Feedback Questions

1. There is a 3D model that shows a nozzle at the aft end of the motor. Is this accurate?

○ No, there is no nozzle at the aft end of the motor. We corrected that in the 3D model. 2. Can the team discuss the individual dimensions of the launch vehicle?

○ Yes, see the figures below in the PDR addendum section for a complete analysis.3. Can the team clarify what the material and dimensions of the recovery harness are?

○ See bottom for a discussion and complete analysis4. The terminal velocities for both chutes are very slow. Please do a drift analysis for the launch

vehicle.○ See below in the PDR addendum (recovery section).

5. 3 rail buttons are not required for the launch vehicle.○ True, we have limited it to a 2 rail button design, see below.

PDR Addendum

I. AGSE OVERVIEW

Figure 1: AGSE with Launch Vehicle (Down Position)

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Figure 2: AGSE with Launch Vehicle (Measurements)

The AGSE is pictured above with the launch vehicle in place to show the complete system that we will be using for the competition. The AGSE has three distinct systems; the payload system, the launch vehicle erection system, and the ignition insertion system. A more detailed description and explanation follows.

Figure 3a: AGSE with Launch Vehicle (Up Position)

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The picture above shows the launch vehicle position at angle of 85 degrees from the horizontal. The picture also shows how our linear actuator will be extended.

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Figure 3b: Basic electrical diagram showing inside electrical box

Electrical Box: The black box shown in the picture above houses all the electrical power system and the

power sources for the AGSE. Not shown are the terminals connected to the gripper, linear actuator and igniter system. The wires will be be run along the outer rails to each of the component systems, each encased in a shielded wrap. Red will run two wires to the ignition wire system, one wire to the geared servo and a second, shielded again, to the ignition wire. The green wire will run to the linear actuator, and the blue wire to the payload system.

We will be using an Arduino Mega 2560 as a microcontroller Board, along with a Motor shield for controlling the gear motor of the linear actuator , the gear motor of the payload gripper, the two servo-motors of the payload gripper and the two gear motors for the igniter system.

Figure 3c: Arduino Mega 2560 board

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Figure 3d: Arduino Motor/stepper shield from Adafruit

The Arduino will also receive signals from the switches on the platform in order to run the different steps of the system. These switches will be triggered mechanically from each system. The team chose to use trigger switches because of its simplicity and accuracy compared to wireless sensors (i.e. sonar, IR). Further explanations on the use of these sensors can be found in the below sections.

II. PAYLOAD SYSTEM

The payload system is made up of a robotic claw attached to a motorized slide rail and a self securing payload bay that is part of the launch vehicle itself.

Payload Gripper:The team chose this Robotic Claw-MkII from Sparkfun because of its compact size and

usability at a good price. The claw itself is composed of steel with a claw span of about 2 inches. Rubber grips on the claw allow it to more easily retrieve the payload. The schematics of the claw are shown below in Figure 4. Figure 4 also includes how the claw mounts onto the servo motor. Figure 5 shows the specs of the servo motor that the team chose as provided by Sparkfun.

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Figure 4: Robotic Claw

Servo Motor: This is the servo motor (see Figure 5) we have selected as it is designed to work with the other components of the AGSE. The torque of this servo motor is 2.8kg-cm , which is overly sufficient for moving the 4 oz payload and 4.65 oz claw.

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Figure 5: Servo Details

Claw Rotation: The claw will be used to raise the payload from the ground as well as to raise the payload into the launch vehicle. To achieve those motions a second servo motor as detailed above in Figure 5 will be mounted between the claw and the belt. This turn will require the addition of the robotic pan/tilt bracket seen in Figures 6 and 7. Also in Figure 6 you can see the attachment of the claw.

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Figure 6: The Claw Assembly

Figure 7:Pan/tilt bracket for the robotic claw assembly. It provide a pivot point for the robotic claw off of the slide rail described below.

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Belt/Slider:

The belt used to raise the claw from the ground to the launch vehicle is the Actobotics Kit - Channel Slider Kit which is shown in Figure 8. The linear slider has an overall length of 24” which will be extended by ordering extension parts offered by Sparkfun for this slider to have a final overall length of 36”. Each end of the slider has a limit sensor that will alert our microcontroller that the claw has reached either the top or bottom of the slider. Also included in Figure 8 below is the motor used to rotate the belt. It is a precision gear motor, 90 RPM and a 6-12V rating.

Figure 8: Belt Slider

Payload Acquisition System:

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Figures 9 and 10 : Rendering of payload system to the top, general dimensions of payload system to the bottom

III. LAUNCH VEHICLE ERECTION SYSTEM

The linear actuator will be positioned as above in the AGSE. The actuator will be positioned at 30 degrees from the horizontal, or negative x-axis, when the actuator is fully compressed. The actuator is attached to two MB1 brackets (shown below) that allow the actuator

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to change angles during movement to allow easy extension of actuator. A bracket is connected to the AGSE and the launch rail that is wide enough to prevent any damage as the launch vehicle is raised (as shown in Figures 11-3). The linear actuator will extend out until the launch vehicle rail is at 85 degrees. We have two methods to achieve the 5 degree off of the normal as required. We can control the actuator to lift for either a certain distance or for a certain time until the 5 degrees is done. As a backup system, we will place a switch placed upon the base of the AGSE, such that when the guiding rail encounters the switch, it will automatically stop the actuator from continuing its lifting.

Figure 11: Dimensions of the actuator

Linear Actuator Calculations:

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Linear Actuator details:

Figure 12: Rendering of Actuator at full extension (not shown is the ignition wire system)

Linear Actuator Appendix

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Figure 13: MB1 Bracket for the linear actuator

The selected bracket to mount the linear actuator to our frame is the MB1 from the table above. Here is a technical drawing provided by Firgelli Automations and the MB1 attached to the actuator.

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Figure 14: Linear Actuator

We will be using a light duty rod actuator. We will be using the 200 lb model from the table above (FA-240-5-12-XX). Also included are several other of the manufacturer's specifications.

Figure 15: Shown at left are the MB1 brackets as they are to be attached to the linear actuator on each end.

Ignition Insertion:

The idea behind the ignition insertion device is a smaller and much

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less violent version of a pitching machine. Instead of propelling a ball, however, our device propels forward the ignition wire. The device uses a single gear motor and a simple series of gears to rotate two wheels at the same speed. The wheels are separated by a distance of .008 inches, the same distance as the diameter of the 32 gauge ignition wire. The gear motor we have selected runs at 20 rpm. The wire will be spooled in an enclosure to protect it from any damage caused by the launch vehicle. A second benefit to the spooling, is that the only power source connected to the actual wire is the ignition controls, minimizing stray currents from causing an accidental launch. After the wire is fed through the two wheels, it will extend into a funnel, which will guide the wire into the launch vehicle motor.

Figure 16: Ignition insertion assembly

As shown above in Figure 16, the wire is spooled in the apparatus on the very end of the rail. Next is two rubber cylinders that will feed the wire forward. The rubber material also helps minimize the risk of unwanted electrical charges going through the wire. There is an outer steel box to hold the motor, gears, and rubber cylinders together but are hidden for easier viewing (Figure 18 has the full model). The wire then goes through a funnel which directs it into the launch vehicle’s motor.

Gear Motor Specification:

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Figure 17: Gear Motor

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Figure 18: Wire extension assembly

The Wire extension assembly is powered by a 20 RPM geared motor. The motor will spin a smaller gear which will be meshed with a larger, one inch gear that is then meshed with another one inch gear. The reduction will depend on the final speed we want. This will be calculated in a later study. A rubber wheel will be on the same shaft as each of the one inch gears. This will allow the rubber wheels to spin inward or outward together to be able to retract or extend our ignition wire.

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Figure 19: Wire Spool enclosure

The wire spool enclosure will house a spool which the wire is wrapped around. It will be free spinning to allow the wire to unravel easily. The enclosure’s purpose is to protect the wire, and it will be coated in a non-conducting paint to ensure that the wire doesn’t pick up stray electric charge.

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Figure 20: Wire funnel

The wire funnel’s purpose is to guide the wire into the motor. It will be made of steel or aluminum to ensure no damage will be done to it during launch.

Launch Vehicle

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Figure 22: 3-D model of complete launch vehicle (side view)

Length: 1.5 inchesDiameter: 6.155 inchesMass with motors: 26.2 lbCenter of Gravity: 57.579 inchesCenter of Pressure: 68.434 inchesSafety Margin: 1.76

Figure 23: Breakdown of the critical flight and payload systems

I. Launch

Ignition:

The igniter was chosen to produce a suitable energy source to ignite the motor. It is also necessary that the igniter not catch on the motor casting during automated insertion. The

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mechanism for this will be a coil dispenser operated by a motor. If necessary, the wires on the igniter will be reinforced prior to insertion to prevent bending and failure of seating.

Launch rod:

The launch rod is 110 in long and allows for acceleration of 36.8 ft/s stable velocity threshold.

Launch lugs:

The number of launch lugs has been updated from three to two because friction becomes an issue with three lugs. If only two lugs are used there is a reduced friction coefficient allowing for a more rapid acceleration. It is also imperative that since there are fewer lugs that they be more heavily reinforced. This can be achieved by distributing the forces at the body tube-lug joint with washers bent to the curvature of the body tube to prevent indentation and provide a more even contact surface.

The lugs will be placed as follows: one near the center of gravity of the rocket and one near the aft of the motor mount.

Figure 24: Locations of launch lugs with respect to the bottom of the body tube.

Motor mount:

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Figure 25: Locations of centering rings with respect to the bottom of the body tube.

The motor mount will be located at the bottom of the launch vehicle with the bottom of the mount being flush with the bottom of the body tube. The centering rings will be located according to Figure 25.

Motor placement:

The motor will protrude from the motor mount tube one inch to prevent combustible material from impacting the body tube as frequently, therefore helping to preserve the useful lifespan of the launch vehicle.Motor specifications:

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Figure 26: Aerotech K480W Thrust per second

The motor is Aerotech K480W. It is a a 78% K reloadable composite motor and was chosen for its ability to lift the mass of the launch vehicle and payload to a safety margin of 784 feet above the desired apogee (3000 feet). This is to insure that the apogee could still be reached if the launch vehicle has an increased coefficient of drag and/or increased mass relative to the designs and simulations.

The motor was also selected because it is able to accelerate the launch vehicle to the necessary stable velocity for flight. The motor is currently being sold and is is certified until 30 Jun 2015 for hobby rocketry.

Aerodynamics:

The aerodynamics of the launch vehicle are affected by three major factors; nose cone shape, fins shape, and paint finish. The nose cone shape selected for the design was an ogive shape. It was decided that a nose cone with a pointed tip would be easier to construct that would a nose cone with a rounded tip. Even though there is a small aerodynamic sacrifice, the tradeoff is worthwhile because a higher degree of precision can be assured. These decisions were made with the knowledge that a rounded tip is more aerodynamic for subsonic velocities and a pointed tip is more aerodynamic for supersonic velocities.

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Figure 27: Nose cone dimensions (in mm)

The base diameter of the nose cone is the same as the outer diameter of the body tube so that turbulence over the changing surfaces is minimized.

Nose cones tend to work best at a length:diameter ratio of 3:1. The ratio for our nose cone is 4:1. There is nothing wrong with this ratio, but we may consider taking 154.2 mm off the nose cone length to bring it to the 3:1 ratio in future designs. This will serve both to reduce mass and to decrease drag forces.

Fins:

The fins were designed with a wide tab that extends into the body of the launch vehicle. These tabs will be connected to the motor tube. This was the most effective method of reinforcement that we found for the fin structure. The fins must necessarily be reinforced because of the high velocities and torque forces that will be applied to them during flight, especially during the acceleration phase. The wood selected for the fins needs to have a high tensile strength, resistance to bending, and resistance to breaking. A multi-ply wood with perpendicularly oriented layers will provide the most strength. The fins will have a thickness of ½ inch.

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The fins are flush with the bottom of the body tube, but are drawn upwards to prevent damage due to recovery impact and combustible material from the motor.

Figure 28: Fin dimensions

Stability:

Figure 29: Launch vehicle stability diagram

The stability of the launch vehicle depends on the locations of the center of gravity and the center of pressure. The safe margin for stability is between 1.00 and 2.00 body tube diameters. The margin for our launch vehicle is 1.76. This puts it on the higher end of stable, which would theoretically allow for a higher apogee to be obtained.

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Stability diagrams are in the appendix in the bottom of this document.·

Recovery:

● Parachutes

Drogue: 36 inch - ripstop nylon (deploys at apogee)Main: 105 inch -ripstop nylon (deploys at 1000 feet at decent)Payload: 38 inch -ripstop nylon (deploys at 1000 feet at decent)

● Parachute Deployment

At 3000ft the altimeters located in the altimeter bay will set off the rear charge to allow the 36in drogue parachute to deploy to slow the descent down to 64 ft/s. The compartment that contains this parachute is marked on the launch vehicle diagram. As we approach the 1000ft mark during the descent, the altimeter will set off the front charge to separate the altimeter bay and the motor compartment (that are connected) from the nose cone and payload bay. The two parachute sizes used are 105in for the main and 38in for the payload main parachute. These two parachute sizes will give a final descent velocity of 23.3 ft/s for the payload and 17.2 ft/s for the rest of the launch vehicle.

To make sure that the payload parachute is deployed instead of being pushed due to the charge, we will place a loosely fitting nylon bag on the payload chute that encloses most of the harnesses and the chute that is connected by a patch or 2 of velcro by the top half of the bag that is held by a cable or harness to the main parachute of the other half of the rocket. So as the main parachute is pulled out of the body of the main rocket, it will pull out the bag but is loosely connected enough such that the bag will be ripped open and the payload chute pulled out.

● Safety Harness

The components of the safety harness design are a ½ inch thickness multi-ply, cross-layered wood and a ¼ inch stainless steel U-bolt.

● Parachute Attachment

The parachutes will be attached by tubular nylon. This material was chosen due to it’s high tensile strength, durability, and light-weight properties. We will be using ½ inch thickness tubular nylon with a breaking strength of 2000 pounds.

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Figure 30a: overview of launch and parachute deployment

Figure 30b: Bulkhead with parachute harness attachment point

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Figure 31: Bulkhead schematics

● Ejection

Separation of the body tube will be accomplished by the ignition of a black powder charge at each separation point. In order to prevent inadvertent separations anytime during flight and recovery, four shear pins at each separation point will be included to maintain a specific force of separation.

Four small pinholes will be created in the body tube of each separation section to prevent ejection failure due to abnormal pressure differences.

● Electronic control

Electronics tend to be a major failure point on a launch vehicle. For this reason we chose to have redundant power sources for each of our altimeters. This way if one power source fails, at least one altimeter is able to take a reading and control parachute ejections.

Ejection will be controlled by two altimeters. The altimeters should be sealed from any other compartment to prevent erratic pressure readings. In order to guarantee

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precision of the altimeter readings, four small pin-holes will be created in the body tube of the altimeter capsule. This will allow for pressure equalization.

● Tracking

Altimeter: The altimeter that will be used is a StratoLogger Perfectflite Altimeter. Two

outputs are provided for deploying a small chute at apogee to and the two larger chute at the same time at 1,000 foot. The StratoLogger collects flight data (altitude, temperature, and battery voltage) at a rate of 20 samples per second throughout the flight and up to 100,000 foot altitude.

Figure 32: StratoLogger Perfectflite Altimeter

Figure 33: 3-D model of avionics compartment (Precise mounting equipment will be modeled in the next report)

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The PerfectFlite altimeters will be set to fire the parachutes at the appropriate altitude as well as keep track of the altitude that the launch vehicle attains.

See Figure 34 for physical diagrams of the altimeter bay. A crude wiring diagram is shown below. Two altimeters will be utilized to make sure the charges are fired, in case one altimeter fails.

Figure 34: rough circuit diagram for the altimeter bay.

Turbulence can have adverse effects on altimeter readings, so the placement of the altimeter bay is such that there will be minimal wind turbulence from the nose cone and fins.

Radio Recovery:We have a Com-Spec AT-2B radio transmitter that is small enough to fit

easily within our payload electronics bay.

Figure 35: Recovery device schematic

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We will use the paired radio finder to do long distance location of the payload after it has landed. Once we are close enough to see the payload, the paint style should allow for easy identification and recovery.

Launch Vehicle Paint:The paint on the rocket will be a fluorescent with reflective helical stripes

painted over the top. This will make the rocket incredibly easy to find and track visually.

PayloadThe loading of the payload will be achieved using the gripper system

attached to the AGSE and dropped into the bay.

The payload will be brought to the capsule by the AGSE. It will be dropped into the opening on the outer surface area of the capsule, passing through the inner tube, through the locking mechanism, and coming to rest at the center of the chamber - preventing movement during flight. After this point the payload can only be removed by manually pressing down the locking mechanism while the power is off.

Figure 36: Payload compartment (3-D view)

The dropping of the payload will trigger a switch that will cause the microcontroller to rotate the inner payload bay (inner tube) to a secure state. The

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rotation will be between 10° and 45°, in order to keep the payload inside the compartment at normal state of the rocket (horizontal). This is an added security measure in case of locking mechanism failure. It also serves the purpose of reducing the coefficient of drag due to turbulence along the capsule. The rotating mechanism will work by aligning two holes cut specifically for the payload. One hole will be cut into the outer tube, the other cut into the inner tube. The only way for the payload to pass through these holes is for them to be fully aligned. This compartment should not be fully airtight since pressure differentials may cause locking of the moving components.

Figure 37: Payload compartment (rear view)

The rotation of the payload compartment will be done by a servo-motor (same as for the payload gripper as the torque will not be high because the compartment is coaxial with the outer tube). The microcontroller that will command this will be an arduino UNO. We might change and choose a smaller Arduino board as it will only be needed to get a signal from a switch and send a PWM signal to the servo-motor. The power supply will be given by a Nano Lipo of 3S by Turnigy.

· ·

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Appendix

Total motion vs. Time

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Drift Analysis at 5mph

Drift Analysis at 10mph

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Drift Analysis at 15mph

Drift Analysis at 20mph

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Drag Coefficient at 5mph

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