Cornell Rocketry Team: Polaris1500° C. Since the burn duration is less than three seconds, this material choice provides a substantial enough buffer from the measured combustion flame

Experimental Sounding Rocket Association

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Cornell Rocketry Team: Polaris

Team 42 Project Technical Report for the 2019 IREC

Harrison D. Fay1 and Liam G. Patterson2 Cornell Rocketry Team, Ithaca, NY, 14853

and

Sanskriti Joshi3, Jacob H. Mensah4, Yashdeep S. Sahota5, Matthew P. Schneider6, Simran Shinh7, and Christopher R. Vann8

Cornell Rocketry Team, Ithaca, NY, 14853

This report details the system design of the launch vehicle Polaris, created by the Cornell Rocketry Team for entry in the 2019 Spaceport America Cup. The target apogee of the launch vehicle is 10,000 feet above ground level, and the propulsion system is a commercial, solid propellant motor made by Cessaroni Technology Incorporated. Overall, the G12 fiberglass airframe is 147 inches long and 6.17 inches in diameter. A central flight computer carries out most in-flight operations, including GPS tracking, data logging, ignition of primary separation charges, and control of an active airbrake. The launch vehicle separates into two independent sections at apogee, each with their own recovery system. During descent, a guided parafoil is ejected from the forward section and descends towards a target location. A successful mission is completed upon safe recovery of all sections of the launch vehicle and the payload. Also included is the design of a student-developed motor using solid, Potassium-Nitrate Dextrose-Anhydrate propellant. This motor was unable to fly due to manufacturing issues.

Nomenclature σh = hoop stress σmax = tensile strength a = speed of sound AR = fin aspect ratio D = diameter dP = pressure differential g = fin tip to root chord ratio G = shear modulus Isp = specific impulse P = pressure t = thickness Vf = fin flutter velocity

1 Team Lead, Cornell Rocketry Team, 141 Upson Hall Ithaca, NY 14853 2 Team Lead, Cornell Rocketry Team, 141 Upson Hall Ithaca, NY 14853 3 Electrical and Software Lead, Cornell Rocketry Team, 141 Upson Hall Ithaca, NY 14853 4 Airframe Lead, Cornell Rocketry Team, 141 Upson Hall Ithaca, NY 14853 5 Radio Communications Lead, Cornell Rocketry Team, 141 Upson Hall Ithaca, NY 14853 6 Propulsion Lead, Cornell Rocketry Team, 141 Upson Hall Ithaca, NY 14853 7 Business Lead, Cornell Rocketry Team, 141 Upson Hall Ithaca, NY 14853 8 Payload Lead, Cornell Rocketry Team, 141 Upson Hall Ithaca, NY 14853


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I. Introduction HE Cornell Rocketry Team (CRT) is an undergraduate project team at Cornell University, supported by the Sibley School of Mechanical and Aerospace Engineering. CRT was founded in 2013 and has historically competed in

NASA's University Student Launch Initiative. This year marks a transition for the team away from the old competition, switching instead to the Spaceport America Cup.

The team is comprised of 41 Cornell undergraduate team members, a faculty advisor, and a faculty mentor. CRT's faculty advisor is Professor Douglas MacMartin, a Senior Research Associate and Senior Lecturer. As advisor, he provides guidance on general administrative and design decisions, as well as assigning grades to the members that participate in the team for academic credit. CRT’s faculty mentor is Daniel Sheerer, the Building Manager of Weill Hall at Cornell. As mentor, he advises the team on the design and testing of the launch vehicle. He also acts as a liaison between CRT and the Upstate Rocketry Research Group (NAR #765, TRA #139) located at Penn Yan, NY. CRT generally conducts certification launches, sub-scale and full-scale flight tests, and static firings at the launch field.

The team is functionally organized into six subteams: Airframe, Propulsion, Payload, Radio Communications, Electrical and Software, and Business. Projects can be wholly contained within a subteam, but most are shared between multiple subteams. Team members are assigned and committed to one subteam, though they may occasionally do work for another. Each subteam is run by a subteam lead, who is responsible for assigning tasks to their members and ensuring that projects are completed in a timely manner. The team overall is managed by two team leads, who have a variety of roles including scheduling deadlines, organizing recruitment events, and interfacing with the Cornell Administration. CRT receives funding from a variety of sources, including the Cornell College of Engineering and community crowdfunding. CRT also has a number of corporate sponsors that provide funding and gifts in kind. Sponsors from this year include Boeing, Lockheed Martin, Solidworks, Altium, and ANSYS, among others.

II. System Architecture Overview Polaris is a Launch Vehicle (LV) designed for entry into the 2019 Spaceport America Cup. As depicted in Fig. 1,

it is made up of multiple critical systems, namely propulsion, recovery, and payload. The LV uses a solid COTS propulsion system with a target apogee of 10,000 ft Above Ground Level (AGL). At apogee, it separates into two independent sections, each with their own recovery system. The payload is a guided parafoil which ejects from the forward section during descent. All systems must be housed within and integrated with the airframe.

A. Propulsion The propulsion system handles the lift requirements of the LV. Design focus is on lifting the launch vehicle to a prescribed altitude while maintaining control of thrust throughout the ascent. For the majority of the design process, it was intended that Polaris would launch on an Student Researched and Developed (SRAD) motor. However, due to manufacturing issues, full characterization of the motor was not completed in time for the 2019 Spaceport America Cup. Instead, the launch vehicle uses a Comercial Off the Shelf (COTS) motor. The design of both systems is detailed below.

T

Figure 1. External and internal views of the overall layout of Polaris, with labeling of major systems.


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2. SRAD System The SRAD propulsion system consists of a single solid rocket motor using Potassium-Nitrate Dextrose-Anhydrate propellant. Burn duration is 2.75 s with an average thrust of 1234 lbf and an operating pressure of 8.5 MPa. Total impulse is 14585 N � s and specific impulse is 136.9 s. The motor classifies as an N5500. The weight of the loaded motor is just under 55 lbm.

The motor measures 38.86” in length and 4.5” in diameter. As depicted in Fig. 2, propellant is divided into four sections in a Ballistic Test and Evaluation System (BATES) configuration. The motor employs a top ignition system to bring the combustion chamber up to operating pressure before igniting the propellant. Spacers are used to separate grains to ensure smooth combustion progression throughout the burn. The nozzle is made of 1018 cold rolled steel with a melting point of approximately 1500° C. Since the burn duration is less than three seconds, this material choice provides a substantial enough buffer from the measured combustion flame temperature of approximately 1400° C. The nozzle is fixed to the combustion chamber using 18 radial directed socket head screws and sealed using a pair of high temperature o-rings. The nozzle employs a 12° divergent half-angle to account for the high concentration (43% by mass) of solids in the combustion product and a 30° convergent half-angle to facilitate smooth transition of the combustion product to the nozzle throat. Performance predictions were corrected from standard gas models to account for particle thermal lag and particle velocity lag resulting from the solids in the exhaust flow.

The nozzle geometry is an 85% (by divergent length) Rao parabolic approximation that gives 99% efficiency of an ideal nozzle. The nozzle has an expansion ratio of 15.4, optimized for operation around 1500 ft AGL at Spaceport America. This was determined using the theoretical gas De Laval nozzle expansion ratio formula. The throat diameter is 24 mm. Quadratic Bézier curves were developed to model the divergent contour using a set of control points and polynomial coefficients. These points and coefficients were determined using expansion ratio values in Rao’s model for throat exit angles and nozzle exit angles. The nozzle, as with all components of the SRAD motor, was manufactured entirely by CRT team members. Solidworks was used for design modeling and Autodesk HSMworks was used for CAM G-Code CNC programming. Turned parts were made on a Southwestern Industries ProtoTRAK SLX CNC lathe. Milled parts were made on a Haas TM-2P CNC mill with a 4th axis rotary (A-Axis). The motor casing for the SRAD rocket motor must withstand the forces and pressures associated with combustion as well as the thrust induced loads on the airframe. CRT used ANSYS Workbench Static Structural to model material stress concentrations at and around the radially directed bolt holes at both ends of the motor casing. CRT used this model to make material selections and geometry determinations. The motor casing is made

Figure 2. Cutaway view of the CAD model of the SRAD motor, depicting all major components.

Figure 3. The completed nozzle.

Figure 4. ANSYS Static Structural analysis of the motor casing, depicting the stress response under peak loads.


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of a A513 1026 DOM steel tube with a yield stress of 500 MPa. This model gives a factor of safety against yield of 2.16. An insulating liner composed of muli-ply polyurethane impregnated paper protects the inside of the casing from heat transfer during combustion. O-rings on the nozzle and upper bulkhead isolate the radially directed holes in the casing from the high chamber operating pressure.

The ignition system consists of an upper bulkhead with a separate ignition cannister that is fixed to the bulkhead when the motor is ready to be primed for launch. This enables CRT to safely transport a loaded motor to a launch or static firing site. The igniter pre-pressurizes the combustion chamber to 8.5 MPa before igniting propellant grains. This enables propellant to combust most efficiently, optimizing performance. Ignition is initiated by activating an e-match placed inside a pyrogen charge in the ignition cannister.

Propellant is composed of four Potassium-Nitrate Dextrose-Anhydrate BATES grains in a 13:7 ratio by mass. Total propellant mass is 10.9 kg with each grain having a length of 220 mm with an outer diameter of 102 mm and a core diameter of 40 mm. Propellant grains are pressure cured after casting in specifically designed apparatus. Negative erosive burning was a design concern due to the

small port to throat area ratio of the propellant configuration and large length to outer diameter ratio of the motor. However, static firing observations and data demonstrated that negative erosive burning most likely does not occur and if it does occur then its effects are marginal. Grains are primed with a charcoal potassium nitrate paste to facilitate smooth burn transition from grain to grain. The outer surface of the propellant grains are burn inhibited by bonding fiberglass resin impregnated flame proof fabric to it.

The motor was successfully static fired on April 20, 2019. The static firing exhibited no burn anomalies or destructive tendencies. No instrumentation was used in this static firing. A second static firing with instrumentation was planned, but was prevented by inability to use the chemistry lab where propellant was cast. As a result, the SRAD motor cannot be used at this year’s competition. 2. COTS System The COTS propulsion system consists of a single solid motor: the Pro98 13628N5600-P, manufactured by Cessaroni Technology Incorporated (CTI). The motor uses the propellant mixture known as “White Thunder,” with an Isp of 218.4 s and a burntime of 2.42 s. The maximum provided thrust is 1526.2 lbf. The wet and dry masses of the motor are 24.87 lbm and 10.84 lbm respectively. It has an overall diameter of 3.86” (98 mm) and a length of 39.76”

1. The thrust curve of the motor is displayed in Fig. 7. A mounting system for the motor was designed with an emphasis on using fasteners rather than adhesives. This allows the system to be readily swapped in and out in the event that the SRAD propulsion system failed. The mounting primarily consists of a 30” long fiberglass motor tube and two aluminum centering rings. The tube is pressed between the rings, which constrain the tube both axially and radially. The motor is then slid into the entire assembly, with the thrust ring resting against the aft centering ring. A thin aluminum motor retainer is then screwed to the aft ring, locking the motor in place. The mounting system is fixed to the airframe via 16 ¼-20 bolts. These bolts run through the airframe and thread into the centering rings, eight to each ring. This allows for the entire system to

Figure 6. Static firing of the SRAD motor.

Figure 5. Cutaway view of the ignition system.

Figure 7. Thrust curve of the COTS motor1.


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be removed from the airframe if desired. Assuming the maximum load is taken upon only four of the 16 bolts results in a factor of safety of 5.07.

B. Aero-Structures The LV has a length of 147” and a 6” inner diameter and a 6.17” outer diameter. It is made up of two independent sections: the forward section, which consists of the nose cone and payload airframe, and the aft section, which consists of the payload switchband, mid airframe, and the booster section.

All sections of the airframe are made of G12 Fiberglass, except the fins, which are made of G10 Fiberglass. The bulkheads and fin mounts are made of Al 6061-T6. All blast charge canisters are 3D printed to prevent static charge buildup and to allow for rapid replacement in the event of damage. 1. Forward Section The forward section is made up of the nose cone and the payload airframe. The nose cone is a commercially purchased 5.5:1 Von Klarman Tangent Ogive shape, which provides minimum drag. The material is G12 Fiberglass, the same as the other airframe sections. The nose cone holds payload electronics. The payload airframe houses the payload as well as the forward recovery system. It is 35” long and contains a 4.2" x 15" hatch that allows for payload deployment. Two Al 6061-T6 bulkheads (bulkhead 1 and bulkhead 2), on either end of the airframe, isolate the payload from the recovery system in the payload coupler and the nose cone. The payload airframe is fastened to the nose cone coupler via the fasteners system, while it is fastened to the payload switchband via 2-56 nylon shear pins. The aft end of the payload airframe (aft of bulkhead 2) holds the forward recovery system. 2. Aft Section The aft section consists of the payload switchband, mid airframe, avionics bay, and the booster section. It also houses the aft recovery system. The payload switchband provides a tether point for the aft recovery system. This tether point is bulkhead 3. The mid airframe is 19” long, its forward end is connected to the payload switchband via shear pins and its aft end is connected to the avionics bay via the fastener system. The middle airframe houses the aft recovery system. In order to secure airframe sections to one another, a fastener system consisting of internal plastic mounts epoxied to the airframe is used. The fastener system provides a mount for threaded inserts to securely attach to screws within the airframe. The avionics bay contains the central flight computer and one RRC3 board. A 2” switchband on the avionics bay provides a location for two key switches to arm the electronics before launch. Two ¼-20 1018 carbon steel threaded rods hold up the avionics bay mount, and have been chosen based on yield stress calculations. Mounted on the switchband externally is an antenna mount, 3D printed out of nylon and designed to be aerodynamic during flight. Aft of the avionics bay is the booster section which is 58” long and houses the motor, motor mounts, fins and fin mounts. The motor mounts retain the motor and are attached to the airframe via ¼-20 pan-head steel screws. Made of Al 6061-T6, the mounts have been verified to withstand the thrust provided by the motor through ANSYS simulations. Four fins are mounted on the booster section via fin mounts machined out of Al 6061-T6. Each fin is 0.25” thieck and is made from G10 Fiberglass. Fin flutter calculations show that the size and shape of the fins will not undergo flutter during flight.

Figure 8. A schematic of the major structures of the LV.


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3. Material Analysis

One potential failure mode of the fins of the LV is aeroelastic flutter, whereby the passing air vibrates the fins during the flight. At high enough speeds, this vibration can grow significantly in amplitude, eventually damaging or breaking off the fins. The following fin flutter formula was used to verify the selection of G10 fiberglass as the material for the fins.

𝑉𝑓 = 𝑎√𝐺

1.337(𝐴𝑅)3𝑃(𝑔+1)

2(𝐴𝑅+2)(𝑡𝑐)3 (1)

Using the shear modulus for G10 fiberglass and the geometry of the fins, the fin flutter velocity is calculated as 4377 ft/s. With the expected peak velocity of 986 ft/s, this results in a factor of safety of 4.43, which is more than sufficient to justify the selection of G10 fiberglass.

Basic analysis was also performed to ensure that the airframe would not be damaged upon separation due to the internal pressure from blast charges. The following pressure vessel equation was used to calculate hoop stress under a pressure differential of 15 psi.

𝜎ℎ =(𝑑𝑃)𝐷2𝑡

(2) The stress in the airframe was calculated to be 529 psi. The tensile strength of G12 fiberglass is approximately

σmax = 40,000 psi. This results in a factor of safety of 76, meaning that there is virtually no danger of airframe failure as a result of separation charges.

C. Recovery Both the forward and the aft section have independent recovery systems. Each section comes with its own

parachutes, hardware, blast charge canisters, and electronics. The forward recovery system includes a 9 ft nylon parachute and a 2 ft drogue parachute. The aft recovery system includes a 12 ft nylon parachute and a 2 ft drogue parachute. Drogue parachutes are used to reduce the descent velocity of the LV sections to 99.93 ft/s (forward) and 106.37 ft/s (aft) and stabilize the vehicle’s altitude during descent. The drogue parachutes deploy at or near apogee. The main parachutes reduce the descent velocity of the LV to 16.43 ft/s (forward) and 17.49 ft/s (aft). They deploy at 1,000 ft AGL. In both systems, the parachutes (main and drogue) are tied to a single shock cord which is then attached to a bulkhead’s eyebolt via a quick link. The forward system attaches to bulkhead 2 while the aft system attaches to bulkheads 3 and 4. Swivels are also used to prevent tangling and twisting of the shock cord and shroud lines. The main and drogue parachutes for each system are folded and tied together using Jolly Logic chute releases that are set to open at 1000 ft. For redundancy, two jolly logics are used in series and tethered to each other using rubber bands to ensure that the system works if only one activates. The folded parachutes and Jolly Logics are then wrapped in Nomex to protect them from the ejection charges.

Separation is triggered by pressure created by detonating black powder charges. The blast charge canisters that hold the black powder charges are 3D printed to prevent stray static charges from setting off canisters. The caps of the canisters are tethered to the canisters with kevlar wire to prevent them from flying off and becoming a hazard. For redundancy, each system has two black powder charges, one connected to the Central Flight Computer (CFC) and the other to a redundant system. The redundant charge is 1.5 times larger than the primary charge for both systems. For both separation systems, the CFC determines when to trigger blast charges. The CFC directly sends the primary separation signal to the aft separation system. It also sends a signal at apogee to the auxiliary flight controller (located in the forward section). The auxiliary flight controller then directly sends the primary separation signal to the forward separation system. A separate RRC3 altimeter is used for redundancy in both systems (one housed in the forward section, one in the AV bay). The approximate mass of the black powder needed for the detonation and separation of the aft system and the forward system was calculated using MATLAB. The script accounted for friction, the resisting shear pin forces, and the mass of each section in order to calculate the mass of black powder needed. These values, listed below in Table 1, were then verified via ground and flight testing.

Table 1. The amount of black powder in the primary and redundant charges for each section.

Forward Recovery System Aft Recovery System

Primary Redundant Primary Redundant

3.85 g 5.78 g 1.95 g 2.93 g


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D. Payload The Parafoil Experimental Guided Aircraft System (PEGASYS) is comprised of two major systems. The first is the Aircraft Deployment Mechanism (ADM), which includes the Payload Enclosure Zone (PEZ). The second is the paraglider which includes the Parafoil Guidance System (PGS) and the parafoil.

1. Aircraft Deployment Mechanism

The ADM opens a hatch in the side of the rocket, rotates the PEZ so that the end protrudes from the airframe and deploys the payload out of the airframe through the hatch. During launch, three locking pins hold the hatch in place and keep it from opening. During descent, these three pins are pulled, allowing springs to rotate the PEZ out of the rocket. After this the door in the top of the PEZ opens, releasing the PGS. These pins are actuated by servos that are triggered at certain altitudes on descent by electronic boards created by CRT.

The PEZ is the structure retaining the PGS. The PEZ is composed of four acrylic panels, five assorted metal bars, and an aluminum endplate. The dimensions of the PEZ are 10 cm by 10 cm by 30 cm. This conforms with the Cubesat Requirement specified in IREC Rules and Regulations, Section 2.7.1.7, which states that the size of the payload must be comprised of multiples of 10 cm cubes. The PEZ also interfaces with the Payload Deployment Mechanism so that it is held in place.

The four acrylic panels are each roughly 10 cm by 30 cm and 3 mm thick. They have jigsaw edges to facilitate assembly. After being laser cut from stock, the acrylic panels were assembled into place using epoxy. CRT considered various materials for which to construct the PEZ from. CRT considered acrylic, balsa wood, and aluminum for the panels of the PEZ. CRT ultimately chose acrylic because it has a low density, a smooth surface, and can be laser cut. The smooth surface is important for ensuring the PGS exits the PEZ cleanly, and the ability to laser cut the acrylic decreases manufacturing time. Although acrylic is not as strong as aluminum, the acrylic walls are not meant to bear loads.

The four steel bars of two different sizes inside the PEZ are intended to be load-bearing. The larger bar has a 1.5” by 1.5” cross-section, while the smaller bar has a 1.0” by 1.0” cross-section. Both bars have a length of 15 mm. They are of 1215 Carbon steel, chosen for strength and easy machinability. In addition, there is one more T6061-T6 aluminum bar that is designed to align the PGS in place and provide structural support for the PEZ. These five bars have threaded holes so that they can interface with the Payload Deployment Mechanism in various locations. The locations of the 5 metal bars are chosen so that one of the principal axes of inertia of the PEZ is aligned along the roll axis of the rocket. This addresses this issue of instability and imbalance caused by unevenly distributed masses.

The deployment controls’ purpose is to read data from the altimeters, send motor signals to servos and send a radio signal to PGS. Additionally, within this module there is a radio module, whose purpose is to communicate with the controls module and with the PGS radio successfully within range of 200 ft with the PGS.

On power up, the board sends a signal to the servos to push the racks into a locking position, keeping the hatch closed. Additionally, it holds the hatch and PEZ door in place until receiving a signal from the altimeter at a designated altitude and only releases the hatch and PEZ door upon receiving a signal from the altimeter. The launch lock system is actuated using two non-continuous, 180-degree servo motors; two Power HD High-Torque Servo 1501MG. These motors have a stall torque of 17kg-cm or around 1.2 ft-lbs. At a moment arm of .5” these motors can output a force of 28.8 lbs, well exceeding the minimum requirement of force to overcome friction forces in the system, which was estimated to be about 4 lbs.

2. Paraglider

The PGS is the core of the paraglider. The parafoil is mounted to it and it holds the motors and electrical systems that control the parafoil. While no instruments are equipped to the current system, there is extra volume within the system to add instruments like cameras or particle sensors in the future. The PGS is comprised of seven major parts; an aluminum box and top panel, which are secured together using bolts through the top panel that screw into threaded holes in the box. The box provides the primary structure for the system and is robust enough to withstand hard landings.

Figure 9. The full PEGASYS system in a stored configuration


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Both aluminum components were manufactured using a CNC mill. The parafoil is mounted to the top panel, connected at the four corners of the PGS. The break lines run out of slots along the center of the PGS.

Within the PGS are five 3D printed parts that together hold the motors, axles, electronics boards and batteries in place. The parts are placed within the box and held in place by pins protruding from the parts and the top panel. The bottom piece holds the batteries, and the electronics board is mounted to the next. The third piece forms the bottom of the motor mount, the fourth completes the motor mount and forms the bottom of the axle mount, and the top pieces completes the axle mount and provides addition support for attaching the parafoil.

The design of the parafoil was driven by the desire for high lift at low speeds, as these were believed to be the conditions that the parafoil would fly under when making a guided descent from a section of the launch vehicle. This required a thick airfoil, thus the parafoil was modeled to be roughly based on the NACA 4418

airfoil. The airfoil shape was imported into XFOIL and analyses were run under a number of conditions to find its coefficient of lift and drag under varying parameters. The parafoil was then manufactured by sewing laser cut ripstop nylon pieces together. The parafoil was drop tested to verify deployment and gliding capabilities.

The parafoil was tested by varying how the parafoil was folded and conducting drop tests with it in each configuration. The precision and compactness of the folding was found to be most important factor in determining whether the parafoil fully opened up when it was dropped. Through multiple iterations, an optimal folding method was determined. When the parafoil was folded carefully, it consistently exited the PEZ and inflated.

The paraglider controls’ purpose is to collect and record pertinent altitude, speed and heading information, to ensure the controlled flight of the payload. The main sensors were the LSM9DS1 IMU, which contains a 3-axis gyroscope, accelerometer and magnometer, an MPL3115A2 altimeter, which contains a barometer and a thermometer, and a GPS module. These sensors were chosen due to their ability to measure values, since the IMU needed to be able to measure up to 24G and the altimeter need to be able to measure up to 10000 ft. The main controls module is a Broadcom BCM2837, which was chosen over the Atmega328p and the SAMD21 chips due to its four CPU cores versus the other chips’ only one CPU core. Additionally, it had a 1Gb core, which is much larger than the next largest, the 32 Kb RAM on the SAMD21 chip. However, with these benefits comes more power consumption, almost 100 times more than the SAMD21, which only had 0.002 W power consumption. These benefits were important to ensure that the chip would be able to take readings and calculate the required adjustments needed to keep the parafoil on path.

The autonomous flight control systems’ purpose is to keep the parafoil stable in flight, maneuver the paraglider and land the paraglider after mission completion. This system needs sustained paraglider controls throughout the duration of the mission and be able to maneuver the paraglider along a defined flight path as laid out before the mission. There are two servos controlling the brake lines of the parafoil, which was able to control the paraglider’s direction in the air using a PID controller to stabilize the movement about each flight axis. Using the GPS module, this system will direct the payload to the target destination. Depending on its position and velocity relative to the target destination, different trajectories for the payload will be calculated. At the target altitude and velocity, this control system will calculate a Dubin’s path trajectory. This will ensure that the payload arrives as close as possible to the target position and velocity.

The paraglider radio communications’ purpose is to communicate the status of the mission with the ground station and to communicate deployment and mission status with the launch vehicle. The RFM69CW radio is used to transmit packets at least a range of 1 mile away from the ground station and must be able to transmit packets at least 20 m to the launch vehicles. This radio was chosen because of CRT’s extensive experience with the module as well as its ability to transmit up to five miles given line of sight while only using 25mW of power. Due to these specifications, the RFM69CW far exceeds mission requirements, ensuring reliable communications for the duration of the mission.

Figure 10. The PGS, with labels for various openings.


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III. Mission Concept of Operations Overview The mission may be broken down into several stages. The major events that define these stages are depicted in

Fig. 11 below. 1) The LV launches. 2) The forward section is separated and the drougue deploys at apogee. 3) The aft section deploys drogue parachutes one second after apogee. 4) The payload is deployed from the forward section at 1500 ft AGL. 5) The forward and aft main parachutes deploy at 1000 ft AGL. 6) Both sections and the payload land.

Ignition begins with a “fire” signal being sent to the ignitor and concludes when the propulsion system increases to chamber pressure. Lift-off begins at the LV’s first motion and concludes when it has reached the end of the launch rail. Motor burnout occurs, where the LV burns through the powered ascent until the propellants are exhausted. After this, the LV continues to cruise to apogee with steadily decreasing velocity. During this ascent, the airbrake is deployed to increase LV drag to more accurately reach target apogee.

The CFC determines apogee and sends the primary separation signal to the aft separation system and to the auxiliarily flight controller in the forward section. The auxiliary flight controller then directly sends the primary separation signal to the forward separation system. The estimated height of apogee and time taken to reach apogee for varying wind speeds are below:

The CFC will record the altitude of the LV and trigger separation when apogee is detected. A separate RRC3 altimeter is used for redundancy (housed in the forward section and AV Bay). Black powder charges are detonated, creating a pressure buildup that pushes against the bulkheads. This building pressure also breaks removable shear pins used on the compartments of both sets of parachutes, allowing them to deploy. Each separation system has two black powder charges for redundancy (connected to the primary development system and a redundant system).

The forward system is triggered first, followed by the aft system. The forward section is separated, and the forward section drogue deploys at apogee. The aft section then separates the separation section from the booster to create

Figure 11. A diagram of the major stages of the mission.

Table 2. The estimated apogee and time to apogee for varying wind speeds. Wind Speed (mph) 0 5 10 15 20 35 Apogee (ft) 11254 11249 11238 11224 11209 11196 Time to Apogee (s) 26 25.9 25.9 25.9 26 26


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independent sections and to deploy the drogue parachute. The aft separation occurs approximately one second after apogee.

The payload is deployed at 1500 ft AGL, and forward and aft section main parachutes deploy at 1000 ft AGL, as signaled by jolly logics. The sections continue their controlled descent to the ground, until landing. The ground station receives CFC data from the ground station including latitude, longitude, altitude, velocity, and acceleration, allowing for a real time display of LV movement. Since the LV can drift depending on wind speeds, this recovery system allows for the LV to be recovered after landing.

IV. Conclusions and Lessons Learned This year was considered a “transition year” in many respects due to the difficulties in transitioning from the

strictly regimented NASA Student Launch to the open-ended Spaceport America Cup. Some of these challenges were anticipated and some caught the team by surprise. However, this year’s success is measured not by our ranking at competition, but how well the groundwork is laid for future years to build off of what was done before. By that metric, the design process of Polaris has been a tremendous success.

From an administrative perspective, the structure of CRT’s design cycle had to be more or less reworked from the ground up. Schedules needed adjusting to accommodate the later competition. Budgets needed expanding to fund new and more interesting projects. Documentation needed restructuring now that immense, periodic reports were no longer required. Safety needed revisiting as the team looked to mix propellant for the first time. All changes experienced issues as they were transformed, such as the realization that the initial phases of the design cycle need to begin earlier, or learning exactly what incentives were required to ensure that documentation is done correctly and on time.

A number of changes occurred from a technical standpoint as well. With the increase in apogee and the possibility of even higher targets in the future, CRT has realized that a more complete understanding and analysis of the airframe will be required, especially as the maximum velocities approach supersonic. The airframe also switched this year to using aluminum bulkheads attached with fasteners, rather than the typical fiberglass attached with epoxy. However, this resulted in significant issues with electronics, blocking radio and GPS signals. In the future, more care must be taken in the placement of beacons and the layout of metal.

Regarding the payload, something of a cultural difference between NASA Student Launch and Spaceport America Cup is that the former favors complex, mobile, mechanically-focused payloads, while the latter often has static payloads with a focus on analysis. As CRT attempts to find its footing in this new competition, it is preferable to follow the trend of simpler payloads, focusing attentions instead on the LV, at least for the time being.

Finally, the propulsion system was an entirely new endeavor for CRT. In many ways it went remarkably well, with a successful static firing, despite not being able to fly at competition. However, there was a significant knowledge gap between members that had worked with motors on their own time and members that were new to the concept, meaning that a great deal of effort had to be devoted to knowledge transfer. It was difficult to design components that could be manufactured on the machines available to the team, particularly give the size of the motor casing. Besides that, it was discovered very late in the design process, after all the parts were manufactured, that the local launch field did not permit motors to be constructed of steel. This meant that new locations were required for static firing, and a test flight was entirely out of the question.

Despite the challenges, CRT has succeeded in designing and manufacturing a LV for competition. In order to ensure that the lessons learned are passed on, a number of procedures are in effect. At the lowest level, it is the responsibility of each subteam lead next year to teach the new members, guiding them and preparing them to hopefully become leaders themselves. As new subteam leads replace the old, transition documents are passed down with advice and guidance. At the highest level, a massive and comprehensive transition document titled “How to Lead CRT” is maintained by former team leads. The faculty advisor is also able to ensure that the larger mistakes of the past leadership are not repeated.


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Appendix

A. System Weights, Measures, and Performance Data


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B. Project Test Reports


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C. Hazard Analysis


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D. Risk Assessment


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E. Assembly, Preflight, and Launch Checklists


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F. Engineering Drawings


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Acknowledgments The Cornell Rocketry Team would like to thank Cornell University, particularly the Sibley School of Mechanical

and Aerospace Engineering, as well as Mike Thompson, Gibran A el-Sulayman, and Kae-Lynn Wilson. Without their support, our team would never have been able to take the strides we have taken.

We would also like to thank our sponsors, who provide us with much needed funds and tools so that we can

continue to reach new heights. We would like to thank our advisor, Douglas MacMartin, for all of his help with administration this hectic

transition year. We would like to thank our mentor, Dan Sheerer, for all his assistance with design and testing and propulsion. We

would also like to thank the entire group at URRG for supporting us. Lastly, we would like to thank our alumni, those that came before us and laid the groundwork for what we do

today. We hope to continue to build and take on new challenges.

References 1Cessaroni Technology Incoporated, “Pro98 ® high power rocket motor reload kits,” URL:

http://www.pro38.com/products/pro98/motor.php [cited 17 May 2019].

Documents

Cornell Rocketry Team: Polaris1500° C. Since the burn duration is less than three seconds, this material choice provides a substantial enough buffer from the measured combustion flame