The
University of Manchester and UMIST 3rd Year Aerospace Engineering
Design Project
Design of an
Autonomous Surveillance UAV
Preliminary Design
Report
1st
November 2002
Group H:
Tim Myall
Andrew Walker
Keith Watson
Russell Hankey
Andrew Chan
Mark Kathurima
Summary
A
customer requires an autonomous unmanned flight vehicle capable of taking high
quality footage of a remote target.
This report considers a range of potential approaches to the customers
design specification. These approaches
are considered with reference to the advantages and disadvantages of their
application using calculations, current designs, and practical engineering
knowledge. The conclusion describes a
design that is optimal with respect to the design specification; this will be
taken on to the detailed design phase.
Contents
2 Analysis of
the Specification
3.2 Basic
aircraft configuration
4.1.1 Power Plant
Considerations
4.1.2 Engine and
Propeller Choice
4.4 Fuel
Fraction and Weight Estimation
4.7.1 Take-off and
Landing Considerations
4.8.4.1 Wing
Position on Fuselage
4.8.4.2 Canard or
Conventional Tailplane?
4.8.4.3 Vee-Tail or
Conventional tailplane?
4.8.4.4
Longitudinal Stability
4.8.4.6 Control
Surfaces Configuration
4.9 Materials
and Manufacturing
4.9.1 Aims of
materials selection
4.9.4 Method of
attachment: Wing to body
A2-8 Alternative
Take-off/Landing Calculations
Wing and Tailplane
Positioning
1 Introduction
The use of Unmanned aircraft for reconnaissance missions began in World
War Two, but it was not until the early 1980’s when Israel made the first
significant use of drones to watch over Syrian and Palestinian forces in
Lebanon [1] that the world realised their potential. The future for UAV’s looks
bright with Pentagon officials predicting that by 2015 10% of all combat strike
aircraft will be unmanned [1] There have been a wide range of UAV’s developed
for both surveillance and attack missions, some of the most successful from a
military point of view have been the Predator [2] which fired the first live
missile at a target in Afghanistan and the Global Hawk a proven reconnaissance
UAV again in Afghanistan. [3]
This aim of this project is to design a UAV to meet the specification
detailed below; this report documents the preliminary design process. The end of this report details a design for
the aircraft that is able to be further optimised in the detailed design phase.
2 Analysis of the Specification
2.1 Customer Specification
Below is a condensed version of the customer’s specification for the
vehicle:
Ø
It must be an
autonomous unmanned flight vehicle.
Ø
Have the ability
to take high quality video footage of a remote target.
Ø
Have a range of
600km.
Ø
Must be able to
maintain height of 1000m for one hour while above target.
Ø
Have an overall
mission duration less than 8 hours.
Ø
A tarmac strip of
100m length can be used for takeoff and landing, however alternative systems
may be used.
Ø
Should be modular
in construction so it can be easily transported, and serviced.
Ø
Components should
be commercial off the shelf (COTS) as far as possible.
Ø
An avionics
package is supplied by the customer consisting of a flight control computer
unit, GPS, air data system, inertial navigation system and battery. The crate has dimensions 110 x 110 x 200mm
and mass of 1 Kg.
Using the data given in the specification a number of important
parameters can be determined for the aircraft
2.2 Minimum speed
The aircraft must have a range of 600 Km, which means possibly flying a
maximum of 1200km plus the flight distance while loitering above the
target. The maximum mission duration is
8 hours, which gives 7 hours to travel to and from the target. Considering the unlikely case with no head
or tail wind then the minimum speed that the aircraft can cruise at is
This speed doesn’t take into account the takeoff and landing manoeuvres
of the aircraft nor does it consider the climb and descent from the cruise
altitude. As an initial estimation the
minimum cruise speed of the aircraft has therefore been taken as 50 m/s.
3 Ideas Considered
3.1 Type of flight vehicle
The customer has not specified the type of aircraft required; however
the range of the vehicle and the time constraints on the mission duration are
important to consider. Many UAV’s such
as the Yamaha RMAX remotely piloted helicopter [4] and the guardian CL-327
tactical UAV VTOL [5] have rotating wings, however these vehicles do not have
the range required by our customer and it is difficult to see any way that a
helicopter could meet the mission specifications. Certain designs including the Bell Textron Eagle-eye [6] UAV use
tilting propellers that can be used to enable vertical takeoff, but can also
cruise like a conventional aircraft, these aircraft are relatively new to the
UAV scene and are considered too difficult to design in a project such as
this. Similarly the use of a blimp or
airship which have been used as successful UAV’s such as the AV Pointer Micro
Blimp [15] is generally
restricted to low speed and low range missions and is therefore not applicable
to this project. This leaves the option
of a fixed wing aircraft.
3.2 Basic aircraft configuration
To meet the mission specifications, the aircraft only has to fly at low
speed. This has a number of benefits in
terms of the design of the vehicle; no compressibility effects have to be taken
into account, no shock waves will be present hence there is no need for the
wings of the aircraft to be swept. This
also means that a conventional piston engine powering a propeller is more
adequate than a jet to allow the aircraft to perform its mission. One of the first
pieces of equipment that must be decided for the UAV is the method of propulsion.
There are many different methods that are theoretically available. For example
piston propeller, turbo-propeller, turbo-fan, turbo-jet, Ramjet, rocket, etc.
However most of these are not feasible for the particular application. The
piston prop arrangement is the only configuration that was considered for
further analysis because of its low weight and good fuel consumption. Also a piston prop engine is most efficient
at low speed, which is the application here. Rocket propulsion may be
considered for take-off assistance due to the short field length.
4 Technical Analysis
4.1 Propulsion
4.1.1 Power Plant Considerations
Despite the fact that the piston prop is one of the simplest arrangements of propulsion unit, there are many variations of configuration possible. There is the possibility of using a high torque engine with a large propeller. This is a very efficient way of producing power as the blade can spin slower, reducing the risk of tip separation and flutter effects. However, engines that produce more torque tend to be much larger and heavier than the smaller racing engines. There is also the problem caused by using a large blade to consider. The ground clearance on take-off and landing will have to be greater to eliminate the possibility of a propeller tip striking the ground during the take-off roll or flare. Also a larger propeller will add to the aircraft’s Operational Weight Empty (OWE).
In order to fly at a greater speed there are two obvious possibilities. Either use two (or more) smaller engines or one more powerful engine. The use of a single more powerful engine (such as the model shown in fig 1) was seriously considered in the design. This would have enabled the design to complete the specified mission with time to spare. However there is a lack of technical information for small model aircraft engines. Despite extensive research, no data concerning fuel consumption was found either in books, on websites or in model shops.
Fig. 1 – OS Engines 1.40 RX 3.5 BHP engine. Impressive power
but no fuel consumption information.
Using two smaller engines was also dismissed because, although the aircraft power has been doubled, the weight has also increased. This would cause many knock-on effects on other technical areas such as structures, materials and stability (depending on engine arrangement). Configurations such as those shown in fig 2 were considered.
Fig. 2 –
Different engine configurations using two less powerful piston engines.
4.1.2 Engine and Propeller Choice
It became clear that our choice of engine was fixed to the OS Engines FS-91 (fig 3). This is the only engine in its class (i.e. model aircraft engines) that reliable data for its fuel consumption is available.
Fig 3 – OS Engines FS-91, which will be used to propel
the UAV
A single engine configuration was decided in order to maximise the power to weight ratio. This single engine will be arranged in a tractor design as shown on the drawing of the preliminary design. The engine will be attached to a generic engine mount on the nose of the plane. This mount is readily available from model shops and used extensively for model aircraft. This arrangement also allows the engine to be easily changed or repaired if a failure occurs. If the engine were mounted inside the main fuselage, any work on the engine would require the fuselage deconstructing. A pusher was considered for the aircraft, as this is the usual arrangement for most UAVs studied. This is probably due to carrying photography equipment in the nose. However a pusher has not been chosen here as this require a specialist mounting designing and manufacturing for the engine.
To determine the best propeller size the equation below was used [7]
From this analysis, the best prop size to choose for the aircraft based on its cruise at 50m/s is a standard sized 32 x 15.
A variable pitch propeller was considered in the design, as this would allow the aircraft to operate more efficiently at a wider range of airspeed. However a fixed pitch propeller has been decided, as this does not require the complicated mechanics and systems of a variable pitch, which would be difficult to incorporate into such a small prop. Also a variable pitch propeller carries a weight penalty compared to fixed pitch due to the additional mechanics and systems. The ability of the aircraft to operate at a wide range of airspeeds is not a strong issue for this UAV as it will spend much of its flying time in cruise configuration to, from and above the target. This would make the potential benefits of a variable pitch propeller difficult to realise.
4.2 Lifting Surfaces
4.2.1 Choice of Wing Section
The wing section chosen to fulfill the requirements of the specification must have enough camber to allow sufficient lift to be generated by the complete wing during flight and manouevres, while also have sufficiently low thickness to ensure the profile drag associated with it is not too high. However, both too much camber and too little thickness will make the construction of the wings more difficult, while a wing that is too thin may also suffer in terms of structural strength and may not be able to support the weight of the aircraft in flight.
4.2.2 Aspect Ratio
The optimisation of the aspect ratio of the aircraft is an important
step in progressing further with the design of the vehicle. Changing the aspect
ratio of an aircraft has impact on a number of parameters:
Ø
The greater the aspect ratio of an
aerofoil, the lower is the induced drag at a given lift coefficient. This is due to the weaker wing tip vortices
generated by the presence of a high aspect ratio wing compared to a low aspect
ratio wing. (See fig 4)
Fig 4 –Graph showing effect of aspect ratio on the
drag coefficient
Ø
High aspect ratio
wings produce poor manoeuvrability characteristics from both aerodynamic, and
structural considerations.
Ø
The higher the
aspect ratio of the wing, the stiffer the wing must be, this can have an impact
on the mass of the structural reinforcement required and can cancel out the
drag reduction achieved by increasing the aspect ratio.
It is clear that the choice of aspect ratio for the vehicle must balance
these considerations. The
manoeuvrability of the vehicle is not an issue that has to be considered at
this stage in the process because the customer has not specified a minimum
turning radius or any manoeuvre parameters; all the vehicle has to do is circle
round a target. In order to decide on
an achievable aspect ratio the structure of the wing and its strength and weight
have to be considered against the possible benefits.
4.2.3 High Lift Devices
The use of flaps and slats increases both the aircraft drag, and the wing lift available at particular flight speeds, as shown in figure 5 below
Fig 5 – Effect of using flaps
This figure also shows that the stalling incidence is reduced with the flaps deflected, while the associated CLMAX is increased, implying that lower angles of incidence are required at take-off. Trailing edge flaps can however produce a strong nose-down pitching moment such that a larger tail surface is required to trim the aircraft. A list of the different flap systems used on aircraft today is given below, together with any specific advantages and disadvantages applicable to their use on the designed UAV, given that all result in increased wing lift:
Leading Edge Slat- relatively complicated construction.
Nose Flap- increase wing camber, relatively simple.
Plain Flap- increases wing camber, simple construction.
Plain Flap With Blowing- increases wing area, overly complex construction relying on high-pressure air.
Split Flap- increases wing area, high drag penalty.
Fowler Flap- increase both camber and wing area, slightly complicated construction.
Slotted Flaps- highly complex construction.
4.2.4 Wingtip sails
These additions to the outer edge of the wing use the spanwise
circulation of airflow around the wing to produce a forward thrust that can
reduce the drag of an aircraft significantly in cruise configuration. However the drag reduction can be outweighed
by the complexities of constructing the devices and extra structural weight
that they contribute to the airframe.
The devices also have to be optimised to the aerofoil section that they
are applied to in order for the full benefits to be seen, and the presence of
winglets can adversely affect the aircrafts drag characteristics in the out of
cruise condition. As an additional
disadvantage the presence of winglets can drastically affect the stability and
control response of an aircraft. Even
though there are many problems with the addition of these devices, they have
been used successfully on many aircraft; the Chapy Corp. Hawk-I 7H is a UAV
design used for short range missions that uses winglets, many commercial jets
also use winglets to maximise economy an example is the Boeing 747-400.
Due to the difficulties in designing a whole wing to take account of the
presence of winglets, and the complexity and structural weight that these
devices add to the airframe the option is going to be ignored, in order to
simplify the design and construction of the vehicle.
4.2.5 Lifting Surface Choice
The preliminary choice for wing section for the UAV based upon the stated requirements is the Gottingen 535 section. A profile of this section can be seen below, below which is a chart depicting the lift and drag characteristics associated with it. [8]
CD CL
1.6 0.6
Fig 6 - Gottingen 535 wing section with corresponding
lift and drag relative to incidence
As can be seen, this wing section provides favorable lift and drag characteristics, with a relatively high CLMAX, allowing good performance in take-off and landing, while the drag in cruise will not be too high. The fact that this section is also not too thin will allow the wings of the UAV to be manufactured relatively easily, and in such a way that structural strength is sufficiently high to allow for high wing loading in all manouevres.
Using the lift characteristics of the aerofoil from fig 6 the cruise CL will be approximately 0.6 as this is the zero incidence lift coefficient. Also as can be seen on the diagram, CLMAX is clearly 1.6. These values will be used in the constraints analysis (section 4.5). It has been decided not to use flaps for take-off and landing. This has been discussed further in section 4.7.
The details with respect to the drawing of this section are shown in a table in the appendices.
4.3 Camera Choice
It has been decided that a commercially available camera would be better to use in the UAV rather than a specialist design. This is because commercially available digital camcorders are now relatively cheap and are easily available from electrical retailers such as Comet and Currys. Also, they’re picture quality is not notably different to a more expensive choice.
The camera that will be used is the Sony DCRPC101 whose specification is listed below.
Fig 7 – Sony DCRPC101 Mini DV camcorder
|
Zoom |
10x optical 120x digital |
|
Image Stabiliser |
Yes |
|
Still Photo capability |
Yes |
|
Removable Memory |
Sony Memory Stick |
|
Dimensions |
55mm x 105mm x 99 mm |
|
Weight |
480g |
|
Price |
£1 099.56 (Currys) |
Source www.currys.co.uk
This camera has been chosen because of its good weight and size. Also the unit has removable memory which will mean that the UAV has quick turn around times as just memory has to be replaced rather than removing the camera. In addition to this, the camera has a facility enabling it to take both video and still images. The image stabiliser is particularly useful, as it will help to damp out the vibrations transmitted through the airframe during flight giving a clearer image. The price is also reasonable for this equipment and as can be seen it is readily available as are replacement batteries, memory and other accessories.
4.4 Fuel Fraction and Weight Estimation
Details of how this analysis was carried out can be found in the appendix (A-3).
Using the relationship between the OWE and take-off weight of current UAVs and the newly designed UAV, it is possible to plot a graph. The intersect of the two lines will give an estimate for the weights of the new UAV. This can be seen in fig 8
Fig 8 – Relationship between OWE and take off weight to
determine weights.
Reading from the graph, the estimated weights are
Ø
Maximum take off
weight – 17 Kg
Ø
Operational
Weight Empty (OWE) – 11 Kg
Ø Fuel Weight – 4.72 Kg
4.5 Constraints analysis
4.5.1 Constraints Diagram
It has become clear from the engine section that the engine choice and therefore engine power was fixed. From the analysis shown in the fuel fraction section (4.4), it can be seen that the design has a fixed power loading due to the fixed weight. Using this power loading figure fixes the position of the deign point on the constraints analysis and enables the aircraft wing area and CLs to be determined.
Further details of how the constraints diagram, shown in figure 9,
was constructed are held in the appendix A2-4. It should be noted that the
constraints analysis has been carried out using Imperial units.
Fig 9 – Constraints analysis diagram for the UAV design
4.5.2 Derived Parameters
Using the engine power (in bhp) and take off weight (in lbs) it is possible to calculate the maximum power loading.
Using this it is possible to fix a point on the constraints diagram where all flight conditions are met. The point chosen is indicated on the diagram.
The constraints analysis has shown that it will be possible for the aircraft to take off in a conventional manner provided the specified CL could be met. This is namely
CLmax Take off - 2.2
Also as the aircraft is to land by parachute, the landing constraint will be the stall speed at its approach. The required speed for parachute deployment is 25m/s and therefore the aircraft will require a CL of 1.6 at landing in order to fulfill this.
Now, taking the corresponding wing loading at the design point gives the aircraft wing area.
And based on the pre-specified aspect ratio of 10, the wingspan and the wing mean aerodynamic chord can be calculated
Therefore for the aircraft, the weight, power and wing section has been defined.
4.6 Structures
4.6.1 Fuselage Shape
The shape of an aircrafts fuselage is an
important parameter in determining the drag and the handling characteristics of
an aircraft. The power plant, camera,
fuel and the avionics package all have to be contained in the aircraft and at
this stage it will be assumed that all of these will be housed in the fuselage. The fuselage will not have to be pressurised,
because of the altitude that the aircraft will cruise at and because there is
no need to pressurise the contents of the aircraft.
Many UAV’s such as the Aerosonde have aerodynamically simple fuselage
shapes that consist of a cylinder with two end caps. These fuselages are easy to produce, and analyse from a
structural and aerodynamic point of view.
Other UAV’s such as the Global Hawk and the Predator have fuselages with
large nose cones used to house reconnaissance equipment, these set-ups allow a
slimmer fuselage main body and have lower drag than that of a cylindrical
fuselage that could house the equipment.
Due to the simplicity in analysing and producing a cylindrical fuselage,
and the fact that the camera unit chosen is no larger than the engine body, it
is sensible to stick with a conventional fuselage.
4.6.2 Fuselage Structure
The fuselage must be strong and streamlined since it must be able to withstand forces that are created in flight. These forces are bending, torsion, compression, shear and tension. All these forces have to be considered in the structure of the fuselage. The structure types to be considered are truss, monocoque and semimonocoque.
Truss
Since any member is stronger in compression or tension than in
bending, members carry end loads better than side loads. In order to achieve
this, the members are arranged in a form of truss or rigid framework. For a
truss to be rigid, it must be composed entirely of triangles. When the load on
a truss acts in one direction, every alternate member carries tension while the
other members carry compression. When the load is reversed, the members that
were carrying compression now are subjected to tension, and those, which were
carrying tension, are now under compression. The truss itself consists of a
welded tubular steel structure with longerons (horizontal members) and diagonal
braces. These features make it rigid, strong, and light. The truss is covered
with a fabric or metal cover, which reducers drag. To produce a smooth surface,
the fabric cover is put on fairing strips (thin flat strips of wood or metal).
These fairing strips run the length of the fuselage in line with the direction
of flight.
Semimonocoque
The semimonocoque is the most often used construction for modern, high performance aircraft. Semimonocoque literally means half a single shell. Here, internal braces as well as the skin itself carry the stress (see figure 1-4). The internal braces include longitudinal (lengthwise) members called stringers and vertical bulkhead. The skin of the semimonocoque structure must carry much of the fuselage's strength, which means it will be thicker in some places than at other places. In other words, it will be thicker at those points where the stress on it is the greatest.
Monocoque
Monocoque (single shell) structure is a thin walled tube or shell, which may have rings, bulkheads or members installed within. Loads can be carried effectively, particularly when the tubes are of small diameter. The stresses in the monocoque fuselage are transmitted primarily by the strength of the skin. As its diameter increases to form the internal cavity necessary for a fuselage, the weight-to-strength ratio becomes more efficient, and longitudinal stiffeners or stringers are added to it.
After some consideration, it has been decided that the semimonocoque is the best method of manufacture. One of the advantages of this method is that the bulkheads, frames and stringers aid in the strength of the fuselage. The main advantage is that it does not depend only on a few members for strength and rigidity as all structural members aid in the strength of the fuselage. This means a semimonocoque fuselage can withstand considerable damage and still remain strong enough to hold together.
4.6.3 Fuselage Access Schemes
A key requirement would be the placement of the fuel for stability
considerations. It was agreed that the fuel would be located beneath the wing;
the avionics and camera could then be placed just aft or just forward of the
fuel. The design would thus entail a detachable wing structure. The wings could
be made to fit into special grooves on top of the fuselage to allow them to be
taken off when access to the avionics equipment was required. The final
position of the avionics had not yet been decided, but a modular construction
was generally thought to be the most feasible in terms of manufacture and
practicality.
4.6.4 Wing Planform
It has already been established that the wings are not required to be
swept because of the speed that the aircraft will be cruising. The simplest form of wing to manufacture is
a rectangular shape, but this design can be optimised to give improved lift and
drag characteristics. For a given
amount of lift, the minimum trailing vortex (induced) drag is achieved when the
downwash is constant along the span. In
order to produce this downwash pattern, the desired wing planform is
elliptical. There are a number of
problems with this elliptical planform. There are inherent manufacturing
difficulties associated with producing the shape, and structurally it is hard
to reinforce to a high degree of stiffness.
The elliptical wing planform has not been employed on many aircraft
because of these difficulties, however the Spitfire is a notable example of its
application.
Structurally the ideal wing has a large chord at the root so that the
lift forces are concentrated near the fuselage and hence produce smaller
bending moments. The wing then narrows
symmetrically about the mid chord point so that the lift force decreases along the
span of the wing. This format would
allow a wing of minimum structural weight to be produced.
There is a compromise to be reached between these two extremes, and in
many cases this involves the use of the straight taper. This wing type is easy to construct, analyse
and gives good lift and drag characteristics. Virtually all commercial jets
employ this straight taper wing planform along with a number of operational
UAV’s such as the General Atomics Gnat 750 and the Lockheed Martin Darkstar.
4.6.5 Wing Structure
There were two main
ways considered for wing structure for manufacture. One was a hollow
rib-and-spar construction; the ribs would be cut from balsa wood sheet in the
chosen aerofoil shape, with a solid spruce single spar. The structure would
then be covered with solartrim.
The second option
would be to use a length of expanded polystyrene and cut it to required shape
both lengthwise and in profile.
The bending moments on
the wing are considered in detail in appendix.
4.7 Undercarriage
4.7.1 Take-off and Landing Considerations
The type of undercarriage ultimately depends on how the UAV is to be launched and recovered. There is the obvious solution of a conventional undercarriage with wheels in either a tri or bi configuration. The advantage of this kind of system is that it allows operational flexibility. The aircraft could be launched and recovered without the need for any other equipment. However the disadvantages include additional weight and drag which the UAV has to contend with. Using a retractable undercarriage will aid the drag but increase the OWE due to the additional servos and systems required. Also if there is an in-flight failure of the retractable undercarriage it will cause the aircraft to have to fly with a drag penalty or crash land.
However if the aircraft is to be launched (i.e. not take-off just under its own power) there are different solutions to an undercarriage. The possibility of using a wheeled trolley, which the aircraft detaches from on take off, would save weight and drag compared to a conventional system. Using this idea would necessitate the use of an alternative landing method, as there is no undercarriage attached to the aircraft itself. Also the aircraft could be launched from the top of a car or van from a roof rack. However most average vehicles cannot accelerate to a high enough speed in 100 metres to allow take-off.
A take off system similar to that used to launch gliders was also considered. This works on the principle that the UAV would sit on a trolley and be pulled along by a towrope in order to accelerate it to a suitable climb speed. Once the aircraft has started to ascent, the towrope is jettisoned.
In order to land, the aircraft could use skids to land on soft grass. The skids could be more streamlined than landing gear so retract ability wouldn’t be such an issue. However there would be no “natural” damping like that found in rubber tyres so the aircraft may experience unacceptable shock on touchdown and require and elaborate damping system fitting to the skids so that the avionics and camera are not damaged. A parachute recovery is also an option. This would mean that the aircraft does not need to use high lift devices on landing and simply ejects a parachute. The main problem with this method is a severe crash landing if the parachute fails. Also there is the issue of calculating the optimum position for the parachute to be attached that will allow a safe recovery of the UAV without the parachute being tangled. However in addition to a parachute, the aircraft could be fitted with a cushion, which inflates when the aircraft is about to touch down. This cushion would be similar to an airbag but would hold its pressure allowing a softer landing for the UAV and reduce the risk of damage to the systems. A far-reaching solution to the landing problem is to consider the aircraft expendable and only recover the avionics and camera package. Using this option would be based purely on cost.
Figure 6 shows some of the undercarriage designs that have been discussed above.
4.7.2 Undercarriage Choice
At this stage in the design process, the choice of landing gear is based on the engineer’s judgment as no detailed force calculation can be carried out.
The method of launching the vehicle has been decided as using a conventional take-off configuration but with the help of a towrope. This will allow the aircraft to reach a greater speed at take-off and therefore the required CL can be reduced. The take off CL will be reduced to a value that can be obtained by the clean wing alone. This eliminates the need for complicated flap systems, which cause increased drag and more importantly a severe weight penalty. This method has been chosen because of the good chance of success of the UAV being able to easily clear any buildings near to the take-off area.
In order for the aircraft to land, a parachute recovery system has been chosen with the addition of an inflatable airbag at touchdown. This method will be able to reduce shock to the aircraft structure and systems considerably compared with a conventional landing gear arrangement. Also this method allows a near vertical landing, which could be beneficial in areas where large open spaces are not available to allow a more conventional approach.
4.8 Stability and Control
4.8.1 Introduction
The stability and control of the UAV is important for it to complete its specified mission. This part of the report looks at the various aspects of stability and control and tries to define a design philosophy that will be used to design the UAV.
4.8.2 Control
For the success of the UAV, it is important that aircraft must be controllable in most foreseeable flight conditions. This ranges from straight and level flight all the way through to emergency manoeuvres. The flight control system comprises an avionics package, which signals the actuators to move the control surfaces through mechanical linkages. The avionics package has been specified by the customer, thus for the
control system we have to decide and specify the size and type
of the actuators and control surfaces, route the mechanical linkages.
Fig 12 - UAV’s Control system
4.8.3 Stability
Clearly the duty cycle on the control system is also dependent on the stability of the flight vehicle. If the aircraft had unstable flight characteristics the control systems would be constantly making corrections to keep the aircraft flying level. This has the following detrimental effects: -
Ø
Higher current
loads on the batteries. If the actuators are constantly moving these will sap
energy from the batteries. By making the aircraft stable the capacities of the
batteries to complete the mission are much less, thus reducing weight.
Ø
The processing
load on the computer will be high. If the UAV stable then it is less critical
to have highly tuned control loops, which should hopefully reduce the cost of
the final product.
Ø High maintenance, wear on bearings/moving parts.
Therefore the UAV has to be stable by configuration.
4.8.4 Configuration
4.8.4.1 Wing Position on Fuselage
The position of the wing along the length of the fuselage can be
determined by considering the desired stability characteristics of the aircraft
and will be addressed later. The
vertical positioning of the wing on the fuselage will be considered here.
Mounting wings above the centre of gravity of
an aircraft aids its roll stability. Consider an aircraft in a banked turn with
sideslip, the lower wing meets the airflow first, hence generates more lift and
a restoring moment is produced. Often
dihedral is not required on high wing mounted aircraft such as the
Islander. This means that for low wing
aircraft larger dihedral angles are required to achieve roll stability. Larger dihedral angles cause manufacturing
complexities, and can add to the structural weight of an aircraft.
The positioning of the wing on the top of the fuselage can also present
a few problems. The stowage space for
the undercarriage is lost when the wing is moved upwards, the high winged
aircraft requires longer legs on a retractable undercarriage, and these will be
heavier, and use up more space than on a low winged aircraft. Moving the wing further from the ground also
reduces the ground effect factor which is a useful phenomenon to aid the
aircraft in lifting off the ground at takeoff.
The higher wing may also interfere with a conventional tailplane
configuration by creating a downwash over it, this often means that the
tailplane has to be raised thus increasing the structural weight required, an
example of this is the McDonnell Douglas C-17A.
Interference effects at intersections between two surfaces such as the
wing and the fuselage disrupt flow patterns around the aircraft and add to the
overall drag produced. Oblique angles
cause the least interference, and these can be produced if the wing is placed
in the middle of the fuselage, however the presence of a wing spar going
through the centre of the fuselage may cause problems with space for the
payload.
4.8.4.2 Canard or Conventional Tailplane?
When Canard aircraft are trimmed for steady straight and level flight,
both the canard and the main wing will produce positive lift. This means that the overall wing area, drag
and total weight can be lower than required on a conventional aircraft. The Canard set up also allows shorter
takeoff runs as increasing the lift on the Canard raises the nose up rather
than pushing the tail down. There are
however a series of problems with using the Canard set-up; the main
disadvantage is that interference effects from the canard affect the airflow
reaching the main wing and can increase the drag produced considerably. Another problem is that in highly turbulent
conditions or violent manoeuvres the canard will stall before the main wing
because of its higher angle of attack, if the wing stalls at the same time then
the aircraft may be impossible to recover. [11]
4.8.4.3 Vee-Tail or Conventional tailplane?
Many UAV’s in current operation employ the use of Vee-tails, a tailplane
arrangement with two surfaces rather than the conventional three. These have a number of advantages:
Ø
Drag is reduced
on the aircraft in cruise because there is lower wetted surface area.
Ø
Weight reductions
are possible because only two surfaces are used.
Ø
Perhaps not so
important is the fact that this arrangement can avoid right angles that cause
large radar reflections. This is the
reason that the Vee-tail arrangement was used on the F117.
Some UAV’s that have benefited from the Vee-tail set-up are the General
Atomics Predator and the Prowler both are currently in use by the US Air force.
However the avionics required to control the aircraft become more complicated, and during manoeuvres the drag off the aircraft can be increased above the level of the conventional set-up.
There are simple steps to take that would mean the configuration of the aircraft is stable. This can be split into two main sections, Longitudinal Stability and Lateral Stability.
4.8.4.4 Longitudinal Stability
The key aspect is the CG position must be in front of the neutral point. Therefore the control of the CG is important, if the CG moves to far during the mission there could be major problems associated with both stability and control. If the static margin was too great then there will be large loads required by the tail plane to control the UAV but if the static margin is to small or even negative then UAV will be unstable. Thus a range of static margins is required. Our UAV will have maximum static margin of 25%MAC and a minimum static margin of 5%MAC.
To minimise drag in the cruise it is important that the tail is unloaded.
4.8.4.5 Lateral Stability
There are various configurations that lend themselves to being laterally stable and directionally stable. These include a high wing position, dihedral and sweep back. These all have an effect on roll stability. The UAV has a high wing position, which already lends itself to roll stability. Dihedral will probably have to be incorporated into the design but it does complicate the wing structure, so it will be avoided initially. Sweep back further complicates the structure and will be avoided for this design as previously mentioned.
For yaw stability it is important to have a vertical stabiliser on a suitably long tail arm, which is also beneficial for lateral stability.
4.8.4.6 Control Surfaces Configuration
Control surfaces are required to change the attitude of the aircraft. These are usually accomplished by having a rudder for yaw control, elevators for pitch control and ailerons for roll control.
4.8.4.7 Ailerons
Ailerons are used to roll the aircraft. Some model aircraft do not employ ailerons for turning the aircraft. They rely solely on the wing dihedral to accomplish this. In practice this leads to a very sloppy turn, as the aircraft has to be put into a sideslip first. This method requires relatively large amounts of dihedral to accomplish. Even though the design of the wing is less complicated and would probably result in a lighter aircraft if ailerons were to be dispensed with, this would be at expense of versatility. As such ailerons will be included on the UAV.
Positioning of the ailerons is quite important. Aerodynamically they will have more effect if they are as far outboard on the wing as possible but this imparts large bending moments on the wing structure. On very high aspect ratio wings, problems can arise due to aileron drag and control reversal. Clearly the aim is to position the ailerons as far out as the structure will allow.
4.8.4.8 Tail plane configuration
There are many types of configuration that can be employed on UAV’s pod-boom class of fuselage. Below is a brief look at the types of configuration that can be employed.
Ø
Conventional
Tails: The single fin and stabiliser is by far the most common type of tail.
Problems can be encountered by tail plane being positioned in the wake of the
wing this would result in a larger fin being required. Another problem
associated with a conventional type tailplane is masking of the rudder when the
aircraft is a spin, this can be overcome by careful positioning of the two
surfaces. The Conventional tail has many advantages; one of the main ones is
being structurally simple.
Ø
“T” Tails: Common
on sailplanes, these tails have some aerodynamic advantages over a conventional
tail. One of the main advantages is that tailplane is usually out of the wake
of the aircraft wing and thus can be smaller. It tends to have a higher
efficiency due to the reduction of vortex drag off the fin. Problems with this
design include the structural rigidity of the high tailplane, complexity of the
control linkages and poor deep stall characteristics. Despite these problems this is the simplest and the most likely
tail design that will be taken to the detailed design phase.
Ø
Multiple Vertical
Fins: These are often found on large multi-engined aircraft and are usually
employed if large rolling moments associated with a tall rudder is undesirable.
There is some aerodynamic advantage if the fins are positioned at end of the
tailplane. The major disadvantage is the complexity of the control linkages.
This design will not be employed on our UAV
Ø
“V” Tail: Are one
of the least popular design of tails on full size aircraft but they are popular
on model aircraft and UAV’s. They have a number of advantages over conventional
tails. These include less weight and drag due to smaller overall area, better
spin recovery and suffer less from interference from the wing. The main
disadvantage is the mixing required for the control surfaces but this can be
easily accomplished electronically by the avionics.
Further analysis is required to decide on whether a conventional tail configuration or V tail will be used for our UAV.
4.8.5 Actuator Sizing
The actuators that will be utilised in this UAV are
commercially available RC servos. Data was obtained from Futaba website
[www.futaba-rc.com] (Part of Hobbico Inc) which gives basic dimensions, weights
and torque ratings. This data has been included in appendix 2-6. Their servos
range from a torque range 0.12Nm all the way to 2Nm. From this data some
initial estimates on the maximum size of the control surfaces can be done.
Fig 13 - Control Forces
Equating work done by the moments gives a relationship between the output torque from the servo and the hinge moment. This is: -
Where G is the gearing ratio. Typical gearing ratios are
approximately 0.6.
The hinge moment
is given by the equation below. [12]
Where Ss is the are of the control surface, cs is the cord of the control surface and Ch is the hinge moment coefficient. The typical rate of change of Ch with respect to d is 0.7 rad-1 [13]. Thus for deflection of 15 deg this gives a Ch =0.183.
For normal cruise flight conditions V=50m/s, r= 0.91kg/m3 and a typical torque output of servo 0.55Nm. This gives a Sscs of 1.1x10-3m3. For a control flap of 5cm cord the area of the flap is 0.022m2, which thus gives a length of 0.44m. From this brief study it shows that a COT servo is clearly capable of moving the control surface which will reduce cost and complexity of the design. Clearly a more in depth method will have to be used to estimate the hinge moments.
4.8.6 Design Criteria
With regards to stability and control requirements for the UAV the following design criteria will be used.
Ø
The UAV should be
inherently stable.
Ø
This implies that
CG position has to be carefully controlled. The range of the static margin will
be between 5%MAC and 25%MAC
Ø
The wings will
probably have dihedral.
Ø
Ailerons will be
included and will be positioned as far out board as possible.
Ø
The tailplane
will either be a conventional T-tail or V tail.
Ø Commercially available actuators will be used.
4.9 Materials and Manufacturing
4.9.1 Aims of materials selection
Having considered the flight envelope using the given data such as
range, mission duration and specification as well as payload, the group then
set about loosely defining what solution would best fit the given parameters.
Choice of materials to use was centred on ensuring that weight was minimised
while structural strength was optimised, a perennial problem facing aircraft
design on any scale. Manufacturing methods were also to be kept as simple as
possible.
4.9.2 Choice of materials
After looking at several possibilities, wood construction was decided on
as the easiest and most cost-effective method. The following materials were
considered:
Balsawood
This was the material considered first. It is light and relatively easy to manufacture. It is widely used in model aircraft. An investigation into possible reasons for this led to acquisition of wood data which yielded a low specific gravity (0.16) and yet relatively high compressive strength (parallel to grain with 12% moisture content – 14,900kPa) [14]
Spruce
Another wood commonly used, spruce is about 2.75 times as dense as balsa
(under the same moisture conditions) [14], but its compressive
strength is proportionately higher as well.
It was decided upon because it can give to the wing structure the
strength required, since the wings will have a high aspect ratio and thus will
need to be stiff enough to withstand aerodynamic loads as well as fuel and
payload weight.
PVC Plastic tubing
After considering the use of expanded polystyrene, it was decided that a
plastic tube of 200mm diameter be used instead of a block of expanded
polystyrene to manufacture the fuselage. This would be easier to manufacture
and analyse structurally. PVC plastic was thought to be suitable, but this had
not been yet agreed on as the final material.
Solartrim
This thin gives a
smooth surface and good aesthetic finish. It is thin, lightweight but tough,
easy to cut and self-adhesive. However, light colours tend to be translucent.
This can be detrimental to the joint integrity (especially for the adhesive) at
high altitudes in clear weather.
Information on
these materials is included in Appendix A2-7.
4.9.3 Manufacturing methods
The most challenging aspect of the manufacture process would be the
wing. These must be light but at the same time be able to take the structural
as well as aerodynamic loads. A simple wing box model was to be employed
initially for preliminary structural calculations. The actual wing
construction, however, would be based on a ‘tried-and-tested’ use of ribs with
reinforcing spars. It is most likely, at this stage that the ribs will be cut
in the shape of the aerofoil, and then the spars will run the length of the wing.
This should give sufficient structural strength while minimising weight. The
leading edge shape can be achieved by using a commercially available balsa wood
of the required length and cross-section. The entire framework will be covered
in Solartrim to give a smooth surface finish. It also gives a close-fitting
overall structure that helps keep the internal wing structure intact.
4.9.4 Method of attachment: Wing to body
After careful considerations, it was decided that the fuel would need to
be stored just under the main wing for stability considerations. This posed a
design problem because another key design point was to be transportability. It
was though that this could effectively be achieved by constructing the fuselage
so that the wing would slide onto the top of the fuselage tube, possibly by
means of grooves. At the same time, the entire front section of the aircraft,
i.e. the engine, forward fuselage and wings, should be detachable from the rear
portion that includes the avionics box and digital camera. This arrangement was
seen to be the most feasible to allow access to both the fuel and the avionics.
A schematic of this is shown below:
5 Chosen Preliminary Design
This section describes
in brief the optimisation process for each of the components of the
aircraft. It details the reasons why
the particular configurations, components and structures have been chosen to
take on to the detailed design phase.
Propulsion
The chosen engine is the OS Engines FS-91, this was chosen because it provides the optimum power for weight. The propeller is an off the shelf 32x16 fixed pitch propeller that has been determined to perform the best coupled to this engine. Also the combination is the only choice where the required amount of technical data is available. The constraints analysis has shown that the power provided by the engine will enable a realistic weight for the UAV. Also the engine and prop are readily available from many model shops should a replacement be needed and have good part availability. Being an off the shelf item means that their price is relatively cheap when compared to a special design made as a one off. The fuel required by the engine is also a standard model aviation fuel so is readily available. The engine has been tested by a third party for this project and has proved reliable in these tests.
Wing Choice
The aerofoil that has been chosen for use on the main wing of the aircraft is the Gottingen 535. This aerofoil was chosen because of its favourable lift to drag characteristics, and its high thickness to chord ratio that allows a strong structure to be built inside of the wing. The aspect ratio of 10 has been chosen to allow a compromise between structural and aerodynamic efficiencies, and a straight taper of the trailing edge aids both the aerodynamic and the structural aspects of the wing.
Launch and recovery
The UAV take-off requirements have been tailored to be simple and re-usable
In order to take-off, the aircraft will require a custom built trolley. However this will be a simple design using many commercially available parts. Towrope attachment will be such that the rope can be re-used for subsequent launches. Despite the need for a take off CL of 2.2 according to the constraints analysis, the towrope will be used to accelerate the aircraft to a higher speed, allowing take off to take place at a much lower lift coefficient.
The landing of the UAV will require a CO2 capsule to be carried aboard the aircraft. Once the UAV is close enough to the ground, this will be detonated causing the airbag to inflate. The airbag will be obtained from a supplier of car airbags that can easily be replaced after each landing. These airbags will not be re-used as there is a chance that the material may become more likely to pop upon charge detonation the more times it is used. The parachute however will be re-used every time on the aircraft and be easily pack able so that turn-around times are kept to a minimum. The use of parachute modules which can be replaced immediately and spend parachutes be packed later will be investigated as will the use of off-shelf parachutes, depending on the required size.
Stability and control
The UAV is required to be stable by configuration to reduce the duty cycle of the control system. This will be accomplished by a high wing position, dihedral if required, and careful C of G management to maintain the static margin within its range. The actuators will be COTS servos used by model aircraft. This means that they are readily available from model shops and cheap. The tail will be a conventional T-tail configuration to ease both analysis, and manufacture of the airframe. The UAV will also have standard ailerons positioned as far out on the wing as feasibly possible.
Structures, Materials and Manufacture
The fuselage needs to be streamlined and
strong to withstand forces created in flight. The choice of materials and the
type of structure will be important to make sure that the fuselage is strong
enough for the whole flight including the takeoff and landing phases. Weight
reduction and simplicity of construction will be the key drivers of the choice
of materials and manufacture. The fuselage will consist of PVC plastic tube cut
to the required length. This would be relatively easy to manufacture as
compared to, say, a solid block of expanded polystyrene cut to the required
shape. The wings would predominantly feature balsa wood for reasons of weight
reduction. A single spar spruce construction would be employed whereby the spar
would run the length of the wing, intersecting with balsa ribs cut in the
aerofoil shape. The entire construction would then be covered in solartrim. All
the materials selected are easily obtainable off the shelf and relatively
simple to manufacture.
APPENDICES:
A1 Chief Engineers Report
This is the first major group project that any of the group had participated in, and for this reason nobody really knew what to expect. The group meetings were a major factor in the success of this report, and because of the set layout of the meetings, and their regular pattern, they formed not only a good way to share ideas and help each other, but also allowed discussion on the progress of the work, and in an indirect way put pressure on everybody to complete what was required of them.
Perhaps the most difficult part of this whole task was actually defining a starting point to allow further work to be carried out, the whole group participated in this phase of the work in order to define a number of parameters that were estimates of the aircrafts characteristics, these were taken from both current UAV data and engineering based reasoning. A major part of the work for this project was the choosing of the powerplant for the aircraft, in order for the aircraft to perform its task. This became a problem when the lack of any data on specific fuel consumptions for commercially available engines was realised. A lot of time was spent on a solution to this problem, with a rough estimate given by Christian Harris eventually being used. The production of a constraints analysis also posed a few problems for the group, a number of members worked together to produce a series of solutions however a final solution was discovered by Russell Hankey the propulsion engineer; it is this that is included in the report. The majority of the remainder of the work however has been research into configurations of the aircraft, and decisions based on this research regarding defining features of the aircraft such as aspect ratio, and fuselage layout.
Overall the team has worked very well as a unit, with attendance at meetings being excellent, team morale has been high, and deadlines have been met easily by following as closely as possible the project plan that has been kept up to date on a regular basis.
The allocation of marks within the group has been discussed extensively within the group with each member arguing their case; the results are shown below
|
|
Tim Myall |
Andrew Walker |
Keith Watson |
Russell Hankey |
Andy Chan |
Mark Kathurima |
|
Attendance |
|
|
|
|
|
|
|
Delivery on allotted tasks |
|
|
|
|
|
|
|
Contribution of ideas |
|
|
|
|
|
|
|
Contribution to the report |
|
|
|
|
|
|
|
Rating |
|
|
|
|
|
|
A2 Technical Appendix
A2-1 Propulsion
Engine Performance Data:
|
|
OS Engines 1.40 RX |
OS Engines FS-91 II |
|
Picture |
|
|
|
Displacement |
23 cc |
14.95 cc |
|
Bore |
32 mm |
27.7 mm |
|
Stroke |
28.6 mm |
24.8 mm |
|
RPM Range |
2 000 – 9 000 |
2 000 – 12 000 |
|
Output |
3.5 bhp @ 9000 rpm |
1.6 bhp @ 11 000 rpm |
|
Weight |
0.830 Kg |
0.678 Kg |
|
SFC |
Unknown |
3.055E-07 Kg/s/W |
A2-2 Lifting Surfaces
Co-ordinates for GOE 535 AIRFOIL
1.000000 .000000.987580 .003927.967555 .010127.945890 .016654.922853 .023386.898832 .030174.874115 .036911.848874 .043539.823231 .050027.797245 .056360.770904 .062549.744190 .068613.717147 .074563.689892 .080395.662602 .086077.635459 .091552.608617 .096770.582196 .101693.556263 .106288.530796 .110530.505701 .114413.480830 .117936.456065 .121114.431435 .123950.407117 .126433.383397 .128547.360537 .130240.338635 .131466.317658 .132205.297502 .132454.278041 .132215.259202 .131501.240964 .130328.223343 .128715.206376 .126688.190104 .124280.174580 .121534.159881 .118509.146083 .115270.133235 .111866.121340 .108302.110341 .104584.100161 .100715.090715 .096706.081932 .092604.073758 .088460.066154 .084316.059090 .080213
|
.052545 .076192.046499 .072282.040943 .068478.035863 .064748.031230 .061065 .027009 .057406.023165 .053753.019680 .050101.016548 .046457.013761 .042824.011311 .039205.009175 .035599.007322 .032005.005725 .028421.004362 .024846.003213 .021279.002261 .017719.001491 .014166.000891 .010617.000449 .007074.000155 .003535.000061 .001776.000000 .000000.000985 -.003034.002139 -.005900.003482 -.008637.005020 -.011233.006765 -.013688.008732 -.016000.010940 -.018166.013414 -.020181.016166 -.022046.019187 -.023771.022460 -.025367.025961 -.026846.029670 -.028214.033601 -.029466.037773 -.030596.042209 -.031597.046939 -.032466.052003 -.033196.057456 -.033792.063374 -.034266.069848 -.034645.076973 -.034978.084838 -.035312.093527 -.035640.103128 -.035940 |
.113730 -.036183.125419 -.036342.138258 -.036378.152328 -.036245.167807 -.035897 .184924 -.035320.203745 -.034535.223900 -.033577.244671 -.032420.265443 -.031032.285957 -.029377.306333 -.027395.326992 -.025031.348408 -.022277.370987 -.019135.394988 -.015639.420307 -.011901.446362 -.008057.472357 -.004247.497569 -.000599.521529 .002793.544131 .005826.565578 .008450.586208 .010649.606421 .012414.626645 .013766.647241 .014770.668461 .015475.690433 .015925.713091 .016163.736150 .016202.759249 .016031.782049 .015651.804273 .015073.825730 .014295.846333 .013276.866085 .011983.885046 .010389.903304 .008474.920971 .006318.938193 .004210.955128 .002426.971928 .001122.988729 .0002921.000000 .000000
|
A2-3 Weight Estimation
Symbols and units used for calculating fuel fraction
|
Symbol |
Description |
Unit |
|
R |
Aircraft Range |
M |
|
h |
Propeller Efficiency Factor |
- |
|
C’ |
Specific Fuel Consumption |
Kg/s/W |
|
G |
Gravitational Acceleration |
M/s2 |
|
m |
Aircraft Mass |
Kg |
|
L |
Lift |
N |
|
D |
Drag |
N |
|
t |
Loiter Time |
Sec |
|
V |
Cruise Speed |
M/s |
|
CD |
Drag Coefficient |
- |
|
CL |
Lift Coefficient |
- |
Fuel Fraction formula:
Fuel fractions for stages engine start-up, take off, climb and land taken from standardised values for a micro light.
To find fuel fraction for cruise, use Brequet Range Equation for piston prop aircraft
Assume constant incidence (lift will be constant)
To find fuel fraction for aircraft to loiter over target, divide above range equation by velocity.
Overall fuel fraction for flight plan
The weight of fuel in terms of aircraft take off weight is given by
(Where Mff is the overall fuel
fraction)
Fuel Fraction Spreadsheet:
The spreadsheet over the next page shows the calculation of each stage of the fuel fraction along with the inputs used. This spreadsheet uses a broad estimate of SFC for the engine as 320cc of fuel last 12 minutes.
The fuel fraction is shown to be
SPREADSHEET TO GO IN HERE
Verification of Weights:
Using the UAVs detailed in the following table, the line in figure 3 was plotted using a simple excel spreadsheet and applying a linear trend line.
|
UAV Model |
OWE (Kg) |
MTO (Kg) |
|
Dara Aviation D-1 |
25 |
38 |
|
Pakistan MkI |
16 |
30 |
|
Aerosonde |
13.97 |
18.96 |
|
ALO |
14 |
20 |
|
AVE |
36.4 |
63.6 |
|
DP-4 |
35 |
64 |
|
Fox AT-2 |
65 |
125 |
Source www.uavcenter.com, www.uavforum.com and www.uvonline.com.
The equation of the trend line was found to be
WTO= 2.308WOWE - 8.21
To find the corresponding for the new UAV, the simple relationship below is used, based on a payload of 1.5 Kg (avionics = 1 Kg, camera = 0.5 Kg)
Plotting the two lines whose equations are given above gives the graph shown in fig 3 and therefore the weight can be estimated by reading off the axis at the point of intersection of the two lines.
The fuel weight is calculated using the take off weight just found. Remembering
A2-4 Constraints Analysis
Symbols and units used for Constraints Analysis
|
Symbol |
Description |
Unit |
|
IP |
Aircraft Power Index |
(hp/ft^2)^0.33 |
|
W |
Aircraft Weight |
Lbs |
|
S |
Wing Area |
Ft2 |
|
P |
Engine Power |
Hp |
|
s |
Density Ratio |
- |
|
CD |
Drag Coefficient |
- |
|
CL |
Lift Coefficient |
- |
|
CD0 |
Drag coefficient due to skin friction |
- |
|
KCL2 |
Drag Coefficient due to lift |
- |
|
L/D |
Lift to Drag Ratio |
- |
|
q |
Aircraft Climb Angle |
Rad |
|
h |
Propeller Efficiency Factor |
- |
|
Vs |
Stall Speed |
Mph |
|
r |
Density |
Slugs/ft3 |
In order to construct a constraints analysis for the UAV, a similar method to that used for large commercial jets was employed. The aircraft has to be able to operate at its 4 main conditions. These are take off, climb, cruise and landing. Formulas have to be determined which will give the relationship between the power loading (W/P) and the wing loading (W/S) at each flight condition. These relationships are then plotted graphically and a design point chosen.
This constraints analysis has been conducted wholly in Imperial units. Most of the information in this section is based on the definitions given in reference 3.
Cruise
As defined previously, the aircraft will cruise at 50m/s at 3000m. The first stage is to determine the aircraft’s power index Ip. This parameter is proportional to the aircraft’s cruise speed and by using a figure in reference 3 which shows the relationship graphically, The power index has been found to be 0.7 (hp/ft^2)^0.33.
The power index can then be substituted into the equation given below to give a straight-line relationship between the power loading and wing loading in cruise.
Where s is the
density ratio at cruise height.
The constraint is met above the cruise line.
Climb
In order to determine the climb constraint, some parameters have to be defined first, namely the climb angle, propeller efficiency, density ratio and the drag co-efficient.
At this stage it is acceptable to take the drag co-efficient as
The climb angle has been set at 10°, which will allow the UAV to clear most buildings, and obstructions on take off.
The propeller efficiency has been taken as 0.75 as this is deemed to be an accurate estimate at this stage in the design process. The efficiency is not as high as would be expected as in cruise due to the slower speed.
The density ratio has been calculated based on a cruise height of 3000m.
The relationship between wing loading and power loading in the climb phase is given below.
By using this formula and the parameters detailed above, the climb constraint can be calculated for a range of lift coefficient. The constraint is met below the climb line.
Take Off
For take-off, a simple quadratic equation relates the take off ground run (in feet) to a take off parameter required by FAR 23. Although the UAV is not bound by FAR requirements, it has been decided that it would be beneficial not to take-off too close to stall speed as this may cause control problems. Based on 100m ground run (=228ft)
Solving this equation for T and taking the positive root. T=60.267
Note s=1 at take off as taking off at sea level.
Using this relationship, the take-off constraint can be plotted at a range of CL. The constraint is met below the take off line.
Landing
The landing case has been constrained based on a 25m/s stall speed where the parachute will be deployed. This constraint has been determined by using the definition of the lift coefficient and rearranging to give.
Note that density is calculated at sea level with unit slugs/ft3. The wing loading for various CL can be plotted and the constraint is met when to the left of the line.
SPREADSHEET TO GO IN HERE
A2-5 Undercarriage
PUT RUSSELL’S DRAWINGS IN HERE
A2-6
Stability
Analogue Servos
|
Description |
Stock # |
Dimensions |
Weight |
Torque |
Speed |
Case Set |
Gear Set |
Application |
|
S136G
Compact Retract |
FUTM0670 |
0.87 x 1.75 x 1.00 |
1.48 |
76.4 |
0.50 |
FCS36G |
FGS36G |
air, sp |
|
Best retract servo; metal
gears; gold plated connectors; water resistant; 5-pole motor |
||||||||
|
S148
Standard Precision |
FUTM0710 |
0.77 x 1.59 x 1.41 |
1.50 |
42.0 |
0.22 |
FCS48 |
FGS48 |
air, sail, heli, off, on,
gas, boat, sb |
|
Most economical servo for
any car, boat or plane; 3-pole motor |
||||||||
|
S3001
Precision BB |
FUTM0029 |
0.77 x 1.59 x 1.41 |
1.60 |
42.0 |
0.22 |
FCS3001 |
FGS48 |
air, sail, heli, off, on,
gas, boat, sb |
|
Natural upgrade of popular
S148 or S3003; 3-pole motor |
||||||||
|
S3002 Mini BB |
FUTM0030 |
0.62 x 1.21 x 1.18 |
1.20 |
46.0 |
0.16 |
FCS3002 |
FGS3002 |
air, sail, on, sp |
|
Metal gears; 5-pole motor |
||||||||
|
S3003
Standard |
FUTM0031 |
0.78 x 1.59 x 1.42 |
1.31 |
44.4 |
0.23 |
FCS3003 |
FGS3003 |
air, sail, off, on, gas,
boat, sb |
|
The standard in low cost
servos for all uses; 3-pole motor; not for heli use
|
||||||||
|
S3004
Light Weight BB |
FUTM0004 |
0.78 x 1.59 x 1.42 |
1.30 |
44.4 |
0.23 |
|
|
air, sail, off, on, gas,
boat, sb |
|
Single ball bearing on
output shaft; precise tight fit throughout the geartrain; light weight; not for heli use
|
||||||||