear, a tail to wing length of 8 ft was set. Substituting the wing span and area, we can
calculate the area of the tail as,
SVT
=16.94 ft2
(48)
Similarly the area of the horizontal tail can be given by,
SHT
= CHT
cW
SW
LHT
(49)
A horizontal tail coefficient of 0.8 was utilized suggested for turbo-prop aircraft. The parameter Wc is the wing
mean aerodynamic chord (3.28 ft) and HTL is the distance between quarter chords of the wing and horizontal tail.
Since the horizontal tail is mounted on top of the vertical tail, its mean aerodynamic chord is farther away from the
wing's compared to the vertical tail. We will estimate an additional 2 feet further based on the leading edge sweep
of the vertical tail, hence HTL is 10 ft. Using Eq. 49, the area of the horizontal tail is 19.50 ft2
.
The horizontal and vertical tail planform shape can now be determined based on desired tail aspect ratio
and taper ratio. These values are typically determined from historical data. The leading edge sweep angles are
usually higher in the tail than in the main wing to increase the critical Mach number more than in the wing. Below
are the tail geometric parameters that were used in the design:
Vertical tail aspect ratio, VTA 1.45
Horizontal tail aspect ratio, HTA 4.39
Vertical tail taper ratio, VTλ 0.70
Horizontal tail taper ratio, λHT
0.5
Vertical tail leading edge sweep, 25°
Horizontal tail leading edge sweep, 20°
The vertical and horizontal tail planforms are plotted below in Figures 6 and 7 respectively. Once the tail planforms
have been calculated, we can select the airfoils for the tail surfaces. These airfoils are selected upon two key
requirements: The need for a symmetric airfoil and a low base drag coefficient. A symmetric airfoil is needed
because the tail surfaces should not be creating lift at a zero angle of attack; only when the control surfaces (rudder
and elevator) are deflected should they generate lift. Thus, the effect of deflecting the rudder and elevator is like
adding camber to a symmetric airfoil creating a negative zero lift angle of attack similar to an airfoil with camber.
A low base drag coefficient is necessary for cruise conditions when the tail surfaces are not creating any lift (thus
no induced drag) and the only drag is due to the base drag and skin friction. However, once the tail surfaces are

deflected, we would like relatively large amounts of lift to be generated so that the aircraft can maneuver
adequately. This translates to a high lift curve slope, which is also desirable in wing design.
Figure 6. Vertical tail planform
Figure 7. Horizontal tail planform
Taking this considerations into account, we can select the following airfoils as candidates: The NACA SC(2) 0010,
NACA 0010-64, and the NACA 64A-010. The NACA 0010-64 is similar to the NACA 0010 airfoil, but moves the
maximum thickness point rearward from 30% to 40% of chord. This increases the drag divergence Mach number
(for low angles of attack) or the point at which the drag begins to rise sharply by several percent to about Mach 0.8.
The lift coefficient at a given angle of attack and Mach number also generally decreases slightly. We will use four
parameters to decide which airfoil to utilize: The zero lift drag coefficient, the lift to drag ratio, lift curve slope, and
the stall angle of attack. The Reynolds number for these values is at 100,000 and although it is not anywhere near
the flight Re, it can be used for comparison purposes. Below the airfoils are compared on Table 3. Based on these
characteristics, we will select the NACA 0010-64 airfoil for the vertical and horizontal stabilizers. Below is a
profile of the NACA 0010-64 airfoil.
dCL
dα
0
3
6
0 3 6
z(ft)
x (ft)
Vertical Tail Planform
0
1
2
3
4
5
6
0 1 2 3 4 5 6
z(ft)
x (ft) symmetry axis
Horizontal Tail Planform

Table 3. Airfoil aerodynamic characteristics
In this section we will calculate the tail drag during cruise. Much of the formulation here is identical to the
wing design thus discussion will be omitted where appropriate. We can express the drag coefficient as the sum of
the base and induced drag coefficient,
CD =CDo
+CDi
(50)
During cruise we assume that the aircraft is in trimmed conditions and flies in level flight. Hence the vertical and
horizontal tail produce a negligible amount of lift and we can neglect induced drag. The drag coefficient is
therefore given by,
CD
= CDo
= Cf
FQ
Swet
S (51)
The skin friction coefficient for the vertical and horizontal tail is given by,
Cf
=
0.455
log10
Remac( )
2.58
1+0.144Meff
2
( )
0.65
Cf , VT
= 0.00314
Cf , HT
= 0.00337
(52)
The tail form factors are given as follows,
F =1.1 1+
0.6
x / c( )m
t
c
!
"
#
$
%
&+100
t
c
!
"
#
$
%
&
4!
"
#
#
$
%
&
&
1.34MC
0.18
cos Λt/cmax
( )
0.28!
"#
$
%&
FVT
=1.39
FHT
=1.40
(53)
The factor of 1.1 is used to increase the form factor by 10% to take into account the rise in flow separation due to
the hinge gaps of the rudder and elevator. The interference factor, Q for the t-tail configuration will be chosen with
a value of 1.05, which is the same as for a conventional tail. The last term in Eq. 5.15 is the ratio Swet
S . This ratio
has the value of 2.03 for both the vertical and horizontal tail because they use the same airfoil. We can now
calculate the drag coefficients with the following values,
CD, VT
= 0.0093
CD, HT
= 0.010
(54)
Finally we can calculate the drag produced by the tail surfaces,
D = qSCD
DVT
= 21.04 lbf
DHT
= 26.22 lbf
(55)
The aerodynamic drag analysis of the wing, fuselage, and tail has been completed. The summation of these drag
components yields the total aircraft drag at the beginning of cruise, which is calculated as,

DAircraft
= 400.96 lbf
The aircraft cruise lift-to-drag ratio when the drag components of the fuselage and tail and considered is as follows,
Airfoil Zero lift drag Lift to drag ratio Lift curve slope Stall angle
NACA SC(2) 0010 0.027 14.46 0.058/deg 15
NACA 0010-64 0.013 28.55 0.053/deg 15
NACA 64A-010 0.014 24.35 0.053/deg 13

L DAircraft = 7.78

III. Turboelectric Propulsion Concept

The main requirements when selecting a suitable propulsion concept for this medevac aircraft design were
determined as the following: (i) attain a forward flight cruise speed of 235 knots (ii) be capable of vertical takeoff
and landing (VTOL) and the ability to quickly transition to forward flight with the same propulsion system, and
(iii) have a relatively small takeoff/ landing footprint to minimize risk in emergency landing areas. The cruise
speed requirement is beyond traditional helicopter average flight speeds of 160 knots limited by a large main rotor,
hence in this study we consider fixed-wing propulsion alternatives. Moreover, large rotors pose a safety hazard
during landing and loading of medical patients onto the aircraft. Another critical factor is the need to achieve
VTOL while allowing stable transition to forward flight. Among the propulsion concepts explored, turboelectric
propulsion emerged as the most viable. In the configuration envisioned, a turboshaft engine would power an
electric generator to develop electric power for 4 distributed electric motors integrated to ducted fans as shown on
Figure 8.

Figure 8. Turboelectric propulsion concept for medevac aircraft

In order to estimate the power requirements for the turbsohaft engine and motors, some preliminary calculations
were made given the takeoff weight of the aircraft. A curve fit empirical equation exists that approximates the take-
off thrust to weight ratio as a function of maximum Mach number for jet transports according to,
TTO
WTO
≈ 0.267Mmax
0.363
(56)
Our maximum Mach number is the cruise Mach number of 0.35. Substituting this value, we obtain,
TTO
WTO
≈ 0.182 (57)
Given the thrust to weight ratio at takeoff, a historical approximate value of takeoff thrust developed by the
propulsion should be 592 lbf. Below is a relationship that gives the engine thrust at a given altitude, H as a function
of the maximum take-off thrust at sea level (SL) and the pressure and temperature ratios, PH
PSL
and θSL
θH
,
TH
= TSL
PH
PSL
θSL
θH
(58)
At an altitude of 8,000 ft, the pressure and temperature ratios are given as,
PH
PSL
= 0.74;
θH
θSL
= 0.95 (59)

Performing thrust analysis at the cruise altitude, given Eq. 58 and an 85% of maximum thrust for sustained cruise
thrust (forward flight) yields a cruise thrust at 8,000 ft of 412 lbf (which is approximately equal to the total aircraft
drag at cruise) and un-sustained maximum thrust of 485 lbf. The takeoff thrust at sea level is 720 lbf which is the
total thrust developed by the engines forward flight,

Forward Flight: Tengines
= 720 lbf

(60)
During VTOL, we will assume that the propulsion system must generate 20% higher thrust than the takeoff weight
of the aircraft,
VTOL: Tengines
=1.2WTO
= 3894 lbf
(61)

With the thrust requirements for cruise forward flight and VTOL, we can now estimate the power requirements for
a turboelectric propulsion system.
A. Blade Element Ducted Fan Study
We will assume that a control system and pitch mechanism can be designed to collectively rotate the motor pods
for VTOL (vertical position) and forward flight (horizontal position) to direct the fan thrust as desired. The 4 fans
are envisioned to be mounted on pods on the side of the fuselage forward and aft of the wing. Given the thrust
required for VTOL and number of ducted fan motors, the required thrust per motor is as follows,
VTOL: Tengine
= 974 lbf
(62)
We will now use blade element theory to calculate how much electrical power is required to generate 974 lbf of
thrust, given a conceptual ducted rotor with the following characteristics given in Table 4.
Concept DF Rotor
No. of blades 8
Airfoil NACA 0012 (similar
data for other airfoils)
Rotor radius 1.5ft
Root cutout 0.41 ft
Root chord 5 in
Tip chord 3 in
Ideal twist
Ideal taper
Prandtl tip and root loss
Table 4. Concept ducted fan rotor characteristics
The blade element theory formulation utilized in this analysis is that as given by the textbook Principles of
Helicopter Aerodynamics by Leishman for hover conditions. The power and aerodynamics coefficients were solved
iteratively to meet the desired thrust. Below are plots of the blade angles, thrust and power coefficients as well as
the rotor operating conditions.
Figure 9. Rotor blade pitch, inflow, and angle of attack
0.00
20.00
40.00
60.00
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
Angle(deg)
x/R
Angle (deg)
EFF. ALPHA INFLOW ANG

Figure 10. Blade thrust coefficient
Figure 11. Blade power
Based on the 4 motor ducted fan requirement, the medevac concept would require a total power of about 1
MW of power for VTOL, which gives a 20% thrust margin above and beyond the calculated takeoff weight of the
aircraft. The total horsepower of 1456 Hp can be utilized to size the turboshaft engine, which must be capable of
generating approximately 10% higher shaft horsepower to account for losses in the generator and electric motor.
Table 5. Rotor performance thrust and power requirement per motor

B. Turboelectric Propulsion Model

The need for an integrated turboelectric model that incorporates the turboshaft engine, electric generator, and motor
is critical. Various analysis tools exist in the scientific community for each element, however one of the challenges
0.0000E+00
2.0000E-03
4.0000E-03
6.0000E-03
8.0000E-03
1.0000E-02
1.2000E-02
1.4000E-02
1.6000E-02
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
Ct-ThrustCoeff
x/R
Ct - Thrust Coeff
MOMENTUM THEORY
0.000
5000.000
10000.000
15000.000
20000.000
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
Power-ft*lb/s
x/R
Power - ft*lb/s
PROFILE POWER INDUCED POWER
Rotor Conditions and
Calculations
Rotational velocity 4650 RPM
Thrust (lbf) 998 lbf
Power (Hp) 364
Power (kW) 271
C_T 0.137
Cp/sigma 0.027
Ct/sigma 0.099

of creating a model and simulation is which specific tools can be leveraged (if any) and which tools must be
developed from the ground. In the turboelectric model presented here, preliminary steps were carried out in the
analysis and identification of tools that can be used to build the model.
The first component of the turboelectric propulsion system is the turboshaft engine. An analytical model
of the engine was constructed by analyzing a turboprop engine, and then making important assumptions about the
thermodynamic expansion cycle. A schematic of a turboprop engine is shown on Figure 12.
Figure 12. Schematic of a turboprop engine for thermodynamic analysis
The goal of the model is to calculate the shaft power based on a given set of parameters that define the operation of
the components upstream of the power turbine labeled as section 4.5 to 5 in the schematic. The first assumption
that is made that effectively allows us to treat a turboprop as a turboshaft is neglecting the power split between the
power turbine and the nozzle. The entire low pressure turbine (LPT) expansion process enthalpy goes towards
driving the shaft. Hence, there is no thrust generated by the nozzle. An analysis of the engine is done by
considering the stations on the engine beginning with the flight ambient conditions, M0, p0, T0 . The total pressure
and temperature are obtained using the isentropic tables for γ=1.4. The station 2 quantities at the face of the
compresor needed are the total temperature Tt2 and pressure pt2 that are given as follows, where πd is the inlet
pressure recovery.
Tt2
= Tt0
(63)
pt2
= πd
pt0
(64)
The station 3 quantities at the end of the compressor include obtaining the temperature ratio, τc from the
compressor ratio and polytropic efficiency, e
τc
= πc
γc−1
γcec
(65)

The input parameters for the burner at station 4 include the burner efficiency, ηb, pressure loss, πb, exit temperature,
Tt4, fuel heating value, QR, fuel-to-air ratio, f, and exit total pressure pt4. These quantities allows us to calculate the
HPT enthalpy at station 4.5 and the HPT pressure ratio,
ht4.5
= ht4
−
ht3
− ht2
ηmHPT
1+ f( ) (66)

πHPT
=
Tt4.5
Tt4
!
"
##
$
%
&&
γt
γt −1( )etHPT
(67)
The thermodynamic expansion cycle and power split, α between the LPT and the nozzle (N) for a turboshaft engine
is shown below. For the case of a turboshaft engine, we assume there is no power split (hence α =1) and the gas
expansion process and temperature change occurs in the LPT from station 4.5 to 5.
Figure 13. Thermodynamic expansion cycle and power split of a turboshaft engine
With the above assumption, we can express the power in the LPT of a turboshaft as follows,
℘LPT
= !m9
ηLPT
ht4.5
1−
p9
pt4.5
!
"
##
$
%
&&
γt −1
γt
!
"
#
#
#
$
%
&
&
&
= !m9
ht4.5
− ht5( ) (68)
To obtain the power delivered to the shaft, we define mechanical efficiency factors in the LPT and the gearbox
such that the shaft horsepower is the following,
℘shaft
=℘LPT
ηmLPT
ηgb
(69)
Finally, the power specific fuel consumption,
PSFC ≡
!mf
℘shaft
(70)
A preliminary test case was implemented where the objective was to determine what combination of engine design
parameters could generate approximately 1 MW of power at the cruise flight conditions of the medevac aircraft.
The input parameters are shown on Figure 13. The values may need to be refined to represent more accurate values
representative of turboshaft engines in this thrust class.
The engine outputs, specifically the shaft power of 1.09 MW is shown on Table 7. The shaft power is equivalent to
to 1,461 shp, which is required to generate approximately 1 MW of electric power at the output of the generator.
The last two components of the turboelectric engine model are the electric generator and motor. Prior to
modeling and identifying what tools can be leveraged to analyze these components, some research was conducted
into the state-of-the-art for the type of generators and motors being utilized for hybrid electric propulsion and
drives. It was determined that for high-power applications, 3-phase synchronous generators are the most common
and efficient. Moreover, having 3-phases allows the turboshaft engine to experience a constant load at all operating
speeds which is ideal. The standard model for a synchronous generator is often described as an equivalent circuit
with a rotor and stator containing inductances and coupled flux linkages which are a function of rotor speed. The

formulation will be omitted as it is beyond the scope of our purposes, however the coupling between the
mechanical power input and electrical output is shown in the following relation,
(71)
Table 6. Input parameters for turboshaft engine model
Table 7. Engine outputs; shaft power and efficiency
where P is the number of poles, ωrm is the rotor speed which is equivalent to the engine shaft speed. The
parameters, λ and i are the flux linkages and the electric current in the d and q axes as described in the generator
model illustrated in Figure 14. Simulations of generators have been built based upon circuit models. An example of
a Simulink model is shown on Figure 15 presented in the textbook Dynamic Simulation of Electric Machinery. The
model contains blocks for the q and d axes circuits, torque, speed, and rotor angle, as well as other flow variables.
The simulation can be used to determine the operational characteristics of the generator. The resistances and
reactance of the circuit model under desired initial conditions under a fixed voltage supply. The mechanical torque
is also an input parameter (based on the turboshaft engine) and the simulation outputs the generator current from
which electrical power can be calculated.
Pem
=
3
2
P
2
ωrm
λd
iq
− λq
id( )

Figure 14. Circuit representation of a synchronous generator
Figure 15. Block model simulation of a synchronous generator
The electric motor has also been previously modeled and simulated. Research suggests that high-power
applications including hybrid electric road vehicles utilize three-phase asynchronous induction motors. The
induction motor is an AC motor where the electric current in the rotor needed to produce to produce torque is
obtained by electromagnetic induction from the magnetic field of the stator winding. A block diagram of an
induction motor simulation is illustrated in Figure 16.
Figure 16. Block model simulation of an induction motor

The model is based on rotor and a stator with a magnetic core that contains laminations with three distributed stator
coil windings. The model contains blocks for the q and d axes and rotor. The inputs and outputs are connected to
sequentially numbered input and output ports as required to obtain the steady-state and small-signal models. The
objective of the simulation is to obtain the output torque and rotational speed for the given input parameters,
specifically voltage and current, as well as inductances within the motor structure.
IV. Conclusion
The conceptual design of a medical evacuation air vehicle designed to accomplish a specific mission was
presented. Based on the mission requirements, specifically higher speed cruise and VTOL capability, a
turboelectric propulsion system emerged as the most viable. This system consists of a turboshaft engine of
approximately 1,500 shaft horsepower (SHP) driving a 1 MW 3-phase synchronous generator and distributing
power to four 3-phase induction motors with ducted fans. Preliminary blade element theory calculations (neglecting
the effect of the duct) suggests that a total power of 1 MW is required for hover. A crucial step towards advancing
turboelectric propulsion technology for aircraft is the development of a unified model and simulation that
seamlessly incorporates the three elements mentioned above. In the preliminary work presented here, a different
approach was taken, a turboshaft engine was modeled analytically based on the operation of a turboprop and some
key assumptions. Existing models and simulations of a 3-phase generator and induction motor using Matlab/
Simulink were described. These models can be utilized as building blocks, together with a turboshaft simulation
model in order to create an integrated turboelectric engine model. NASA’s toolbox for modeling and analysis of
thermodynamic systems (T-MATS) which currently models various gas turbine engines is an attractive tool to
create such a tusboshaft model with separate blocks for the generator and motors. Finally, the flexibility for
distributed power with electric motors and propellers or ducted fans allows many motor mounting locations on the
aircraft and downwash effects, which interfere with the aerodynamics of the vehicle in different ways. Secondary
considerations for a turboelectric propulsion simulation would include these factors, which are important for
analyzing the flight dynamics and performance of the entire aircraft.
Acknowledgements
The authors would like to thank the NASA Transformative Aeronautics Concepts Program (TACP) / Convergent
Aeronautics Solutions (CAS) Project for supporting this research.
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