Rocket Project PROJECT MANAGEMENT & DESIGN II MAE 435. March 4th, Class Supervisors: Group Members: Nathan Akers. Dr.

Transcription

1 PROJECT MANAGEMENT & DESIGN II MAE 435 Rocket Project March 4th, 2015 Group Members: Nathan Akers Chad Bagley Paul Campbell Class Supervisors: Dr. Sebastian Bawab Dr. Colin Britcher Mr. Michael Polanco Charles Juenger Brian Mahan Chris Skiba Project Advisor: Dr. Thomas Alberts Michael Weber Michael Wermer

2 2 Contents Nomenclature Abstract Introduction Structures Introduction Completed Methods Proposed Methods Propulsion Fuel and Combustion Chamber Design Introduction Completed Methods Proposed Methods Nozzle Design Introduction Completed Methods Aerodynamics Introduction Completed Methods Proposed Methods Avionics Introduction Completed Methods... 19

3 3 6.3 Proposed Methods Results Discussion Appendix Additional Figures Gantt Chart Budget References List of Figures Figure 1: Rocket Components... 7 Figure 2: Rocket Assembly... 8 Figure 3: Thrust vs Time Figure 4: Combustion Chamber with Convering-Diverging Nozzle Figure 5: Velocity Contour Figure 6: Temperature Contour Figure 7: Coefficient of Drag vs Mach Number Figure 8: Fuel Performance Figure 9: Predicted Flight Performance Figure 10: Rocket Nozzle Data Figure 11: Dual Deployment Wiring Diagram Figure 12: Avionics Bay Arrangement Sketch Figure 13: Rocket Motor... 26

4 4 Figure 14: Rocket Upper Body Figure 15: Avionics Bay Figure 16: Rocket Lower Body Nomenclature r Propellant Burn Rate P 0 K n ρ p P exit P atm Combustion Chamber Pressure Propellant Burn Area to Nozzle Throat Area Ratio Propellant Density Nozzle Exit Pressure Atmospheric Pressure M throat Nozzle Throat Mach Number M e T V exit Nozzle Exit Mach Number Thrust Nozzle Exit Velocity!m Mass Flow Rate A * D * D e R T 0 γ Nozzle Throat Area Nozzle Throat Diameter Nozzle Exit Diameter Gas Constant Combustion Chamber Temperature Ratio of Specific Heats!m Mass Flow Rate X * Throat Position

5 5 1 Abstract The goal of the project is to design a high-powered rocket capable of taking a cosmic ray detector payload to an apogee of 10,000 feet, deploying the payload, and being recovered. Detailed designs of the structure, propulsion system, aerodynamic fins, avionics, and recovery system were developed using computer aided drafting software (CAD). Structural stability will be determined through finite element analysis (FEA). The propulsion system will be powered by a potassium nitrate (KNO 3 )-dextrose (C 6 H 12 O 6 ) solid propellant that will undergo combustion and expansion through a converging-diverging nozzle, delivering up to 5,120 Newton-seconds of impulse. The expected thrust for the fuel was calculated and an optimized nozzle was designed for the fuel. Aerodynamic analysis was performed through modeled computational fluid dynamics (CFD) to determine coefficient of drag versus Mach number. Scaled rocket supersonic wind tunnel testing will be performed to determine the accuracy of the CFD analysis. The avionics system will autonomously record in-flight data and control the deployment of the payload, the drogue and main parachutes, ensuring safe recovery after flight. Fundamental vibration analysis will be performed to minimize in-flight vibrational effects on the payload and avionics system. All analysis and testing performed on the rocket will allow the team to extract the necessary information to validate the design and ensure a safe and reliable flight.

6 6 2 Introduction Rockets are used to carry payloads into orbit, but are expensive and difficult to design. They need to be designed for the type, size, and weight of the payload to be carried. They need to be designed for the desired payload deployment height. The design requirements of this project are to design a high-powered rocket capable of deploying a five pound cosmic ray detector at 10,000 feet. The rocket design was divided into four key areas necessary to satisfy the goal of the project. These include: structures, propulsions, aerodynamics, and avionics. 3 Structures 3.1 Introduction A high powered rocket is composed of cylindrical body, nose cone, avionics bay, tail fins, motor assembly, and recovery system. CAD software is used to create an accurate visual representation of the rocket and its components. The representation is used to perform analysis that determine design feasibility prior to construction [1]. FEA is a method for determining stress and strain properties on complex solid bodies due to applied forces [2-4]. FEA is used on the rocket assembly to design for structural integrity and to verify material selection [5]. The analysis is needed to determine if the rocket design is capable of successful flight. Every new rocket design needs to be tested for structural integrity. 3.2 Completed Methods A 3D model of the rocket was created using Autodesk Inventor (Autodesk Inc., San Rafael, CA) and was developed by designing part assemblies of the different rocket components. These components (Figure 1) were then joined together to produce the main rocket assembly (Figure 2).

7 7 Figure 1: Rocket Components The first part assembly to be designed is the avionics bay (Figure 15), which houses the electrical equipment and connects the top and lower portions of the rocket. The ends are fitted with two circular end caps and are secured to the tubular housing by two threaded rods with accompanying fastener hardware. Eyebolts are fitted to the center of the two end caps by fasteners. These bolts serve as an anchor point for the shock chords after in flight rocket separation has occurred. The nose cone (Figure 14) is designed with a hollow cavity surrounded by the nose cone walls. A circular end cap with a U-bolt and the accompanying fasteners are fitted to the end of nose cone. The U-bolt serves as another anchor point for the top shock chord. Connecting the nose cone and the avionics bay is the upper rocket body (Figure 14). The lower

8 8 rocket body is a tube fitted with six fins (Figure 16). The engine assembly is made up of a nozzle, outer casing, resin liner and end cap (Figure 13). This is attached to the lower rocket body using six circular motor mounts attached to inner walls of the bottom rocket body. Connecting the avionics bay to the nose cone are Kevlar shock chords. Connecting the nose cone to the upper body and the avionic bay to the lower body are three shear pins for each connection. The avionics bay is connected to the upper rocket body using three riveted joints. Figure 2: Rocket Assembly

9 9 3.3 Proposed Methods The FEA will be carried out through the application of NASTRAN In-Cad (Autodesk Inc., San Rafael, CA). The test to be performed will consider the outer rocket body and the stress put on it from thrust. The maximum thrust force limit will be applied during the FEA analysis. Stress testing will be carried out on the shear pins to determine the force required to cause shear. A testing fixture will be constructed with two interlocking metal tubes that form a couple. A bulkhead will be fixed to each end of the couple with an eyebolt running through the center of each. Three equally spaced 5/64-inch holes will be drilled through both pipes in the couple. After the three shear pins are inserted into the corresponding holes in the couple, the test fixture will be secured to the MTS Alliance RF/300 electromechanical load frame in the ODU materials laboratory. The MTS machine will then be set to expand causing tension in the testing fixture. The data will be recorded once the shear pins fail and the couple separates. The MTS machine will be reset and a new set of shear pins will be secured into the rejoined couple. This testing procedure will be repeated until six tests have been performed. A test stand is under development that will be used to measure rocket thrust. The stand will utilize an Interface load cell and be mounted in the vertical position. The test will use the same motor that was designed for the rocket, but with half of the amount of fuel. The data will then be used in conjunction with the fuel calculations to predict the amount of fuel that will be needed to deliver the designed total impulse. The rocket will be assembled in steps starting first with the motor mount. The fins will be installed through the outer body of the rocket and be glued to the motor mount. Then the avionics

10 10 bay will be assembled. After the avionics bay is built, the upper sections of the rocket will be assembled. The motor will be the last piece to be installed. 4 Propulsion 4.1 Fuel and Combustion Chamber Design Introduction Rocket motors are thrust engines, which operate by generating a pressure and momentum thrust. This is achieved by combusting fuel inside a pressure vessel and expanding the combustion products through a nozzle. This is typically accomplished in small-scale rockets by using solid rocket fuels. Commercially available high-power hobby rocket motors are typically made of Ammonium Perchlorate, HTPB and Aluminum, which is similar to the formulations used in rockets designed for space flight as well as missiles. These formulations, commonly referred to as APCP (Ammonium Perchlorate Composite Propellant), provide a high specific impulse relative to other solid formulations. Specific Impulse is a key concept in the design and of a rocket motor because it is used to describe the efficiency of a fuel based on its weight. Specific Impulse is the ratio of the amount of thrust a fuel is able to produce to the weight flow of the propellants [6]. One drawback to using an APCP is the production cost. The material cost is high and the manufacturing process can be difficult. An alternative to APCP, which is commonly used in homemade rockets, is Potassium Nitrate-Sugar propellant, sometimes referred to as R-Candy. This formula is easy to manufacture using readily available materials. The greatest drawback to using these R-Candy formulations is the relatively low specific impulse they produce; meaning to reach a given altitude, more fuel is required. These formulations can be modified using different sugars such as sucrose, sorbitol, or dextrose, which can impact the material properties of the casted propellant, and can influence the ease of manufacturing, and the

11 11 reliability of the motor itself. While general formulations are available, the amounts and types of fuel need to be modified in order to reach the target payload deployment altitude Completed Methods The rocket fuel chosen is an experimental fuel, meaning that the motor is to utilize a homemade propellant formulation in lieu of a commercially available motor. The formulation was taken from Richard Nakka s Experimental Rocketry Website [7]. The need to select a preinvestigated formulation was necessitated by the compressed budget and timeline of this project as there were not enough resources available to develop a unique formulation which could function safely and reliably. The propellant selected for this project was a mixture of 65% by mass Potassium Nitrate (KNO 3 ) and 35% by mass Dextrose (C 6 H 12 O 6 ) and is seen below [8]. C 6 H 12 O KNO 3 à CO CO H 2 O H N K 2 CO KOH (Equation 1) (theoretical combustion reaction at 68atm [8]) This formulation was tested extensively and is commonly used by high-power rocketry hobbyists. This relatively inexpensive formula has been proven to be safe and reliable while being relatively easy to manufacture, and produces results within a reasonable margin of error by using the published values. A preliminary motor design was developed using a freeware Microsoft Excel (Microsoft Corporation, Redmond, WA) spreadsheet ( SRM_2014 ) published by Richard Nakka [9] to target a maximum operating pressure within the combustion chamber of 1100psia. This target pressure was based off a preliminary motor casing design, to be constructed from 2.95 outside diameter, 6061-T6 Aluminum tubing with a wall thickness of 0.11, giving an inside diameter of The selected pressure of 1100psia used a conservative factor of safety

12 12 of 2, this conservative value was selected because of the need to re-use the motor casing multiple times and the relatively low-impact of the casing weight on the rocket performance and design. The size of our motor was scaled to deliver as close to 5120 Newton-seconds of impulse as possible, without exceeding that limit. The desired pressure is achieved by modifying the K n value, which is the ratio of burning surface area of the propellant to the area at the throat of the nozzle [10]. This was done by selecting a commercially available propellant casting mold and modifying the nozzle diameter to meet the required K n value. Using the freeware Microsoft Excel (Microsoft Corporation, Redmond, WA) spreadsheet ( SRM_2014 ) published by Richard Nakka [9] and the fuel amount dictated by the size of the combustion chamber, plots of Thrust vs Time (Figure 3) were obtained. Additionally, the predicted altitude, flight time, and max velocity using a conservative coefficient of drag of 0.35 was estimated. SRM 2014.XLS Thrust vs time Thrust (lbf) Time (sec.) Figure 3: Thrust vs Time

13 Proposed Methods Now that the fuel has been selected it will be manufactured. The process of manufacturing the fuel requires the two components to be thoroughly mixed together and then heated until the dextrose melts, homogenously binding the two products together. The particles must be desiccated and finely milled before mixing and heating to ensure the maximum performance of the propellant. Errors in manufacturing can cause results which are significantly less than predicted. The particles will be mixed together using a rotating drum or other low-speed mixture to ensure an evenly distributed mixture. The dextrose will fully melt at 123 C [7], in order to create the propellant; the particle mixture will be heated between C. The sugar will begin to caramelize at 157 C [7] so it is important to use a calibrated hot plate to precisely control the temperature; caramelization of the sugar will negatively impact the performance of the propellant. While heated, the mixture will be packed into a inside-diameter, cylindrical casting tube, placed around a cylindrical core with a diameter of To form the motor, six of these castings will be made at a length of 4.83 ; each of these castings is known as a grain. The completed motor will be comprised of six grains, totaling long with an outside diameter of 2.562, and a core diameter of , for a total volume of 132in 3. After the fuel has been mixed, the rocket motor manufactured, and the test stand manufactured, the motor will be test fired. Pressure will be measured using a PX-303 load cell (Omega, Stamford, Connecticut) and the thrust data is measured using a PX-LCJA (Omega, Stamford, Connecticut). The data will be collected from the static test using a DI-194 (DATAQ Instruments, Akron, OH). The data from this test will be used to validate the software simulations and used to adjust the fuel component percentages in order to achieve the required thrust.

14 Nozzle Design Introduction In order to achieve maximum propulsive efficiency, a rocket motor must sustain a high flow of heavy particles and elevated combustion temperatures while maximizing the speed of gases through the nozzle exhaust such that the exit pressure is equal to the atmospheric pressure [11, 12]. Exponentially increasing temperatures in the combustion chamber can cause rapid deformation of the materials of the nozzle and chamber [11]. Combustion gas expansion induces high internal pressures within the combustion chamber, which further exacerbates component deformation. Gases that travel through the nozzle carry burnt aluminum particles, which cause ablation to the throat section of the nozzle and agglomeration in the nozzle cone [13-15]. This results in decreased efficiency by altering the nozzle inner dimensions. Additionally, heat transfer in the chamber must be controlled, as heat conducted through the chamber walls results in heat loss to the combustion gases [16]. The inability to accommodate these operating conditions results in performance inefficiencies as well as potential component failure. Determination of a nozzle design, which accommodates maximum operating conditions, will increase performance and longevity while reducing cost. The nozzle dimensions must be individually designed for every rocket because of unique fuel.

15 Completed Methods The nozzle (Figure 4) was assumed to be isentropic, and compressible flow equations were used to determine the thermodynamic properties at the subsonic, sonic, and supersonic regions. The properties at each station and the constraint of three inches on the converging section were used to determine the dimensions of the nozzle. To reach maximum propulsive Figure 4: Combustion Chamber with Convering- Diverging Nozzle efficiency, the exit gases must be expanded through the diverging section of the nozzle so that P exit equals P atm. The mass flow rate in the nozzle was determined by!m = Thrust V e (Equation 2) where V e = 2γRT c γ 1 " 1 P % e $ ' # & P c γ 1 γ (Equation 3). Using the mass flow rate and exit velocity, the throat diameter was determined using the equation D * = 4!m RT c π P c X * (Equation 4) and X is a function of the ratio γ+1! of specific heats X * 2 $ 2( γ 1) ( γ) = γ # & (Equation 5). Next, the nozzle exit diameter was found " γ +1%

16 16 by first finding the Mach number at the exit by M e = ( 2 *" $ γ 1* # )* P c P e % ' & γ+1 2 γ 1 ( ) (Equation 6).,- Substituting the Mach number into the equation for area ratios provided the diameter for! the exit D e = A * 4 # 1 (! 2 $! # & 1+ γ 1 π M e " γ +1% 2 M 2 $ + * # e &- # ) " %, " γ+1 2 γ 1 ( ) $ & (Equation 7). A simulation of the nozzle & % was run at the determined area and pressure ratios using the CD Nozzle Simulator (engapplets, Blacksburg, VA). Using Simulation CFD (Autodesk Inc., San Rafael, CA) a CFD analysis was performed of the nozzle with the inlet conditions found in the fuel analysis. The assumptions were a fluid of air, compressible flow, sea level standard air conditions, steady flow, and no heat transfer. The CFD analysis supplied velocity (Figure 5) and temperature (Figure 6). Figure 5: Velocity Contour

17 17 Figure 6: Temperature Contour 5 Aerodynamics 5.1 Introduction Aerodynamics is examined to reduce drag and shock effects. Shapes of nose cones and fins have been examined in an effort to reduce drag and shock effects [17-22]. Nose cone designs showed that increased ratios of nose cone length to body length decreased drag, and reduction of the angle of the nose cone resulted in a smaller shock transition between the nose and the rocket body during supersonic flight [18, 19]. Fin designs were primarily focused on the flight stability, but as the size and bluntness of the leading edge increased, the drag and shock effects increased [17,20-22]. While the basic aerodynamics of fins and nose cones are known, the final fin designs and nose cone designs are modified for every rocket based upon mass, and length of the rocket in order to move the center of pressure below the center of mass. This ensures flight stability. 5.2 Completed Methods The software chosen to perform CFD analyses on the fins and nose cone was Simulation CFD (Autodesk Inc., San Rafael, CA). The assumptions for the flow to be used in the CFD

18 18 analyses were a fluid of air, compressible flow, sea level standard air conditions, steady flow, and no heat transfer. The inputs were Mach numbers from 0.1 to 1.2 in step sizes of 0.1. The output criteria to judge the different designs were coefficient of drag and center of pressure. The Coefficient of Drag vs Mach number are plotted below (Figure 7). Cd vs Mach Cd Mach Figure 7: Coefficient of Drag vs Mach Number 5.3 Proposed Methods A model rocket will be used to test in the supersonic wind tunnel at Old Dominion University. A supersonic wind tunnel test was chosen because it will be inexpensive and easier to perform. This is because the models are smaller and there are a few model rockets already available for choosing, but the fins need to be modified. The model will also be threaded for attachment to the wind tunnel apparatus. A Mach number of 1.8 will be used since that is the lowest speed this wind tunnel can operate. A CFD analysis will be done at a Mach of 1.8 to ensure that the CFD program used is accurate. Schlieren photography will be used to see the flow of fluid and shockwaves across the nose cone.

19 19 6 Avionics 6.1 Introduction The avionics system will consist of a redundant, dual-deployment altimeter system capable of ejecting the drogue and main parachutes. The typical wiring diagram for dual deployment altimeters is seen in the appendix (Figure 11). The dual-deployment altimeters will be programmed to deploy two separate parachute ejection charges that will allow failure of the rocket airframe shear pins. The ejection charges and airframe shear pins will be sized using common rocketry calculators [23] and experimental bench testing. The first ejection charge will cause shear pin failure of the rocket s forward airframe and will release the drogue parachute and cosmic ray detector payload. The drogue parachute will be deployed at the rocket s apogee and used to decelerate and stabilize the descent of the rocket. When the drogue parachute is deployed the payload will be ejected. The second ejection charge will cause shear pin failure of the rocket s rear airframe and will deploy the main parachute at approximately 800 feet above the ground. 6.2 Completed Methods The electronics, electrical mounting structures, parachutes, shock cords, and rocket airframe shear pins have been ordered. A conceptual sketch has been developed for the arrangement of the avionics bay section of the rocket and is in the appendix (Figure 12). Software Simulation Mechanical (Autodesk Inc., San Rafael, CA) has been acquired and will enable the team to perform the vibration and stress analysis using the rocket model. To gain an understanding of how the software can be used, the team has imported the Inventor model of the avionics bay of the rocket s airframe and have ran a preliminary design analysis report. 6.3 Proposed Methods Once construction of the avionics bay is complete, the bay will be fully modeled in Autodesk Inventor and incorporated into the rocket model. The rocket model will then be

20 20 imported into Autodesk Simulation Mechanical (Autodesk Inc., San Rafael, CA) to determine the effects of vibration and stress on the avionics bay section of the rocket. Using Autodesk Simulation Mechanical (Autodesk Inc., San Rafael, CA), constant random vibration analysis (modal superposition) will be conducted. The results provided from the analysis are the root mean square response of the displacement is equal to the standard deviation. From this analysis, a design analysis report will be generated and included in the final report. Based on the results from the analysis, action will be taken to mitigate the effects of vibration on the electronics (e.g. adding rubber isolators to the electronic mounting points). Ejection charge and shear pin sizing will be performed, first using commonly available calculators used in model rocketry [23]. Data obtained from these calculators will be used as a baseline to size the ejection charges and shear pins for experimental testing. Bench tests will be performed on the assembled rocket testing the rocket s separating force. Bench testing will verify the charges are sized appropriately, allowing failure of the rocket airframe shear pins during flight. The manufacturer of the dual-deployment altimeters provides free downloadable software (mdacs Missle Works Data Acquision and Configuration Software) that will allow the team to extract detailed dynamic data captured during flight. Flight data will be used to verify the performance of the rocket after flight and will allow the team to make adjustments to the rocket to achieve the rocket s performance objectives. This software will be downloaded prior to flight, allowing the team to become familiar with its functions and capabilities.

21 21 7 Results A preliminary rocket design was completed based upon commercially available rocket designs. These include the nose cone, avionics bay, rocket outer body, motor mount, fins and engine (Figure 1). The rocket fuel propellant was chosen from a formulation commonly used by rocketry hobbyists. Using this fuel the combustion chamber pressure reached the desired value of 1100psia given a factor of safety of 2. The fuel performance results are seen in Figure 8. Fuel Diameter (in) Max Operating Pressure (psi) 1050 Max Thrust (lb f ) 460 Average Thrust (lb f ) 414 Total Impulse (lb f - sec) 780 Burn Time (s) 1.89 Specific Impulse (s) 134 Figure 8: Fuel Performance Using the fuel performance and a conservative coefficient of drag of 0.35, which is greater than that predicted by the CFD analysis, the flight performance was estimated (Figure 9). Peak altitude Z peak = 9593 feet Time to peak altitude t peak = 23.6 sec. Predicted (with Max velocity V max = 964 feet/sec. drag) or V max = 657 MPH Burnout altitude Z bo = 886 feet Figure 9: Predicted Flight Performance

22 22 The rocket nozzle dimensions and properties at the inlet throat and exit are seen below (Figure 10). Additionally no shock formed inside the nozzle because there are no sudden drops in velocity (Figure 5). Rocket Nozzle Data Inlet Throat Exit Diameter (in) Area (in2) Discussion Mach Velocity (ft/s) Pressure (psi) Temperature (F) Gamma mdot (lbm/s) Figure 10: Rocket Nozzle Data The purpose of this project was to design high-powered rocket capable of carrying and deploying a cosmic ray detector at 10,000 feet. This was done by designing the structure, propulsions, aerodynamics, and avionics sections of the rocket. The results focused mainly upon the propulsion and fuel design. Analysis of the fuel resulted in a predicted max altitude of 9593 feet. This was done with a conservative coefficient of drag which was higher than the coefficient of drag predicted by the CFD analysis, so the rocket will exceed the maximum predicted altitude. A conservative coefficient of drag was used due to the inability to determine the accuracy of the program used to perform CFD analysis. Additionally, the CFD analysis was run assuming sea level standard conditions which will not be the case as the rocket s altitude changes. However, a CFD analysis will be run at Mach 1.8 and compared with supersonic wind tunnel testing to determine the accuracy of the CFD program.

23 23 In addition to the wind tunnel testing, future work includes manufacturing the fuel, building the rocket, building a test stand, performing a vibration analysis on the avionics section, determining the force and amount of black powder needed to shear the shear pins, and performing test burns with the combustion chamber and nozzle on the test stand to ensure the accuracy of the thrust calculations, and testing the recovery system.

24 24 9 Appendix 9.1 Additional Figures Figure 11: Dual Deployment Wiring Diagram

25 Figure 12: Avionics Bay Arrangement Sketch 25

26 Figure 13: Rocket Motor 26

27 Figure 14: Rocket Upper Body 27

28 Figure 15: Avionics Bay 28

29 Figure 16: Rocket Lower Body 29

30 9.2 Gantt Chart 30

31 9.3 Budget 31

32 32 10 References [1] W. A. D. Michael T. Heath, "Virtual Prototyping of Solid Propellant Rockets," Computer Science & Engineering, vol. 2, pp , [2] J. Dues, "Avoiding finite element analysis errors," in 113th Annual ASEE Conference and Exposition, 2006, June 18, June 21, 2006, Chicago, IL, United states, 2006, p. Dassault Systemes; HP; Lockheed Martin; IBM; Microsoft; et al. [3] J. F. Dues Jr, "Stress analysis for novices using autodesk inventor," in 2006 ASME International Mechanical Engineering Congress and Exposition, IMECE2006, November 5, November 10, 2006, Chicago, IL, United states, [4] W. Younis, "CHAPTER 9 - The Stress Analysis Environment," in Up and Running with Autodesk Inventor Simulation 2011 (Second edition), W. Younis, Ed., ed Oxford: Butterworth-Heinemann, 2010, pp [5] S. Ping, "Dynamic Modeling and Analysis of a Small Solid Launch Vehicle," in Information Engineering and Computer Science, [6] T. Benson, "Specific Impulse," NASA Glenn Research Center, 12 June [Online]. Available: [Accessed 3 November 2014]. [7] R. Nakka, "KN - Dextrose Propellant," Richard Nakka's Experimental Rocketry Web Site, 15 July [Online]. Available: [Accessed 19 November 2014]. [8] R. Nakka, " KN-Dextrose Propellant Chemistry and Performance Characteristics," Richard Nakka's Experimental Rocketry Web Site, 10 December [Online]. Available: [Accessed 19 November 2014].

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34 34 [17] N. F. Krasnov Ed, V. N. Koshevoy, A. N. Danilov, and V. F. Zakharchenko, "Rocket aerodynamics," [18] J. P. Reding and L. E. Ericsson, "Hammerhead and nose-cylinder-flare aeroelastic stability revisited," Journal of Spacecraft and Rockets, vol. 32(1), pp , [19] A. Fedaravicius, S. Kilikevicius, and A. Survila, "913. Optimization of the rocket's nose and nozzle design parameters in respect to its aerodynamic characteristics," Journal of Vibroengineering, vol. 14(4), pp , [20] J. Morote and G. Liano, "Flight dynamics of unguided rockets with free rolling wrap around tail fins," in 43rd AIAA Aerospace Sciences Meeting and Exhibit, January 10, January 13, 2005, Reno, NV, United states, 2005, pp [21] J. Simmons, A. Deleon, J. Black, E. Swenson, and L. Sauter, "Aeroelastic analysis and optimization of FalconLAUNCH sounding rocket fins," in 47th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition, January 5, January 8, 2009, Orlando, FL, United states, [22] J. T. Ross, M. R. Risbeck, R. J. Simmons, A. J. Lofthouse, and J. T. Black, "Experimental fin tips for reusable launch vehicles (ExFiT) flight data validation," in 52nd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference, April 4, April 7, 2011, Denver, CO, United states, [23] T. Apke. (2014, Nov.1). Black Powder Usage [Online]. Available: