Feasibility of Parachute Recovery Systems for Small UAVs

Transcription

1 Feasibility of Parachute Recovery Systems for Small UAVs Kirk Graham Stewart Cartwright University of New South Wales Australian Defence Force Academy School of Aerospace, Civil and Mechanical Engineering Canberra, ACT 2600, Australia Damage to model UAV research aircraft, whether it is structural or systems damage can be costly. To combat this problem this project focuses on determining the mechanical feasibility of a parachute recovery system for small model UAV research aircraft. The economic feasibility of such a system will also be briefly considered. The findings of this project indicate that a PRS is capable of recovering these aircraft within an appropriate altitude. The low cost of implementation for a PRS also make it financially cost effective. This project extends from the initial task of designing and constructing a model aircraft for competition in the SAE Aero Design Competition. The design criteria for this competition were to develop a remote control aircraft with set engine and geometric limitations to lift maximum payload up to a takeoff weight of 25 kg. NOMENCLATURE m = mass [kg] g = gravity [ms -2 ] C d = drag coefficient S = surface area [m 2 ] D o = nominal parachute diameter [m] D p = inflated parachute diameter [m] a = semi-major axis length [m] b = semi-minor axis length [m] V = velocity [ms -1 ] F d x t f n C x X 1 A dv/dt dθ/dt θ β PRS UAV SAE ADFA = density [kg/m 3 ] = drag force [N] = parachute filling distance [m] = parachute filling time [s] = canopy fill constant = opening force coefficient = opening force reduction factor = ballistic parameter = acceleration = rate of change of attitude = attitude angle = power of filling function = Parachute Recovery System = Unmanned Aero Vehicle = Society of Automotive Engineers = Australian Defence Force Academy 1

2 CONTENTS Nomenclature... 1 I. Introduction... 3 A. Aim... 3 B. Scope... 3 C. Background & Justifications... 3 D. Methodology... 4 E. Assumptions and Limitations... 4 F. Management... 4 II. Literature Review... 5 A. Canopy Shape... 5 B. Parachute Activation... 5 C. Inflation Characteristics... 6 D. Parachute Filling Distance... 7 E. Attachment Considerations... 7 F. Summary... 7 III. Concept Selection... 8 A. Introduction... 8 B. Concept 1 Uncontrolled Deployment... 9 C. Concept 2 Ballistic Deployment... 9 D. Concept 3 Spring Release E. Concept 4 Pilot Chute Deployment F. Conclusion IV. System Design A. Parachute System Configuration B. Canopy Sizing C. Pilot Chute Design D. Initiation Method E. Containment F. Deployment Bag G. Canopy Inflation Predictions H. Summary V. Testing and Analysis A. Introduction B. Method C. Results and Validation D. Analysis E. Summary VI. Cost Benefit Analysis A. Purpose B. Scope C. Proposed System D. Costs E. Benefits F. Summary VII. Conclusion A. Concluding Remarks B. Future Direction C. Recommendations VIII. Acknowledgements References

3 I. Introduction A. Aim The aim of this project is to determine if a Parachute Recovery System is a mechanically and economically feasible method of safely recovering the ADFA SAE aero UAV from flight failure. An aircraft parachute recovery system (PRS) is a procedure that relies on the deployment of a parachute to aerodynamically decelerate an aircraft allowing for a safer touchdown (Knacke, 1992). B. Scope This project ultimately began with group design and construction of the ADFA SAE Aero UAV. Once completed individual projects were selected for all group members. The scope of this individual project begins with the examination of existing PRSs, and a review of the principles involved with PRS design. The development of initial concepts will follow, then a focus on system design. More detailed design of the chosen concept will be pursued with the aid of experimental results, gathered through testing. The testing results will be aimed at determining the mechanical feasibility of the PRS. The ability to safely recover UAVs in the event of flight failure is of significant importance. Research will also go into the benefit of a PRS for such small research UAV platforms. A determination of how beneficial such a system will be to small UAV research will try to be reached by carrying out a brief cost time benefit analysis. The revised tasks within the scope of this thesis are listed below; Obtain design details and develop the design requirements analysis. (Group) Design and construct the aircraft. (Group) Flight test the aircraft. (Group) Review existing parachute recovery systems. (Individual) Design a system suitable for this aircraft. (Individual) Test validity of design. (Individual) Determine feasibility of the design. (Individual) The ultimate goal of this project is to determine if a PRS can safely recover this small balsa constructed UAV within an appropriate altitude. Because radio controlled aircraft operate at low altitudes it is unknown if a parachute can open quickly enough to save the aircraft. Furthermore, is it worth the school of ACME investing in a system that may turn out to be more expensive than the UAV itself? C. Background & Justifications The initial task leading up to this project was to design, build and test a UAV to compete in the 2008 SAE Aero East Competition. This task was predominantly undertaken by three students, at the Australian Defence Force Academy Christopher Kourloufas, Gerad Markham and Kirk Cartwright. The design report for this vehicle is attached in ANNEX A. Each student was entrusted with different aspects of the UAV design, and once completed; individual project research was undertaken in different areas pertaining to the aforementioned UAV. 3 Figure 1 - ADFA SAE Aero UAV

4 The design and construction of the ADFA SAE aero UAV was completed in early march 2008, after several months of work. As demonstrated in Fig. 1 the UAV consisted of a canard design, powered by a single piston engine, with conventional tricycle undercarriage. After the construction of the UAV, there was reluctance for the flight testing to begin for fear of damage to the aircraft as a result of flight failure. This reluctance stemmed from the need for an intact aircraft fuselage to be tested within the scope of another student s thesis project. It was suggested that a PRS could be used to reduce the risk of total aircraft damage if flight failure were to occur. The justification for this individual project stems from this need to reduce the risk of total aircraft damage during flight testing. Furthermore, the University of New South Wales at the Australian Defence Force Academy already owns and operates several small UAV research aircraft similar in size and weight to the ADFA SAE Aero UAV. Further justification for this project is a direct result of damage to one of these aircraft, during a flight test. The damaged aircraft was a small research UAV with a gross weight of no more than 25kg, which crashed due to radio controlled interference. Damage caused to the aircraft was considerably costly, and the inability to continue testing caused delays in research. D. Methodology The overall goal of this project is to determine the mechanical and economic feasibility of a PRS for the ADFA SAE Aero UAV. The methodology used to reach this goal is to; 1. Carry out preliminary design of a suitable PRS for the ADFA SAE Aero UAV. 2. Develop a method of testing the effectiveness of this design 3. Test the design. 4. Analyse the test results to determine the mechanical feasibility of the design. 5. Undergo a brief cost benefit analysis as a supporting document to determine the economic feasibility. This methodology outlines the basic structure of this project. E. Assumptions and Limitations To ensure that the basic requirements for the PRS remain the same throughout the design process, it will be assumed that the ADFA SAE Aero UAV is the dominant aircraft that the system is designed for. This will provide a constant base aircraft to design from. The adaptability of the PRS to different platforms will still be considered during the design process. Due to the scope of this project, coupled with the time constraints associated with its completion, the design will be limited in its depth and detail. The project will be more focused on the overall system design, with an underlying goal of determining the feasibility of a PRS. The design however, must be sufficiently in depth to determine both the mechanical and economic feasibility of a PRS for use on small research UAVs. F. Management Appropriate management documentation has been attached in ANNEX G 4

5 II. Literature Review PRSs are not a new concept, and there has been significant research undertaken into several of the more complex problems associated with their design. While keeping the aims of the project in mind, this literature review summarizes some of the research that has occurred in this area. Parachute terminology is used quite extensively throughout this review, so to aid understanding; a diagram of parachute parts is attached in ANNEX B. A. Canopy Shape Principally, UAV PRSs use three different canopy shapes; cruciform or cross-type canopies, hemispherical canopies and parafoils (Wyllie, 2001). Parafoils are gliding parachutes, designed to be steerable, allowing for a small level of navigation after deployment. Their internal cell structure is ram-air inflated which forces the parafoil into a classic airfoil shape. To operate as intended parafoils need to stay inflated and are therefore constructed out of a low porosity fabric (Wyllie, 2001). This causes an increase in the opening shock forces experienced during inflation and a complex reefing mechanism is generally required to reduce these loads. Deployment is further complicated by the need to protect the control line servos from these opening loads (Wyllie, 2001). A follow on effect of inputting systems to reduce parachute opening loads causes a much slower deployment speed and therefore greater height loss during deployment. Cruciform canopies are the simplest of the three canopy shapes consisting of two pieces of rectangular cloth overlaid and sewn together as shown in Fig. 1. These canopies have the smallest drag coefficients, and lower opening forces. The small opening Figure 2 - Cruciform Canopy Configuration (Wyllie, 2001). forces, attributed to gentler parachute inflation, means that the falling body losses more height before full inflation is attained. Cruciform canopies produce lower oscillation than hemispherical canopies, which is one of the reasons they have been researched for use in precision airdrop systems are used as drogue stabilizing parachute (Keith Stein, 2001). Hemispherical canopies have high drag and opening force coefficients (Wyllie, 2001), affording them the advantage of better reliability on opening. Hemispherical non-steerable parachutes are used for aircraft recovery because their simplicity enhances their reliability. Simplicity pertains not only to the parachutes reliability but also to ease of construction and packing, an imperative requirement for this project. B. Parachute Activation Parachute activation refers to the method of deployment prior to inflation. It can be assumed that a key factor of parachute deployment systems is reliability. Three principal deployment methods are forced ejection systems, drogue or pilot parachute systems, and rocket extraction systems (Huckins, 1970). Forced ejections systems are common extraction methods due to their simplicity. The mortar, catapult, and pressure bellows are examples of mechanisms designed to produce a forced ejection of the packed parachute (Huckins, 1970). These systems tend to be heavy and they also produce high reaction loads, which is important when considering the platform in which the system will fire. Parachute deployment using a drogue or pilot parachute has numerous advantages. The system is quite flexible since the parachute extraction force is applied continuously over the entire deployment sequence, and the system is also lighter. This system relies on aerodynamic force to extract the main parachute, thus problems may arise due to pilot chute interference with the wake turbulence of the descending body (known in skydiving as hesitation ). Used in tandem, individual extraction systems increase their effectiveness as demonstrated by the Gemini Spacecraft, which used a drogue gun to launch a drogue parachute to stabilize the re-entry vehicle, until a height at which the pilot chute was extracted, pulling out the main chute (Vincze, 1966). 5

6 A rocket extraction system for parachute deployment has all the advantages of a drogue parachute system, but does have a slight weight penalty. Furthermore, the rocket extraction system produces very light reaction loads, and is only slightly dependent on the characteristics of the vehicle wake (Huckins, 1970). The rocket extraction does however increase the risk of damaging the parachute fabric on extraction, and has the added complexities of dealing with pyrotechnics. C. Inflation Characteristics In view of the fact that parachute inflation is a very complex and unsteady process, it is well known that parachute theory is a difficult problem in the aerodynamic field (Calvin, 1984). PRSs in UAVs require parachute inflation to be reliable and quick, to ensure minimum loss of height during opening. In manned PRSs, such as the Ballistic Recovery System (BRS) used in the Cirrus SR20 (Ballistic Recovery Systems Inc.), complex dis-reefing mechanisms are put it place to slow the inflation process and reduce shock forces caused by the opening canopy. The process of dis-reefing shown below in Fig. 2 is done to reduce the forces felt by the manned occupants, and is important for unmanned PRS design from s structural integral aspect. Figure 3 - Dis-Reefing of a Hemispherical Canopy (Cao & Xu, 2004). Reefing a parachute slows the inflation, meaning more height loss, however there is another reason for reefing a parachute aside from reducing shock forces. There is a phenomenon called wake recontact, sometimes called canopy collapse. This phenomenon occurs when the parachute decelerates the payload so rapidly that the air behind the parachute catches up to the canopy: causing it to deform ( collapse ) and lose drag (Peterson, Strickland, & Higuchi, 1996). One of the single most important aspects in aiding the canopy inflation process is the use of a parachute deployment bag. A deployment bag ensures a controlled deployment phase. The deployment bag also ensures a lines first deployment. This ensures that the canopy only begins inflation after the suspension lines have reached full stretch. If a lines first deployment is not followed the canopy often gets tangled in the suspension lines and fails to inflate ending in catastrophic PRS failure. The increase in drag area of the parachute canopy during the opening process causes a deceleration of the recovery vehicle. A simulation of the deceleration experienced by the recovery vehicle was carried out and a plot of a typical g-loading experienced is shown in fig 11. This plot will be referred to later in this document, when the g-loading curve for this PRS is plotted. Figure 4 - G-Loading Plot Experienced by a Recovery Vehicle During Canopy Inflation (Mohaghegh & Jahannama, 2008). 6

7 D. Parachute Filling Distance Parachute filling distance is defined as the distance required for the parachute canopy to open, taken from the point of initial line stretch to full inflation. Fig. 4 demonstrates this definition (Mohaghegh & Jahannama, 2008). Mueller and Scheubel reasoned that, based on the continuity law, parachutes should open within a fixed distance, because a given conical volume of air in front of the canopy is required to inflate the canopy (Knacke, 1992). With the confirmation of drop tests the parachute filling distance was found to be proportional to the inflated parachute diameter D p, multiplied by the canopy fill constant n, as shown in Eq. 1 below (Mohaghegh & Jahannama, 2008). Figure 5 - Canopy Filling Distance Diagram (Mohaghegh & Jahannama, 2008). (1) The canopy fill constant, typical for each parachute type, is an indicator of the filling distance as a multiple of nominal parachute diameter. Having found the canopy filling distance only one further step is required to determine the canopy filling time. Given speed is distance over time, the canopy filling time is simple given as Eq. 2 (Mohaghegh & Jahannama, 2008). Canopy filling distance and canopy filling time are very important in PRS because they are a direct reflection of how much height loss may occur during the inflation process. Small UAV research aircraft have to operate at low altitudes and it is therefore imperative that the campy opens is a short distance. E. Attachment Considerations The attachment of the parachute to the UAV directly affects the operation of the system. The attachment points determine the behaviour of the aircraft during canopy inflation, and also the attitude at which the UAV will fall once inflation is complete and the PRS is in the steady state condition. Conventional PRSs deploy in such a way that the aircraft falls undercarriage first in order to protect the airframe; however this is not always the case. The Phoenix UAV, for example, has a PRS that allows the aircraft to roll over and land upside down (Wyllie, 2001). This is done to protect some of the sensor equipment underneath the aircraft. Manipulation of the aircraft attitude during steady state descent is generally achieved by changing the position of the PRSs attachments in relation to the aircrafts centre of gravity. This allows the designer the freedom to choose how the aircraft touches down; main undercarriage, or nose wheel first for example. In general however, PRSs are attached at several points with the centre of gravity of the aircraft roughly in the middle to keep the system balanced. From a structural perspective it is important to make sure that the attachments are connected to structurally sound aircraft fixtures, able to handle the large forces that can be experienced due to the rapid deceleration of the aircraft during canopy inflation. The attachments to the aircraft are also vitally important during the inflation stage, where careful placement of the attachment points can protect the parachute canopy from entering the wake of the aircraft. If the parachute is attached forward of the centre of gravity, deployment will cause a strong pitch up moment forcing the canopy into the wake of the aircraft. This situation may even cause the to aircraft fall backwards through the suspension lines, tangling the parachute. When attached behind the centre of gravity it causes a pitching down moment allowing the canopy to inflate in the free stream air and also stopping the aircraft from stalling. If used carefully the attachment method can be instrumental in controlling the pitch dynamics of the aircraft during the deployment cycle. F. Summary It is apparent from the literature review that there has been extensive research carried out on the subject of PRSs. There is little research however available on the use of PRSs for small balsa constructed UAV aircraft. There have been PRSs used for large scale UAVs, but the research available for UAVs under 25kg is very limited. (2) 7

8 III. Concept Selection A. Introduction The purpose of this chapter is to consider the different concepts available for the design of a PRS. Each concept will have merits, however the design will be chosen on its suitability for balsa wood constructed model aircraft.the requirements of the PRS will give the design process a direction. Listed here are the PRS customer requirements. These requirements have been used to generate engineering specifications, attached in ANNEX C 1. Minimise Damage to airframe on ground impact 2. Deploy Reliably 3. Minimal impact on aircraft aerodynamic stability when attached 4. Capable of decelerating a 25kg payload within respectable g-loading limits 5. The possibility of adaptation for different aircraft 6. Light weight 7. Minimal Height loss during parachute inflation The first stage of a PRS is the deployment phase. Parachute deployment denotes the sequence of events that begins with the opening of a parachute compartment and continues with the extraction of the parachute until the canopy and the suspension lines are stretched behind the recovery vehicle and the canopy is ready to start the inflation process. The deployment system must achieve this sequence of events in a progressive manner so as to reduce the shock forces felt by the UAV. The system must also and more importantly release the parachute into the airflow with sufficient speed so as to allow inflation to occur before the UAV losses to much altitude. This requirement is of significant importance to the design of this PRS as it is to be utilised for low altitude recovery of small model UAVs. A good parachute deployment system provides the following benefits as outline by the PRS Design Manual (Knacke, 1992): 1. Minimizes the parachute snatch force by controlling incrementally the deployment of the parachute, and by keeping the parachute canopy closed until line stretch occurs. 2. Keeps Tension on all parts of the deploying parachute. 3. Minimizes opening time. 4. Supports uniform deployment. Once the deployment phase is complete the system is ready for inflation and the most problem prone phase of the working system is over. It is apparent why the selection of the deployment system defines and directs the remainder of the PRS design. Four separate parachute deployment system concepts will be analysed, the best system for the use on the ADFA SAE Aero UAV will be selected for further design. 8

9 B. Concept 1 Uncontrolled Deployment This method of deployment essentially releases the parachute into the free-stream and the forward moving airflow does the work in opening the parachute to full inflation. It is therefore given the name uncontrolled deployment. The simplified diagram below demonstrates an uncontrolled deployment. Figure 6 - Uncontrolled Deployment (Cartwright, 2008). The appeal of this method of deployment for use on a UAV is its simplistic and lightweight design. An uncontrolled deployment would allow for a very fast developmental phase and also reduce the cost of the overall PRS. Furthermore, this deployment method could be constructed with very few components and would in turn allow the PRS to occupy a very small proportion of the recovery vehicle. This system is not without its downfalls however. Due to the uncontrolled nature of this concept, there has been research that has shown it is only effective for parachutes less than 5 feet in diameter (Knacke, 1992). When larger parachutes are released using this method it results in high shock forces and partial canopy inflation prior to the full line stretch of the suspension lines. It is possible to account for these problems by clever design of a complicated reefing system and sophisticated packing methods; however this would undermine the simplicity of the system which was one of the major reasons for its consideration. C. Concept 2 Ballistic Deployment Ballistic deployment refers to the deployment of a parachute by the use of some kind of pyrotechnics. There are two main types of ballistic deployment systems, and although they are subtly different their principle is essentially the same. The diagram below shows the two main ballistic deployment systems, rocket deployment on the left and drogue gun deployment on the right. Figure 7 - Ballistic Deployment 9 (Cartwright, 2008).

10 Rocket deployment uses an angled nozzle rocket to pull out the main parachute. This is a very high speed deployment, and has been used successfully for larger aircraft PRS. The benefit of the rocket deployment system is that it provides a constant pulling force over the whole deployment cycle. The disadvantages of this system are numerous when considering the type of vehicle in which it is intended for. The small balsa constructed ADFA SEA Aero UAV does not have a large amount of space in which a rocket could be mounted. Rockets also produce very hot exhaust gasses that may be an added risk for the internal balsa structure and electronic equipment such as servos. Drogue gun deployment is a method that fires a weighted slug from a drogue gun to pull out the main parachute in a very similar way to the rocket deployment system. This method is advantages in that the explosive energy is contained within the drogue gun and it leaves with the slug in the form of kinetic energy. As shown in the diagram above the slug is tethered to the main deployment bag and pulls out the main parachute from its compartment within the UAV. Unlike rocket deployment the slug relies on its inertia to provide the pulling force needed over the whole deployment. Problems with this method may arise due to the large reaction forces that occur due to the large slug being fired from the drogue gun. The simple balsa frame would struggle to support the drogue gun and damage may occur during deployment. Ballistic deployment has its merits with regards to a favourable deployment time. Having a fast deployment is paramount for low level UAVs and ballistic deployment has proven effective in this area. The complexity of ballistic deployment however, makes it less of a viable option for such a small scale recovery vehicle. The lack of space within the ADFA SAE Aero UAV also means that the system must have as few components as possible. Another consideration is the overall weight of a ballistic system, which, if installed, would require the aircraft to sacrifice this weight in the form of its payload carrying capacity. D. Concept 3 Spring Release This spring release method, represented diagrammatically below, was an attempt to combine the high speeds associated with ballistic deployment with the simplicity of the uncontrolled deployment. This concept uses a high powered spring internally mounted to eject the main parachute into the free stream airflow. Figure 8 - Spring Release Method (Cartwright, 2008). The side effect of using the spring release method is the weight of the internal mechanical components. One possible solution is to use compressed air instead of a mechanical spring however, problems also may arise with mounting the system within the ADFA SAE UAV fuselage, and although much less than a ballistic deployment the reaction force may still prove significant. One further point to note is that force provided by the spring to deploy the parachute must be significant enough to complete the whole deployment. Unlike a Rocket deployment where the force extracting the parachute is constant throughout the extraction phase, the spring release mechanism merely ejects the parachute and allows the airflow to complete the unfurling process. This reduces the reliability and repeatability of the deployment. 10

11 E. Concept 4 Pilot Chute Deployment The final concept to discuss is the method of pilot chute deployment. During this method of deployment a small parachute called a pilot chute is used to drag out the main parachute as demonstrated in the diagram below. Figure 9 - Pilot Chute Deployment (Cartwright, 2008). The pilot chute, which is generally about a quarter of the size of the main parachute, is small enough that it can be released at high speeds without producing large shock forces. The pilot chute then helps to decelerate the fore body while pulling out the main parachute so that when the main canopy opens the speed is slowed enough such that the shock forces are significantly reduced when compared to an uncontrolled deployment. The pilot chute utilises the drag force it creates when in the free airstream to deploy the main parachute and as a result has the benefit of having a constant force throughout the whole deployment phase. This does mean however, that at the very slow speeds the pilot chute can fail to produce enough drag to deploy the main parachute fully. This is one of the reasons that some aircraft PRSs have an operational envelope that defines its limitations. The issue of containment space is less of a problem using this method. Because the pilot chute is made of the same material as the main canopy it can be folded to take up very little space within the fuselage of the recovery vehicle. This naturally applies to the weight of the PRS. Because the pilot chute is made of lightweight material the overall deployment system is very light. The simplicity of this system is a direct result of the number of components required for its function. The deployment system is essentially two parachutes of different sizes opened one after the other. The construction and installation of it is therefore greatly simplified and as a result the cost is reduced. Using a pilot chute for the main parachute deployment does however, have a certain unknown quality to its function. Because the parachutes are different sizes, unique to their specific function; and the recovery vehicle has a known range of speeds defined by its operational envelope, the inflation times and distances are not known with great accuracy. For this type of deployment system testing would be required to determine the feasibility for use on a small balsa UAV. F. Conclusion Four concepts have been introduced within this chapter and one will be selected for further design. Uncontrolled deployment is easily the least complex system and would require very little detailed design. It would be a cost effective design and would require very few resources to take it from concept to construction. The possible complications do however make this a poor choice for further design. The over simplification of the deployment leads to a loss of tension on the suspension lines and canopy which causes entanglement between different parts of the parachute. The speed of a ballistic deployment system would provide a good range of altitudes from which the aircraft could be recovered. The better performance achieved in deployment speed is far outweighed by the increased complexity of the design. Both rocket and drogue gun methods also increase the risk of damaging the aircraft during the violent deployment phase. Ballistic deployment methods carry an inherent risk and as a result the expense of its 11

12 implementation is significantly increased. Unfortunately this makes it a poor choice of deployment system for this type of UAV. To make the PRS a viable option it is necessary to make it as cost effective as possible. The goal behind the spring release method was to reduce the risk of a ballistic deployment by using a mechanical type system to release the parachute into the free airstream. The side effect of using an internal spring system is that a large force must be produced at the start of the deployment to allow the parachute to complete its full unfurling process. This large force has to occur over a very small period of time, resulting in large impulse, which the fuselage was not deigned to handle. To produce a large impulse it would require the use a large and heavy spring system that would make this method of deployment a bad choice for an aircraft application. The final concept shows the most promise for further development. It is the method that uses a pilot chute to extract the main parachute. This system is simple, lightweight and has the potential to be very cost effective. There is however the question of inflation times and distances, and testing will be required to confirm if it is a feasible system. Pilot chute deployment is used primarily for personnel parachutes, which is a testament to their reliability, an important factor considering its intended purpose as an emergency recovery system. The table below is a method of comparing the individual concepts to help with an informed selection. Each concept is ranked against the amended system requirements. The essential requirements carry greater weight in the selection process and are therefore doubled within the table. The best concept for a given system requirement is given a score of 1 and the worst a score of 4. The values attained are summed and the concept with the smallest total number is the concept that best fulfils all system requirements. Table 1 below is the results of the concept selection analysis. Concept 1 Uncontrolled Concept 2 Ballistic Concept 3 Spring Concept 4 Pilot Chute Reliability* Speed of Deployment* Lines First Deployment* Complexity Reaction Forces on Aircraft Adaptability to Other Aircraft Weight Cost Results Table 1 Concept Selection table (Cartwright, 2008). This method of selection confirms what the logic predicts, that concept 4 the pilot chute deployment system is the best system for this application with a score of 19. Ballistic deployment is the second choice; however its complexity rules it out as a viable option. 12

13 IV. System Design A. Parachute System Configuration During the conceptual design phase outlined in chapter 4 of this document, it was decided that the PRS would use a pilot chute deployment method. A more detailed breakdown of the parachute recovery system configuration is represented below. Figure 10 - PRS Layout (Cartwright, 2008). 1. Recovery Vehicle 2. Shock Force Attenuation Bridle 3. Main Parachute Riser 4. Main Parachute Swivel 5. Main Parachute Suspension Lines 6. Main Canopy 7. Main Parachute Deployment Bag 8. Pilot Chute Bridle 9. Pilot Chute Swivel 10. Pilot Chute A. Main Canopy Shape Understanding that this PRS is to be used as a lower atmosphere, subsonic aerodynamic decelerator steers the choice of parachute canopy shape to three different options. These options are Para foil, Cruciform, and hemispherical parachutes. As discussed in the literature review parafoil shaped designs, common in sports parachutes, are steerable systems, unnecessary for this application. The parafoil design also adds unnecessary complexity to the construction of the intended PRS. Cruciform or hemispherical parachutes are more commonly used in aircraft recovery systems. This PRS will use a hemispherical parachute design due to the larger drag coefficient than the cruciform parachute allowing for the use of a smaller parachute for the same descent speed. The cruciform parachute has a smaller opening force coefficient which means less shock force occurs during parachute inflation; however this is outweighed by the reliability of the hemispherical parachute. Hemispherical parachutes also have a lower canopy filling constant which reduces filling distance. Therefore less height is lost during inflation. Further reasons for choosing a hemispherical parachute shape are listed below. 1. Ease of construction. 2. Ease of packing. 3. Better consistency on opening. 4. Large inflation shock forces acceptable due to unmanned aircraft. 5. Large Drag coefficient. 13

14 One negative aspect of using hemispherical parachutes is their tendency to oscillate. Oscillation is a result of an unstable parachute where the vortex shedding initiates a rocking motion back and forth. Hemispherical parachutes often have a tendency to have large oscillations, up to ± 30. Research has shown however that these oscillations only occur at the higher descent velocities ( 9m/s), which is as much as 4 m/s faster than the descent velocity this PRS is designed to achieve. B. Canopy Sizing Having decided that the canopy shape will be hemispherical, the initial canopy sizing can be evaluated with the objective of calculating inflated parachute diameter. Parachutes rely on the aerodynamic drag force, represented by Eq. (3), to slow the descent of a body. Assuming that the parachute is in a steady state descent means that the drag forces F d can be equated to the weight of the descending body, as represented by Eq. (4). Rearranging Eq. (4) for surface area yields Eq. (5). (3) (4) The unknowns in Eq. (5) are, payload mass m, and descent velocity V. All other variables in the equation are already known quantities, ρ=1.225kg/m 3, C D =0.7(standard for hemispherical parachutes) (Knacke, 1992), and g=9.81m/s. Payload mass can be reasoned out by understanding that the ADFA SAE Aero UAV gross weight must not exceed 55lbs ( 25kg), therefore m=25kg. The descent velocity can be obtained from historical data, which shows that aircraft PRSs have descent velocities between 4-6 m/s (Wyllie, 2001). Taking the average and using 5 m/s for V means that surface area can be quantified. Now that surface area is calculated the nominal parachute diameter can be found. The nominal diameter of a parachute is simply a reference diameter found by using Eq. (6), (Knacke, 1992) below. (5) Using table 5-1 (Knacke, 1992), found in ANNEX D, the ratio between inflated diameter D p and nominal diameter D o for a hemispherical parachute is highlighted. This allows for the calculation of the inflated diameter using the Eq. (7) below. (7) The excel spreadsheet shown in table 1 was used to calculate the parachute data, using the equations outlined previously. The excel spreadsheet shows that for a 5m/s rate of descent, a parachute with a 3.56m inflated diameter is required. Research has shown that this method of parachute sizing is highly accurate, and conforms well to the results of parachute drop tests. Knowing the Canopy Sizing (6) Parameter Value Units Mass kg Gravity 9.81 m/s^2 Density 1.23 kg/m^3 Descent Velocity 5.00 m/s Drag Coefficient 0.70 Surface Area Required m^2 Nominal Parachute Diameter 5.40 m 14 Inflated Parachute Diameter 3.56 m Table 2 - Main Canopy Sizing Data (Cartwright, 2008).

15 parachute dimensions it was possible to purchase a hemispherical parachute of the appropriate size. The parachute was purchased early as it was known that testing would be required before the conclusion of this project. The main canopy cost a total of $350 including suspension lines and mounting bracket. C. Pilot Chute Design The pilot chute is a small parachute designed to pull out the main parachute during the deployment phase. Now that the main parachute has been sized and purchased the pilot chute can be designed using the main parachute data. Two different methods have been used to confirm the size requirement of the pilot parachute. The first method of sizing uses historical data from many parachute tests to determine a Pilot-to-main parachute drag area ratio. Table 3, right is taken from the Parachute Recovery System Design Manual, and was used to get a first estimate of the pilot chute size. Table 3 - Pilot-to-Main-Parachute Drag Area Ratio (Knacke, 1992). The first method estimated that a pilot chute with an inflated diameter of 0.62m was required. Pilot Chute Sizing Parameter Value Units Main Parachute Mass 1.40 kg The second method is a mathematical Gravity 9.81 m/s^2 method similar to that of the calculation of the main parachute size. Given that Density 1.23 kg/m^3 the main canopy had been purchased we Descent Velocity 4.00 m/s could get an accurate weight by a simple Drag Coefficient 0.70 measurement. The parachute including bridle has a total weight of 1.4 kg. Knowing the weight of the main Surface Area Required 2.00 m^2 parachute and assuming the worst case scenario, that the pilot chute must have the ability to pull out the main parachute Nominal Parachute Diameter 1.60 m at the stall condition of the aircraft (approx. 4 m/s), it is a simple Inflated Parachute Diameter 1.05 m calculation to solve for a surface area. This calculation was performed in parallel to the main canopy sizing method using Microsoft Excel and is displayed in table 4. The mathematical calculation determined that a hemispherical pilot chute of 1.05m diameter would be required for a successful deployment at stall speed. This size is 40% larger than the pilot chute diameter determined using statistical data. To ensure the reliability in deployment of the PRS and to build in another level of safety it was decided to use the larger of the two values. Hence the pilot chute is set at a diameter of 1.05 meters. The pilot chute was also purchased so that its reliability could be tested at a later stage within the scope of this project. D. Initiation Method From the conceptual design chapter a pilot chute system was selected for use as the deployment method. The initiation system is characterized by the functions of the PRS that occur up to and including deployment of the pilot chute. Several simultaneous actions must occur prior to the start of the deployment process. First the PRS must be initiated by the observer on the ground. A signal will be sent to the UAV that will release the pilot chute and also kill the engine. The engine will be killed to prevent line tangling in case the suspension lines come in the vicinity of the propeller. 15

16 The pilot chute is ejected by using a very lightweight spring mechanism. The spring is sewn into the pilot chute crown area and collapses for easy storage at the rear of the aircraft fuselage just behind the aft bulkhead. The pilot chute spring is compressed and held in place by the release hatch. The release hatch is pinned closed and a small electronic servo is mounted in place to release the hatch when required. The diagram right shows the folded pilot chute and spring aft of the rear bulkhead. The entire PRS is initiated by the pulling of a pin which holds the release hatch to the recovery vehicle. The hatch falls away as the pilot chute is sprung into the free airflow and the deployment process begins. The servo that pulls the release pin to activate the system is run off an independent power Figure 11 - Folded Pilot Chute and Compressed Release Spring Inside Release Hatch (Cartwright, 2008). source and is also wired in to an independent radio receiver. This is to try and avoid accidental parachute release by radio interference. E. Containment Now that the main parachute and pilot chute have been purchased, elements of how the system will be contained within the aircraft can be considered. To ensure that the PRS has no effect on the aircraft during normal flying operations, an internal storage will be used. The ADFA SAE Aero UAV does not have large amounts of fuselage space. By looking at the constructed fuselage the best location for containment can be determined. The newly designed fuselage for the ADFA SAE Aero UAV, designed by Christopher Kourloufas is shown below. Figure 12 - ADFA SAE Aero UAV Fuselage (Cartwright, 2008). The different sections annotated above indicate the lack of space within the main fuselage section. The two sections where it would be possible to fit the parachute internally are the internal electronics mounting space, or over the wing carry through section. To facilitate deployment it is important that the parachute be in a location where it can deploy quickly into the free airstream and be unaffected by aircraft appendages such as the wings or canards. This makes the obvious choice for the parachute containment at the rear of the aircraft, such that it can deploy already 16 Figure 13 - ADFA Final SAE Thesis Aero Fuselage Report 2008, with PRS UNSW@ADFA Location (Kourloufas, 2008).

17 free and clear of the wings, avoiding possible suspension line tangling. Figure 13 is a sectioned cut-away of the fuselage design which indicates the PRS location within the UAV. Knowing the location of containment produces the need to design a method of storing the parachutes. This requires the design of a deployment bag. F. Deployment Bag Using a Parachute deployment bag has many advantages. It allows for a repeatable parachute unfurling process and as a result increases the reliability of deployment. It has been shown to provide the most frequent and successful method of producing a controlled deployment (Knacke, 1992). A deployment bag also permits easy handling, packing, rigging and storage of the parachute assembly. The deployment bag contains the parachute and holds it until it is ready to be unfurled. This means that it must have a few very specific qualities to allow for a good controlled deployment. Firstly the deployment bag must be light. The pilot parachute attaches to this bag and pulls it out of the containment area, so the lighter the bag the smaller the pilot chute required. The deployment bag must also have a low porosity. This ensures that the main parachute canopy can easily slide past the bag when unfurling is required. The deployment bag also ensures that the canopy only inflates after the suspension lines are at full stretch. As brought up in the literature review, a lines first deployment is imperative to reduce shock forces and ensure reliable canopy inflation. Generally speaking a deployment bag is specifically made for the main parachute. In this case an existing product was available that adequately performed all of the necessary functions of a deployment bag. A 2.4 Litre plastic bottle was found to be an excellent deployment bag for this situation. The bottle was light, cheap, fitted perfectly into the allowable space, and has a sufficiently low porosity to allow the parachute to unfurl quickly. The deployment bag encompasses the main parachute into a small space and has the added benefit of easily being attached to the pilot chute. Figures show the deployment bag, and its location within the PRS. Figure 14 - Unpacked Main Parachute and Deployment Bag (Cartwright, 2008). Figure 15 - Main Parachute Packed in Deployment Bag (Cartwright, 2008). Figure 16 - Main Parachute in Deployment Bag with Pilot Chute Attached (Cartwright, 2008). The final requirement for the deployment bag is to ensure a lines first deployment. This is achieved by using a method called skirt hesitation. This design uses a tensioned band around the suspension lines at the base of the deployment bag. It is demonstrated in the figure 17. This red band prevents the skirt of the parachute leaving the deployment bag until line stretch occurs. At line stretch the suspension lines are pulled out from underneath the band and the parachute can be deployed. G. Canopy Inflation Predictions The canopy inflation process is by far the most complex aspect of the PRS. There are several mathematical models to help predict the effect of the canopy inflation process and some of these models will be utilised within this preliminary design chapter, however the Figure 17 - Deployment Bag with Red Sirt Hesitation Band (Cartwright, 2008). 17

18 canopy inflation process must, and will be supported by testing. The reason for such focus on the canopy inflation process is to try and answer the question will this PRS deploy quickly enough before the aircraft looses too much altitude. There is also a need to calculate the opening forces of the canopy for the use in stress analysis of both the parachute and airframe. As discussed in the literature review regarding canopy inflation time and distances, a simple equation exists to give estimate the canopy inflation time. This equation is shown below; D o is the parachute nominal diameter calculated earlier this chapter. V is the velocity at line stretch and n is the canopy filling constant. Using this technique the first estimation of canopy filling time for both the main parachute and the pilot chute was calculated for different deployment velocities and displayed in the Figure 18. This mathematical prediction of canopy filling time indicates that for speeds above 5m/s both parachutes will inflate in less than 5 seconds. (8) Based on the continuity law parachutes should open within a fixed distance; because a given collum of air in front of the canopy must be used to inflate it. This distance is proportional to the diameter of the parachute (Knacke, 1992) and by using the equation below; Figure 18 - Mathematically Predicted Canopy Inflation Times (Cartwright, 2008). A constant value or parachute opening distance for the main canopy of m is predicted. This equation was also applied to the pilot chute and a much smaller canopy inflation distance of 4.21m was calculated. This method of calculating canopy inflation time and distance is used only as an initial estimate for the design. Given that it predicts the main canopy will inflate in a distance as little as 14.25m it provides evidence that this PRS design will deploy to full inflation even at low altitudes. It is imperative that these initial estimates however be validated by testing. Now that inflation times and distances have been estimated, parachute opening forces should also be considered. The inflation distance calculation gave an indication as to whether the canopy could open within a given altitude. Canopy opening force calculations will help determine whether or not a small balsa constructed UAV can handle the opening forces of a parachute at a given deployment speed. This force calculation will be important in trying to determine the mechanical feasibility of this PRS. There are several methods of determining the maximum canopy opening force. The method used here is known as the Pflanz Method, which was developed during WWII by E. Pflanz (Knacke, 1992). This method makes the assumption that the aircraft if flying on a horizontal path at the time of deployment. The first problem is to determine how much velocity decays from the UAV after the parachute is deployed but prior to the commencement of canopy inflation. This is done by using the equation below; (9) (10) 18

19 Velocity decay was found to occur prior to the inflation process. The magnitude of the velocity decay was taken away from the initial deployment speed and the resultant was used in the remainder of the opening force calculation. The amended velocity was now appropriate to use within the force equation below; The force equation above is very similar to the standard drag force equation 3. The different variables are the opening force coefficient C x and the force reduction factor X 1. The opening force coefficient is a tabulated value in ANNEX D, and the force reduction factor is found by calculating the dimensionless ballistic parameter, A using the equation shown. (11) The force reduction factor can now be found by using the relationship between force reduction factor and the dimensionless ballistic parameter A displayed in ANNEX E. Using this value for force reduction factor the parachute opening force can be determined for given deployment velocities and was graphed using excel to give the graph in figure 19. The mathematical data predicts a peak parachute opening force of approximately 800N at 20m/s deployment velocity. This indicates that even for lightweight recovery vehicles there is a substantial canopy opening force. Because this force occurs within the short time that the canopy inflates it is often referred to as the shock force. A deployment at 20m/s will produce considerable shock force and the (12) Figure 19 Mathematically Predicted Parachute Opening/Shock Forces for Different Deployment Speeds (Cartwright, 2008). maths indicates that a shock force attenuation system may be required. This may be in the form of an aerodynamic reefing system, or even a simple system such as an elastic riser to absorb some of the shock force during opening. H. Summary The major components of the PRS have been designed within this chapter. These components include the main canopy, pilot chute and deployment bag. At the very least these systems needed to be designed prior to determining if it were possible for a PRS to safely recover the UAV. To begin to understand if the PRS can safely recover the UAV in an appropriate altitude some preliminary calculations were made to determine the parachute filling times and opening distance. Canopy opening force predictions were also made which seem to indicate that the PRS may produce too much opening force for the airframe to handle. More accurate predictions about how much altitude will be lost and the amount of force the airframe will receive during a parachute deployment, need to be made. The chapter to follow is aimed at answering these questions so that an indication of the feasibility of the PRS can be deduced. 19

20 V. Testing and Analysis A. Introduction Due to the nature of the flight envelope of small model UAV research aircraft questions arise as to the suitability of PRSs. The mathematical calculations in the previous chapter seem to indicate that it is possible for a parachute canopy to fully inflate without losing to much altitude. These calculations are only an estimate however and testing needs to occur to back up the mathematics. This chapter describes the appropriate strategies, and methodologies used to plan, and execute testing of the parachute opening characteristics, to determine their mechanical feasibility for model aircraft. The objectives of the testing is to gain an insight into the; Parachute inflation times. Parachute inflation distances. Parachute reaction to wake flow. Confirmation of parachutes structural integrity. Parachute Recovery System Reliability and Repeatability. Deployment method verification. B. Method The most common method of testing parachutes is in the wind tunnel. This off coarse is the safest method and produces good reliable results. Unfortunately you are limited by the size of the parachute that can fit inside a wind tunnel, and using a smaller model parachute can compromise the dynamic similarity and give meaningless results. For this reason, along with the limited resources that are available at the Australian Defence Force Academy, a simple method designed to reproduce the effects of a wind tunnel test. The basic test approach will involve dynamic parachute testing. The method will make use of a motor vehicle to induce different parachute opening speeds. The parachute will be contained in the deployment bag until the vehicle reaches the appropriate deployment speed. Once this occurs the parachute will be deployed and allowed to inflate, while being recorded for analysis. This will be repeated several times at different deployment velocities. A frame by frame demonstration of the parachute inflation process resulting from the testing method is displayed in ANNEX F. C. Results and Validation The testing occurred on a calm day to ensure that the results would be as reliable as possible. All of the test runs were filmed for later analysis. The freeze frame image in figure 20 illustrates the fully inflated canopy during a test run and demonstrates how the testing was carried out. Runs were conducted, at 10km/h and increased in 10km/h intervals until 50km/h was reached. Each deployment speed was repeated several times so that an average of the results could be attained. The average of the results is displayed in table 4. Deployment Speed (km/h) Inflation Time (s) Inflation Distance (m) Table 3 - Experimental Results (Cartwright, 2008). 20 Figure 20 - Captured Image from Dynamic Test Run (Cartwright, 2008).