PSU Zephyrus

PSU Zephyrus: Design Of A Human-Powered Aircraft
For Sport
Alan R. Campbell,∗ Andrew J. Weinert,† Jason C. Slaby,∗ Kirk P. Miles,∗
Nathan T. Depenbusch,∗ and Kevin T. Show∗
The Pennsylvania State University, University Park, PA, 16802, U.S.A.
Following the announcement of a new Kremer Prize by the Royal Aeronautical Society,
the Flight Vehicle Design and Fabrication class at Penn State University has been working
on the design and construction of the PSU Zephyrus, a human-powered aircraft (HPA),
for the Sport Challenge. The design phase is currently nearing completion, testing of some
construction methods is ongoing, and full-scale construction is beginning. The Sporting
Kremer Prize mission requires the aircraft to traverse an equilateral triangle with 500meter length sides once in each direction. A total flight time of seven minutes or less is
required. The major factor in this competition is a wind requirement; during the entire
flight the wind must maintain an average speed of 5 m/s, and must not drop below that
limit for more than 20 consecutive seconds. The design process began with a survey of
previous HPAs and iterated toward the present configuration. For the “nearly frozen”
design, the pilot will be seated, reclined, in a pod hanging from the main boom and will
power a tractor propeller with a diameter of 3.0 m. The empty weight of the aircraft
will be approximately 25 kg and will require about 285 W output by the pilot. The wing
will consist of a straight center section (7.0 m) and outboard tapered sections (5.25 m)
each. The root chord will be 0.75 meters with a tip chord of 0.5 meters. The wing shall
be supported by a carbon fiber tube spar made up of two ±45◦ heat-shrunk sleeves and
14 plies of uni-directional carbon at the root. The empennage will be of a conventional
configuration with an all-flying horizontal tail for pitch control. The rudder will provide
yaw control, as well as some roll due yaw, which will also be supported with ailerons,
powered by local servos, on the outboard sections of the wing.
Nomenclature
C
L
R
Re
S
T
V
b
e
g
n
x
α
γ
General force coefficient
Lift force, N
Turning radius, m
Reynolds number
Reference area, m2
Tail volume coefficient
Velocity, m/s
Wingspan, m
Span efficiency
Gravitational acceleration, m/s2
Load factor
Reference length, m
Angle of attack, degrees
Climb angle, degrees
∗ Undergraduate Student, Department of Aerospace Engineering, 229 Hammond Building, University Park, PA 16802, Student Member, AIAA.
† Undergraduate Student, College of Information Sciences and Technology, 332 IST Building, University Park, PA 16802,
Student Member, AIAA.
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ν
ρ
ϕ
ω
Kinematic viscosity, m2 /s
Density, kg/m3
Bank angle, degrees
Rotational velocity, rad/s
Subscript
D
Total drag
H
Horizontal tail
L
Total lift
V
Vertical Tail
W
Wing
d
Two-dimensional drag
i, vor Vortex-induced
l
Two-dimensional lift
max Maximum
I.
A.
Background and Motivation
AERSP 204/404H
As a project-based learning experience, AERSP 204/404H (Sailplane) offers an opportunity for students to
work on real-world design issues under the shelter of a scholastic setting. One of the current projects is
the design and construction of a Human-Powered Aircraft (HPA) to compete for the Sporting Kremer Prize
offered by the Royal Aeronautical Society (RAS).
The HPA team is part of a Penn State course, AERSP 204/404H, Flight Vehicle Design and Fabrication,
affectionately known as Sailplane. Sailplane is an upper-level engineering and capstone design course taught
by Dr. Mark Maughmer. The course is partially funded as a member of the Space Grant Colleges and
Pennsylvania Space Grant Consortium and holds the following course objectives:
• Complete the preliminary design for an aircraft such that it satisfies assigned specifications
• Design a system, component, or process that meets given requirements in aircraft systems
• Identify, formulate, and solve engineering problems
• Function on multi-disciplinary teams
• Communicate and present effectively the results and consequences of their technical efforts
• Determine what the ethical responsibilities are to themselves, to employers, and to society
This course is unqiue because it of its strong “hands-on” component and that the curriculum is based
on the German Akafliegs. Starting in 1989, the curriculum has been broken into three components: lecture,
design, and fabrication; with projects typically spanning multiple semsters. The main purpose of the lecture
component is to provide the students with the theoretical background required for their design and fabrication
activities, the lecture topics are specifically requested by the team to address current design or construction
problems. The second component is concerned with design groups of four to six students, in which the
students design and analyze sailplanes, such as their performance, structure, stability and control, etc. The
third component is the fabrication of parts that have been designed and analyzed theoretically, with an
engineering lab assigned to the course for this purpose. Overall, the course gives students the experience
of a cooperative, multi-disciplinary team environment that is essential for solving problems related to the
design of an aerospace vehicle.1
In recent years the traditional class focus on sailplanes has shifted; with current projects including the
annual AIAA Design, Build, Fly competition and wing loading experiments, in addition to the Zephyrus.
The goal of all projects in the Sailplane class is to give the students a practical knowledge of the aerospace
engineering process, particularly focusing on aeronautics. The HPA was begun with this in mind as well. It
didn’t take long, however, for the traditional mindset of “let’s get this to fly” to be taken over by “let’s get
this to win.” It is a result of the added competitive drive and motivation from the Kremer Prize that has
pushed this design to its current state.
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B.
Kremer Prize
The Kremer Prize was initially funded by Henry Kremer, via the RAS, in 1959 for progress toward the goal
of human-powered flight. Two Kremer Prizes are currently available: a Sporting prize worth £ 100,000 and
a Marathon prize worth £ 50,000. Past Kremer Prizes have been awarded for completion of various tasks
in human-powered flight. Prior winners include Paul MacCready’s Gossamer aircraft, MIT’s Monarch B,
and the German-based Musculair aircraft. The Gossamer Condor won the initial Kremer Prize for its milelong oval flight in 1977. The next prize was won by the Gossamer Albatross in 1979 for flying the Engligh
Channel. Following the Gossamer’s victories, numerous smaller prizes have been awarded for improving
human-powered aircraft speed records. The current prizes were established in 2007.2
This particular mission provides several interesting design problems that must be resolved. To start, the
basic mission goal is to traverse an equilateral triangle with sides of 500 meters once in each direction in
seven minutes (see Fig. 1 for course details).
A particular area of interest, as noted on the course diagram, is the wind and its direction. The competition specifies a minimum average wind speed of 5.0 m/s during the flight. In addition, for a flight to be
considered official, the wind cannot drop below 5.0 m/s for a period of more than 20 seconds. The course
shall be aligned with the prevailing wind direction as shown.3
Figure 1. Course Setup for Sporting Kremer Prize
II.
A.
Design
Fuselage
Sizing the fuselage began by essentially measuring out a dimension range for the pilot, putting these dimensions in the correct orientation, and drawing a fuselage around it. Dimensions were based on a 5’10”
(1.78 m) pilot, with the assumption that someone of the necessary power to weight ratio to fly the aircraft
will be no more than this height.
Once the internal characteristics of the fuselage were completed, the external shape was considered. To do
this, a minimum volume “bubble” was drawn around the internal structure. The aerodynamic performance
of this shape was then considered. Constraints considered in this process include minimum widths for pilot
comfort and desired center of gravity of the aircraft. The shape of the pod was designed to be a low-drag
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body that will not generate lift regardless of angle of attack. In addition, the length of the shape was reduced
to allow for aircraft maneuverability in crosswind.
The internal structural members are designed to firmly hold the seat configuration in place, yet still
provide a maximum field of vision for the pilot. Structural members attach to the main boom at a hardpoint located behind the trailing edge of the wing. The chain will rise vertically from the gear to the drive
shaft to provide torque to the propeller.
B.
Propeller
To successfully complete the sport challenge, it was determined that the propeller would need to produce
27.5 N of thrust when cruising at 12.5 m/s and spinning at 135 rpm (as opposed to the 150 rpm by the
Musculair 2). To meet these requirements, the propeller design underwent many alterations during the design
process. Gottingen 796 (inboard) and 795 (outboard) were initally chosen, however the final design employs a
series of Eppler propeller airfoils: Eppler 858 at the hub to Eppler 850 at the propeller tip. Additonally during
the design process, the blade length was increased from 1.34 meters (based on Musculair 2) to 1.5 meters
which will yield an increase in efficiency. The propeller is theoretically more than 80 % efficient in the speed
range 10-16 m/s, which is advantageous given the chance of gusts during flight.4 Also, the propeller should
be capable of creating more than 20 N of thrust when the aircraft is still, enough to get the plane moving.
The blade is designed in two components: a structural shell of blasa wood, unidirectional carbon and kevlar
for the shear web and a kevlar fabric D-tube to take the torsional load.
C.
Wing
Starting from a parent aircraft approach, primarily using the Musculair 1 and 2, but also including the
Monarch B, Daedalus, and Velair models, a first iteration choice for an airfoil was made. A modified version
of the FX-76MP, as used for the Musculair 2 was chosen. Taking characteristics from this airfoil, using an
initial weight buildup, the wing planform size was determined. Then, assuming a take-off weight of 81 kg
(27 kg empty weight), sea-level air density, and using a CL at cruise of 0.8 obtained from the Musculair
reports on our chosen airfoil, a range of velocity values were obtained through the lift equation (see Eq. 1)
to examine the correlation between wing area and possible cruise velocity.5
1 2
ρV SW CL
(1)
2
From this, a range of span and chord lengths
were determined for varying velocities. Considering the time and distance requirements as stated in
the competition rules, a cruise speed of 12.2 m/s
was chosen. This cruise speed necessitates a wing
area of 10.90 m2 . A span of 17.5 meters was chosen
based on parent aircraft sizes and differences in mission requirements. This resulted in a mean chord of
0.62 meters.
Following a more accurate structural design and
weight build-up, further iteration of the planform
was performed. Beginning with a range of airspeeds, roughly 6.7 m/s to 13.4 m/s (15 to 30 mph),
and standard sea-level density and viscosity, the lift
equation was again used iteratively to size the wing.
This process resulted in a span of 17.5 m and a
mean chord length of 0.625 m, giving a wing area
of 10.93 m2 , a stall speed of 9.8 m/s (22 mph), and Figure 2. Profile, induced, and total wing drags for the
a cruise speed of 12.5 m/s (28 mph). The opera- PSU Zephyrus
tional CL ’s range from a CLmax of 1.3 to a CL of
0.7 at 13.4 m/s, with a cruise CL of 0.8. The calculated Reynolds number (Eq. 2) range at the mean chord,
over this velocity range, is from 440,000 at stall to 600,000 with cruise at 560,000.
L=
Re =
Vx
ν
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(2)
Following this step, an initial leading edge taper was introduced from a root chord of 0.75 m to a tip
chord of 0.5 m resulting in a taper ratio of 0.667. This taper results in an increased Re at the root of 530,000
at stall to 724,000 at maximum speed, with a cruise Re of 675,000. In addition, the airfoil selection was
changed to the Eppler 395 human-powered aircraft airfoil.
Further planform iterations were performed via a panel method drag build-up code written in MATLAB.
For this part of the process, the code was isolated for wing drag calculations. In addition, construction
ease and dynamic performance was taken into consideration. The current planform consists of a 7.0 meter
rectangular center section with a chord of 0.75 meters and two outboard sections each 5.25 meters in length.
The outboard sections are straight-tapered from 0.75 m to 0.50 meters at the tip. In addition, the trailing
edge is now the tapered side to allow for straight spar connections. The wing area is 11.81 m2 .
Drag calculations were performed for a set number of wing panels. The mean chord of each panel was
used to determine a panel Reynolds number. XFOIL results for varied Re were interpolated for the twodimensional lift and drag coefficients (Cl and Cd ) for each panel.6 These were then integrated over the
spanned to determine the wing lift and drag coefficients (CL and CD ). Induced drag was calculated via
Eq. 3. See Fig. 2 for the calculated drag of the wing design.
CDi,vor =
D.
CL2
2
πe bS
(3)
Spar
Two spar structure options were considered, an Ibeam and D-tube construction as well as a tube spar
with additional caps for bending. To decide between
the two, rough weight calculations were preformed.
For the tube spar calculations, the structure was
broken into two parts: a tube to take torsion loads
and caps to take bending loads. For the I-beam
and D-tube method it was assumed that the I-beam
took all the bending and the D-tube took all the
torsion. Scaling the E395 airfoil to the wing platform and using the maximum thickness point as the
location of the spar, the size of the structure in both
methods was calculated. Applying a linear lift distribution the weights of the structures required were
compared and it was found that the tube spar structure was lighter than the equivalent I-beam.
The tube spar renders the D-tube as a primarily aerodynamic feature, taking neither bending nor
torisonal loads. The D-tube would then only be a
layer of finished glass or sheets of foam 3-5 mm thick,
similar to past HPAs. Using this assumption, 10 mm
was subtracted from the thickest point of the airfoil
to allow for the D-tube. This new outer dimension
became the furthest distance from the two caps on
Figure 3. Cross-section of the wing spar at the root.
the spar.
The width of the spar caps was determined based
on the size of the workable area, considering that wider caps bending around the tube would be less effective.
For a first iteration, the caps were set as geometric rectangles and a linear lift distribution was assumed.
Basic beam bending calculations were completed to set the initial cap thickness. After it was seen that
the thickness seemed reasonable and fit inside the allowable distance, a second round of calculations were
preformed using an elliptical load distribution. The number of cap plies was determined by tabulating the
calculated thickness and dividing it by the thickness of unidirectional carbon. Since composites are subject
to large stress concentrations at the edges of layers, great care was taken to reduce the number of drop-offs
of plies along the spar caps. The thickness of the top and bottom caps was then accounted for in the outer
diameter of the torsion tube (see Fig. 3 for dimensions).
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This new outer dimension underwent basic beam torsion and aerodynamic loading calculations to determine the wall thickness of the tube along the span. The wall thickness was subtracted from outer dimension
to determine the minimum diameter of the carbon tube. It was determined that two full carbon sleeves
would be required for the tube.
E.
Control Surfaces
The control system was one of the hardest aircraft components to design because of the lack of data available
and the mission specific design of the aircraft. It has been noted that on previous Kremer Prize winning
flights, the pilot is constantly actively attempting to control the aircraft. For example, Bryan Allen in
the Gossamer Albatross could not even look at his watch during the 2 hour and 49 minute flight due to
constant adjustment of the controls.2 The necessary pilot input resulted in a design approach where multiple
calculations of the same parameter with different variable values was used. This results in a lack of constant
variables. Each calculation, however, had a key variable identified. Also, a majority of the control parameters
were calculated at the cruise velocity of 12.5 m/s. Since each of the control surfaces effects all of the others,
one had to be chosen as the starting point. The ailerons were chosen since more prototyping had been
conducted on the wing, along with the aileron sizing being necessary to determine D-tube specifications and
rib spacing.
First, the flight mechanic parameters of turning radius (Eq. 4), load factor (Eq. 5), centripetal acceleration
and force, and angular velocity were calculated for a range of bank angles from 10◦ to 45◦ at 5◦ increments.
These parameters were calculated using a student generated Java GUI, increasing time efficiency. All parameters determined at a bank angle of 30◦ were confirmed through hand calculations to ensure accuracy
(see Eq. 6). Based on the course layout and structural strength of the aircraft, 30◦ was targeted as the ideal
bank angle. The aircraft also must be capable of operating at other bank angles because the desired bank
angle will most likely not be achieved every turn. The load factor produced by a 30◦ turn was determined to
be 1.155, however, an averaged value, from other bank angles, of 1.162 was used. The averaged load factors
have only a percent difference of 0.63% from the the desired load factor, this is within the 1% tolerance
specified by literature.7 Basic roll oscillations were also calculated to measure the damping forces along the
wing during a turn, as a result of the bank angle.
V2
g × tan(ϕ)
(4)
cos(γ) p 2 2
ω V + g2
g
(5)
R=
n=
1
ϕ = arccos( )
(6)
n
The hinge moment was calculated next using the same methodology to be 0.88 N-m with a moment of
inertia of 2.22 N around the spanwise axis. The hinge moment was set at 1.06 N-m instead due to the servo
control design; many servos are rated for 150 oz-inches, which is 1.06 N-m.8
Multiple tail hinge moments were calculated based on velocity and angle of attack. Due to the control
design, the hinges take majority of the load, and therefore, analysis of this structure becomes a priority. It
was important to note the difference in the moment at different speeds because the lack of a constant power
output. Using XFoil and an MS Excel spreadsheet, the hinge moments were calculated and the max and
min was identified for each velocity. The tail control elements were then based off these max and mins.8 See
Fig: 4 for hinge moment as a function of α.
The aileron geometry was then sized, dependent upon two factors: percent chord length and percent
span length. Additionally density (ρ), velocity, and climb angle (γ) were constrained to be constants at
1.28 kg/m2 , 12.5 m/s, 0.0553◦ respectively. The calculated aileron weight was minimized and the force on
the surface along the control span was maximized in the calculated range. Using the values from the flight
mechanics process and comparing them to the calculated force on the main surface along the control span
values, a range based upon percent span length was identified. Based upon aileron weight and force on the
main surface along the control span, along with the overall geometry of the outboard wing sections, ailerons
spanning 5 meters on each outboard section were chosen with a chord of 0.1425 meters. These ailerons will
cover an area of 0.428 m2 and were estimated to weigh 592 grams.
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Moment (Nm)
With the size calculated, three
aileron designs were considered:
Kevlar strip, submerged, and piAngle of Attack vs Tail Hinge Moment
20.000
ano hinge, all of which could be
controlled via servo motors. The
15.000
PSU DBF teams utilize the Kevlar
strip method, which consists of con10.000
structing the wing and cutting the
wing at the hinge point into two
5.000
sections, which would be then reconnected with a spanwise Kevlar
0.000
-15
-10
-5
0
5
10
15
20
25
strip. Although this is a low weight
-5.000
design, it is too weak for a full
scale aircraft and does not provide
-10.000
enough structure. The submerged
method is utilized in professional
-15.000
Angle of Attack, alpha (degrees)
sailplanes; the front of the aileron
14.75 m/s
12.5 m/s
10 m/s
is rounded off and mounted to the
airfoil via socket and control arm.
This method provides the greatest structural support but exceeds
weight requirements, along with be- Figure 4. Aileron hinge moment as a function of angle of attack for 10,
12.5, and 14.75 m/s
ing the most complex considered design.
The current design incorporates a piano hinge attached near the upper surface, so the upper surface
incorporates the leading edge radius, so that the upper surface maintains smooth flow. The bottom surface
can be covered with Kevlar or a tape which moves easily when the aileron deflects downward. Two equal
spaced rods are mounted on the drag spar, and a single rod is placed on the aileron. The rods are then
aligned and strung together using piano wires that are run through the larger mounted rods. This design
was selected based on its weight efficiency and ease of construction. Furthermore this design has the simplest
attachment of the control rods and servos.
Initial tail sizing was performed using a parent aircraft approach, primarily focusing on the Musculair.
Since the Zephyrus has a similar mission to that of the Musculair 2, it was deemed acceptable to estimate
chord and span dimensions for the horizontal and vertical tails. The horizontal planform is rectangular while
the vertical planform features swept-back rectangular sections above and below the boom. The horizontal
distance of the elevator from the wing was calculated using a chord of 0.6 meters and a span of 2.0 meters
and a tail volume coefficient range of 0.5 ≤ TH ≤ 0.65. For the vertical tail the volume coefficient range is
0.02 ≤ TV ≤ 0.05. The chord and span are 0.75 meters and 1.75 meters respectively. Until the sweep angle
is determined, a rectangular platform for the vertical tail is estimated.5
III.
A.
Construction and Testing
Wing
Wing construction was divided into three major sections: spar, ribs, and D-tube. Spar construction was the
most challenging, with various construction methods being considered. Three particular methods included
winding carbon rovings to form the tube, constructing a half-mold to make a permanent inner tube of
fiberglass around which the carbon would be wrapped, or constructing a removable foam core. The foam
core method was chosen because winding the carbon rovings is not feasible for the scale of the spar and
there is difficulty in joining the two half-molds. The core was constructed out of large cell polystyrene foam
which is dissolved by acetone. Then, it was wrapped in painter’s plastic and sealed with tape and Super 77
spray adhesive to prevent the acetone from affecting the carbon tube. After the carbon and epoxy was set,
the tube was wrapped in shrink tube or tape (depending on the construction step) and applied heat via a
heat gun to expel excess epoxy.
Rib construction consisted of cutting out the rib shape (E395 airfoil) out of large cell polystyrene foam.
To allow for the aerodynamic D-tube a notch the thickness of the D-tube foam was cut into the rib shape
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around the front half of the lower airfoil and around the top surface as far back as the estimated transition
point. Unidirectional carbon was laid onto the foam covering the entire outer surface to add structural
stiffness. After the carbon cured, 5 mm wide ribs were cut out from the airfoil shaped foam using a band
saw. No lightening holes were included in the design, because tests indicated there would be an additional
cost of adding structural support causing an increase of weight.
The D-tube was constructed from small cell
Foamular 250. A CNC
hot-wire machine was
used to cut out one meter wide and 0.61 meters long sections of Foamular at 0.0036 meters
thick.
The Foamular
strips were trimmed to
0.46 meters long for use
on the top surface and
0.30 meters long for use
on the bottom surface
Figure 5. Wing rib cross-section
(see Fig. 5 for rib crosssection).
B.
Fuselage and Main Boom
The structure for the fuselage and main boom will be constructed in a similar method as the wing spar. The
main difference is that the fuselage and boom must be designed for point loads rather than the distributed
lifting load on the wing. For the main boom the structure is divided into two parts: the front, from the
propeller to the rear connection of the fuselage, and rearward from the fuselage to the connections with the
stabilizers. The front section is practically identical to the root section of the wing spar with carbon sleeves
making a tube for torsional loads and unidirectional carbon strips for the bending. The aft section is more
complex because it is a long tapering tube subject to point loads at the end. This leads to a greater threat
of buckling. There are two general ways to solve the buckling problem. The first is to add more structure,
in the form of carbon, to make the ultimate buckling load greater than what will be seen in flight. The
problem with this method is that a great deal of weight is added. The second option, which has been used
by previous HPAs, is to make the walls of the tube a sandwich beam. This is the option that will be tested
with a removable foam core to lay-up around, similar to the spar. Then a layed-up carbon sleeve on the
inside is followed by a thin layer of foam and finished with cloth and unidirectional caps. The advantage to
this is method is that it reduces weight while increasing the wall thickness, further avoiding buckling.
The fuselage structure to hold the pilot is also made out of carbon tubes constructed with the same
removable foam core method. The fuselage structure will see a direct air stream, unlike the spar; therefore,
it will not be a round tube. Most will be oval in shape to reduce their drag and to fit into a thin glass fairing
for aerodynamics. The connection points will use foam inserts for gussets and strips of carbon for bindind.
For the adjustable pedals, a thin aluminum tube will be bonded inside the lower fuselage tube during its
construction. This will allow bolts to be drilled through to hold the sprocket and pedals.
C.
Control Surfaces
Aileron construction follows the full scale rib construction. Ribs are first marked according to whether
they support an aileron. Then, the marked ribs will be cut into two, based upon the aileron chord of
0.1425 meters. The aileron ribs will then be attached to a support structure of 2 ounce, fiberglass, sandwich
panel. In addition, another sandwich panel will be constructed to act as the drag spar. The drag spar will
be attached to the ribs on the wing with an epoxy-cotton flocks thick mix. Carbon tubing will be fashioned
to the drag spar and aileron support spar to act as the hinge. There will be two hinges on the drag spar to
each on the aileron. Each aileron will have three pairs of hinges placed 1.66 meters apart. Piano wire will be
run through each hinge pair to join. Local 1.0 N servos will manipulate a control arm for each hinge pair.
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Each servo will be placed on the trailing edge of the wing supported by a double rib junction. The aileron
will have a balsa wood trailing edge and will be covered with Mylar.
The empennage construction method will model that of the wing. An all-flying horizontal tail is utilized
for ease of construction, reduced weight, and increased performance. The elevator will be mounted at its
quarter-chord so that excess weight is not added for structure. A standard configuration vertical tail will also
be used. The vertical stabilizer will be mounted offset from the boom’s centerline and the rudder attached
via carbon tube hinge. Both the elevator and rudder will be controlled mechanically using a pull-pull system
of braided Kevlar control rods and nylon pulleys.
D.
Weight Estimate
The weight of the HPA is very important due to the limited power available. Weight placement is also a
delicate balance; if focus is entirely on weight, it will not be strong or rigid enough structurally. Therefore,
the weight and structure were the object of a careful trade study to balance necessary power and the strength
to handle flight loads.
At each stage of the design and construction process, a weight buildup for the aircraft was performed.
Since the wing has been most extensively designed and tested, it has the most detailed weight breakdown.
The amount of epoxy used in the composite lay-ups was a major concern; the optimal ratio is dependent
on the material but generally is 30-40 % epoxy by weight and never more than 50 %. Professional lay-ups
utilize pre-preg carbon and high pressure autoclaves to produce a finished lay-up with 33-35 % epoxy by
weight. “Homebuilder” lay-ups utilize a wet lay-up and shrink tape or vacuum bagging resulting in a ratio
of about 50 % epoxy by weight. Through prototyping, it has been demonstrated that for relatively large
lay-ups, 45 % epoxy by weight can be achieved. This was then set as the goal ratio.
The base weight was calculated as the total weight of the carbon used in construction. This was calculated
using density of the carbon and multiplying by the epoxy ratio; this estimate for the wing is 9.5 kg. The
total rib weight was then calculated by measuring test ribs and calculating the required number of ribs based
on a spacing of 0.33 meters. This estimate was calculated to be 1.28 kg, increasing the wing weight estimate
to 10.78 kg. The aerodynamic D-tube of Foamular 250 was calculated next using test cuts leading to weight
per unit area calculations. XFOIL was used to find the transition point on the airfoil to determine the area
of the airfoil to be covered. This was used to calculate a D-tube weight estimate of 1.3 kg, which increased
the wing weight estimate to 12.08 kg. The Mylar covering and structural hard points along with the drag
spar and servos were the last components to add to the estimate, increasing the total weight to 14.1 kg.
This estimate was used to scale the other aircraft components based on the assumption that the vertical
and horizontal stabilizers have similar construction methods and structures to the outboard wing sections.
The weight estimate of the empennage was based on the outboard sections scaled to the size of the each
stabilizer, which results in a weight of approximately 1 kg each. The main boom and fuselage weight estimates
were scaled based on the main spar. This was done assuming that a majority of the forces on the spar are a
result of the pilot’s weight in the fuselage. Adding the fuselage, propeller, control stick, and the drive-train
weight estimates to the wing and stabilizer weight results in a total weight of approximately 25.7 kg.
Comparing our results with previous HPAs, our estimate appears reasonable. The Velair 89 and the
Musculair 2 had empty weights of 30.67 kg and 25.0 kg respectively. Also, each had slightly larger wingspans
than the Zephyrus but were not designed to fly in adverse wind conditions requiring higher wing loading.
The PSU Zephyrus wing weight estimate is also comparable with the Velair 89’s wing weight of 15.96 kg.
IV.
Progress and Future Work
The team is currently finalizing the testing phase and has begun full-scale construction of major components. The wing spar, ribs, propeller and fuselage are under construction. A working propeller-powered
tricycle has been built and is undergoing thorough testing of propeller thrust and drive-train components.
Further dynamic stability calculations are being performed to ensure appropriate control surface sizing, and
detailed fuselage and empennage drag components are being implemented into the drag build-up.
The next step is to continue construction of Zephyrus to the point of completion. Then, intensive flight
testing will be performed to confirm the aircraft’s ability to perform as expected. If necessary, additional
design iterations will be constructed to improve performance. Eventually, transportation must be obtained
to an undetermined location in the United Kingdom to, hopefully, claim the Sporting Kremer Prize for PSU.
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Appendix
Figure 6. PSU Zephyrus Three-View (1:20 scale)
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Penn State Zephyrus
Theory
Propulsion
Flight
Geometry
Parent Aircraft
Wing Airfoil
Propeller Airfoil
Propeller RPM
Propeller Thrust (N)
Cruise Velocity (m/s)
Stall Reynolds number
Cruise Reynolds number
Minimum Reynolds number
Root cord (m)
Mean Cord (m)
Tip Cord (m)
Wingspan (m)
Half-span (m)
Trailing Edge Length (m)
Tapered Section Area (m2)
2
Weight
Musculair 2
E395
E858 / E850
135
27.5
12.5
600,000
560,000
440,000
0.750
0.6096
0.500
17.5
8.75
5.274
2.40
Total Wing Area (m )
11.8
Total Aileron Area (m2)
Aileron Deflection (degrees)
Epoxy by weight ratio (%)
Wing weight estimate (kg)
Total weight estimate (kg)
0.428
20°
45
14.1
25.7
Figure 7. PSU Zephyrus Summary Table
Acknowledgments
First and foremost, the authors would like to pay homage to Harry Kremer and the RAS, the late Paul
MacCready and his Gossamer team, Dr. Mark Drela and the MIT Daedelus team, and Prof Günther Rochelt
and his Musculair group for laying the foundations in human-powered flight. Without their contributions,
the field of human-powered flight would truly still be struggling to get off the ground.
The authors would also like to collectively thank Dr. Mark Maughmer, Professor of Aerospace Engineering, Pennsylvania State University for inspiration and theoretical support in our attempt at the prize. In
addition, we would like to thank Dave Benson and DuPont for their generous support of the project. Last
but not least, we could not be where we are and cannot get to where we want to go without the help of the
rest of the AERSP 204/404H class.
References
1 Bramesfeld, G. and M. Maughmer, “The Penn State Sailplane Project - The First Decade”, Pennsylvania State Univeristy,
20th AIAA Applied Aerodynamics Conference, St. Louis, Missouri, June 24-26, 2002.
2 Grosser, M., The Gossamer Odyssey, Zenith Press, St. Paul, MN, 2004.
3 Royal Aeronautical Society, Human Powered Flight Group Website, http://www.raes.org.uk/cmspage.asp?
cmsitemid=SG Hum Pow Home. [cited 15 November 2007].
4 Schenk, H., Propeller Calculator, http://www.drivecalc.de/PropCalc/index.html, [cited 20 January 2008].
5 Thomas, F., Fundamentals of Sailplane Design, College Park Press, 1999.
6 McCormick, Barnes W., Aerodynamics, Aeronautics, and Flight Mechanics, John Wiley and Sons Inc. 1995.
7 Pajno V., Sailplane Design, Salvo D’Acquisto, 2006.
8 Morelli, P., Static Stability and Control of Sailplanes, Levrotto and Bella, Torino, Italy, 1976.
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