Charybdis

Charybdis (ARES-3)
Launch 1 - L1
Launch date	February 6, 2016
Launch site	LUNAR
Launch 2 - L2
Launch date	Pending
Launch site	Pending
Launch 3 - L3
Launch date	Pending
Launch site	Pending
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Team Summary

Stanford SSI Rockets Team - Charybdis,
Leland Stanford Junior University
Stanford, CA
Ian Gomez
Project Manager
iangomez@stanford.edu
Calvin Lin
Team Lead
calvinlin@stanford.edu
William Alvero Koski
Structural Integration Analyst
walverok@stanford.edu
Navjot Singh
De-spin Specialist and Senior Advisor
navjot@stanford.edu
Derek Phillips
Avionics Specialist
djp42@stanford.edu

Launch Vehicle Summary

The Charybdis team under Project Daedalus is in the process of designing a spin-stabilized rocket to launch and reach an apogee of approximately 11,000 feet and dual deploy a main parachute and drogue parachute. The rocket will split between the aft airframe and forward airframe to release the drogue parachute and between the nose cone and forward airframe to release the main parachute.
Charybdis will be launching an L1 test rocket on February 6, 2016 to test the effectiveness of canted fins in inducing a roll in the rocket and the subsequent de-spin prior to reaching apogee at a structurally small-scaled level. Then, the team will be be fulfilling similar objectives in spin-stabilization with the L2 rocket. The overarching goal of the rocket is to build an L3 rocket to test spin-stabilization and de-spin in a large-scaled rocket unit.

Payload Summary

The payload will include an avionics suite to support the dual deploy recovery system and the de-spin unit. The avionics suite will include core electrical components such as the Big Red Bee, Teensy 3.2, and custom altimeters provided by Stanford Student Space Initiative. Miscellaneous items of the payload include a GoPro Hero® and Adafruit with nine degrees of freedom.

Vehicle Criteria

Mission Statement

The objective of Charybdis is to demonstrate spin stabilization and de-spin of a rocket using canted fins and a de-spin unit.

Success Criteria

The mission will be considered successful if the rocket unit reaches spin stabilization, demonstrates successful de-spin prior to reaching apogee. The rocket must be recovered intact and reusable and pass the L3 certification inspection authorized by Tripoli.

Constraints

1. Tripoli height ceiling of 16,800ft
2. Rocket construction to be made using a “minimum of metallic parts” excepting those necessary for airframe integrity
3. Motor impulse to not exceed 10,240 Ns
4. Redundant avionics, wiring, and safe arm systems
5. Vertical descent speed of 20 ft/s maximum upon landing
6. Budget of $4000
7. Back up recovery system with a main parachute
8. Redundant systems
9. A landing within the radius of 300 feet of launching pad

System Overview

Immediately after the rocket is launched, the canted fins will begin to induce a spin in the rocket. At a moment before apogee, the SSI custom altimeter within the rocket will ignite Pyrodex powder (a black powder substitute) to trigger a cord cutter. The cord cutter will activate the de-spin unit to halt the rocket’s angular velocity. This gives a frame of time for the rocket to deploy the drogue parachute. After the rocket descends 5000 ft, the altimeter will cause the main parachute to deploy.
The motor has dimensions of 98 mm in diameter and 548 mm in length. Given the rocket diameter, the motor will permit space for the canted fins to fit in the aft airframe.

Propulsion System

The Cesaroni M1060-P was chosen because it is commercially available, reloadable, complies with the Tripoli and California restrictions, keeps our rocket sub Mach-1, and should achieve a maximum height under the height ceiling with the current mass estimates.

Motor Performance

m0.3 m0.7

[t]0.3

Motor	M1060-P
Type	Reloadable
Diameter
Length
Avg. Thrust
Max Thrust
Total Impulse
Burn Time
ISP

&

[t]0.7

M1060 Thrust curve (lbs. vs. s) image

Flight Characteristics

m0.3 m0.7

[t]0.3

Max Altitude
Max Velocity
Max velocity
Max Acceleration
Wet Mass
Dry Mass
Max Thrust
Center of Gravity
Center of Pressure
Stability Margin
Max Moment of Inertia
Time to apogee
Flight time
Impact velocity

&

[t]0.7

Vertical Velocity curve (ft/s vs. s) image

Altitude curve (ft vs. s)

image

Thrust curve (N vs. s)

image

Mass curve (oz vs. s)

image

Roll Rate Curve (rev/s vs. s)

image

Structural System

Materials

The Charybdis rocket will have various materials, but primarily fiberglass for its resistance to high temperatures. The forward airframe will contain some polycarbonate for the window for the GoPro, but be primarily of fiberglass. The aft airframe, nosecone, and fins will also be fiberglass. The coupler linking the two airframes will be made from kraft phenolic and the bulkheads will be made from both aluminum and wood. The fiberglass was chosen for the body tube as the rocket will reach velocities of 0.92 mach and weaker materials may begin burning up.

Nosecone

The chosen nosecone is the fiberglass nose cone from Public Missiles with a base diameter of 6.1 inches, length of 24 inches, and wall thickness of 0.125 inches. The shape of the fiberglass nose cone will be tangent ogive and have a capped end. Because the end of the nosecone is capped, deployment of the nosecone is much more convenient. The fiberglass material allows the nosecone to withstand up to . Furthermore, the shoulder has a diameter of 5.97 inches and a length of 5.5 inches.

Fins

Three trapezoidal fins constructed from fiberglass with a square cross section and thickness of 0.12 inches will be included in the rocket. The root chord will be , a sweep length of , sweep angle of . Fiberglass is currently the best choice for our rocket because of ease of machining, cheapness, and amplifying stability of the rocket. Because the fins will not extend past the aft end of the rocket, there is no concern in fin damage during landing. After completing the flutter analysis, we can determine and finalize the effectiveness of our shape and dimensions.

Airframe

Both the aft and forward airframe will be long with an inner diameter of and wall thickness of . The airframes both weight . The materials of both airframes will be fiber glass. The aft and forward airframe will be attached with a PT-6.0 coupler from Public Missiles of length .

Spin-Stabilization

Spin-stabilization is the method of stabilizing a launch vehicle by means of spin. Spinning creates angular momentum, which induces resistance to deflecting forces such as wind. The Charybdis rocket will act on a two-axis stabilization.

Canted Fins

To create spin, the three fins on the rocket will each be canted at . Any thing higher in magnitude will induce an incredible amount of spin. Any thing lower will not induce any spin. The fins will be inserted in cut slits at the aft end of the rocket.

De-spin Unit

The de-spin unit will involve a yo-yo de-spin system. Two weights, each attached to a single string, are wound up around the outer side of the rocket. The two weights, each pre-calculated to have a certain mass, will be held within slots on the sides of the rockets. The two weights will be kept snug in the slots with a single string that passes through the airframe. A cordcutter, when charged with Pyrodex, will cut this intersecting string and release the weights.
The reason the yo-yo de-spin was favored is because the unit does not require any calculations of initial angular velocity in order for the de-spin to halt the angular velocity. Rather, the only calculations required are the lengths of the two strings that wind around the airframe and the mass of the weights.

Avionics and Telemetry

In the forward airframe, the rocket will be split into the forward airframe avionics bay, herein referred to as the main avionics bay, and the aft airframe avionics bay, herin referred to as the secondary avionics bay. The main avionics bay will collect sensor information and communicate mission critical parts with the ground station. The unpressurized section is for the altimeter which will deploy the main parachute. The secondary avionics bay will contain an altimeter that controls the de-spin unit and drogue chute.

Avionics Teensy Pinout and XBee Transmitter

In the main avionics bay, a Teensy 3.2 microcontroller will communicate with an XBee 9B 900 Mhz 250MW radio transmitter. The XBee will have a rubberduck RPSMA antenna and communicate information to the ground station in real time. There will be an Adafruit 9DoF sensor providing acceleration, velocity, and orientation information to the Teensy, which will store the data. Parts of the data, such as orientation, that are deemed mission critical will be transmitted through the XBee to the ground station. Everything described above will be connected to a single battery, an Anker Astro E1 5200mAh Portable usb battery pack and have an estimated maximum power consumption of 2 Watts.

APRS Transmission

The second main component to the main avionics bay is a Big Red Bee connected to both an APRS transmit antenna and GPS receive antenna. The purpose of this component is to enable tracking of the rocket at the ground station. This component has its own power supply, and an estimated consumption of 1 Watt.

Camera Payload

A GoPro Session® camera will be loaded directly below the main avionics bay in the forward airframe. The lens will be given a cut window slit on the forward airframe, and covered in polycarbonate in order to prevent any drag. The camera will be self-powered with its own battery, to save consumption from the polymer lithium ion cells.

Altimeters

The two altimeters and their fuction are described in detail in the recovery section.

Power Budget

The Anker Astro E1 Portable usb battery pack provides high energy density of , but is the heaviest component with a weight of . each. This battery will power the Teensy 3.2 and associated components: the XBee Pro 900 RPSMA and the 9 DoF Adafruit. The Big Red Bee and both altimeters are provided with their own power supplies. The total power consumption for the Teensy and associated components is approximately 2 watts. The Teensy uses 5.5V power draw, so the single Anker battery pack contains 28.6 Wh, which can power these components over r14 hours, which far exceeds the time frame of the launch and landing. In the unlikely even that the rocket goes off course, it will continue transmitting its location and other critical information to allow for proper recovery.

Recovery

The Charybdis rocket will employ a dual deploy recovery system, in which the drogue chute is released at apogee and the main parachute is released below apogee.

Drogue Chute System

A Cert-3 L drogue chute from b2 Rocketry will be used to slow the rocket’s descent until the main chute can be deployed. It has a surface area of and a coefficient of drag of . With the parachute the rocket has a calculated descent rate of . The parachute weighs and has a packed size of long with a diameter.

Main Chute System

A Cert-3 XXL main chute from b2 Rocketry will be used as the main chute for the rocket’s descent. It has a surface are of and a coefficient of drag of . With the parachute the rocket has a calculated descent rate of . The parachute weights and has a packed size of long and a diameter of . The main chute will deploy when the rocket is on the descent and is above the ground.

Storage

The main parachute will be stored just behind the nosecone approximately from the front of the forward airframe. The drogue chute will be stored in the aft airframe just behind the coupler, approximately from the top of the aft airframe.

Deployment

Both parachutes will be deployed when the airframes they are in respectively are pressurized by CO2 canisters in the coupler. As the parts separate, the shock cord pulls the rocket out for deployment. One possible idea for preventing tangling of the chute lines due to axial spin is having the hardpoint attachment for the shock cord be able to rotate, to allow the rocket and the chute to rotate independently. At apogee the rocket will be rotating at a speed of . This is before the de-spin unit is activated and is already slow enough that catastrophic failure should not be a problem.

Guidance

The drogue chute will slow the rocket down moderately for the majority of the flight; however, it is not enough to ensure a smooth landing as the rocket needs to come down at a moderate speed so as to prevent the wind from blowing it too far away to retrieve. The main chute deploys at to slow the rocket down to a landing speed, but at a height that it should not be blown too far off course.

Landing

The rocket will land at a speed of ; slow enough to ensure a safe landing so that recovery of the rocket in its entirety is possible.

Manufacturing and Assembly

For small-scaled testing, SSI’s Firestorm kit was used to build the L1 rocket. The only component of the kit not used was the fins. Included in the L1 is the de-spin unit, which includes a cordcutter, two weights, strings, an altimeter, and Pyrodex. The fiberglass fins will be hard cut from a chemical resistance fiber glass of in thickness. For L2 testing, the same process will occur. Any conclusions made from L1 testing will determine if any modifications in manufacturing or assembly will exist.
For the final launch, structural parts will be purchased from trustworthy rocket parts vendors. The fins and airframes will be modified in the Mission control room.

Risk Management

Safety Hazards

The largest concern of Charybdis is the de-spin unit. Because the weights of the yo-yo de-spin must unwind and be released, there is a possibility of inflicted danger. Not only will Charybdis have to perform tests, but also determine numerous methods in which the weights should descend. Some methods include placing mini parachutes on both weights or pack the weights with beads so that the beads separate when released and individually descend. For the L1 test, due to the miniscule mass of the weights (), dropping these weights without attaching a recovery system to them will prove sufficient to fall within a safety range.
The next risk is the de-spin not completely halting the rocket’s spin. If the spin is still significant after the de-spin system is triggered, parachute lines of both the parachute and drogue chute will entangle when deployed.
Another risk is the precessional rate of the rocket. As opposed to inducing rotational velocity, canted fins can also create precessional velocity. If the angle between the vertical of the precessional rate amplifies, it is possible that the rocket loses complete stability. In order to avoid high precessional rate, the fins must be analyzed not only with simulations but also mock launches and physical testing to predict success.
Lastly, weather must be a precaution to determine whether or not Charybdis will launch. If windspeeds exceed on the ground or at , the launch must be postponed.

Budget Risk

The total budget for Charybdis is $4000. Because Charybdis is comfortably under the budget ($2142), the only possible risk to our budget is the expenditures from multiple rounds of testing. Most parts can be reusable, but in case something breaks down, costs will rise.

Activity Plan

Timeline

January 18, 2016	Preliminary Design Review
February 3, 2016	Complete L1 Charybdis Build
February 6, 2016	L1 Charybdis Launch
February 9, 2016	L2 De-spin Building
February 9, 2016	L2 De-spin Test
February 20, 2016	L2 Charybdis Launch
February 24, 2016	Begin L3 Charybdis Assembly
March 1, 2016	L3 De-spin Testing
March 3, 2016	L3 Building Completion
March 5, 2016	Critical Design Review
March 19, 2016	L3 Charybdis Launch