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Washington State University Aerospace Club

Direction for Spaceport America Cup 2019

Noah Thompson, Rocket Team Lead

Matthew Dickson, Control System Electronics Lead

Current Air Brake Design Prototype 


In order to align objectives with the current aerospace industry and establish WSU as a contender on the world intercollegiate rocketry stage, the aerospace club is developing custom simulations and control systems for complex aerodynamic surfaces. The team’s initial design for submission to Spaceport America Cup 2019 included two payloads: a Doppler effect based velocity measurement and an egg. The club has now replaced the doppler payload with an active air brake. The original payloads were auxiliary to the primary goal of the submission, which was to serve as a test platform for the future development of aforementioned control systems. This lead to resources being removed from the original goal of the rocket. In late January the decision was made to remove the placeholder payloads, instead pursuing a simple active braking surface directly. Given a time of only five weeks to produce a functioning prototype for test launch, the group determined that use of an airbrake on this year’s rocket would only be during ascent. The P.I.D. and simulation platform developed for this system will directly apply to later, more ambitious endeavors. With the use of such control surfaces, the club can obtain a level of trajectory control unattainable through other means.


This system was considered to contain primarily only an electrical and mechanical system. The physical structure of the rocket increase (an inherent requirement of the technology’s function). The mechanical components selected also were evaluated on their relative response time, as fast adjustments will need to be made during flight. Lastly, each proposed design’s significance with regard to the current state of astronautical engineering was considered as well. The proposed designs are illustrated in Figure 1.

Figure 1. Considered Air-Brake Design

A weighted case study was performed, comparing each design on the criteria of mechanical simplicity, electrical simplicity, drag capabilities (ability to significantly effect overall drag), Cost, Response, and Engineering Significance (how impressive is it?). Simplicity was weighted heavily due to the time constraints. This study is summarized in the Table 1.

Table 1 – Qualitative trade study of air brake designs.

Air Brake Trade Study
Criteria Weighting #1 #2 #3 #5
Mech. Simplicity 0.3 8 5 6 9
Electr. Simplicity 0.3 7 7 6 9
Drag Ability 0.1 10 10 2 2
Cost 0.1 3 4 4 7
Response Speed 0.1 9 4 8 10
Relevance 0.1 8 6 2 2
TOTALS 7.5 6 5.8 8.1

Ultimately design five, the rotating fin section brake scored the highest in the trade study for its simplicity. However, due to its very low hypothesized drag ability, more in depth analysis of the design was performed.

Two SolidWorks flow studies of design five, one with holes open and one with holes closed, was performed. The total force against flow only increased 3N (from 230N to 233N). These preliminary findings suggest that such a drag modification device might not be significant enough for our purposes. The more traditional air brake was chosen.

Figure 2. Rough conceptual model of design #5, ring air brake

The air brake will be designed to be so electrically and mechanically simple as possible. Almost all internal structures related to it are designed to be cut in two dimensions using a water jet (or a laser cutter for the wooden test rocket components), thus minimizing manufacturing cost and time to production. The mechanism is designed such that a single linear actuator or large servo could engage all brakes by pushing on a single platform.


Figure 3. Base model of design #1, unoptimized, configured for test rocket. Left shown closed, right shown open


Final design and optimization for test rocket will occur from February 15th to 22nd, in tandem with test rocket primary structures fabrication and implementation of control schemes on stationary test bed.


The test rocket will have the centering rings cut from sheet metal, the center platform and linkages 3D printed with solid infill, and the fin supports will be wooden. On the spaceport model, the entirety of the structure will likely be either composite or metal, with the center structure being the only machined component.



The design shown in this document is mostly conceptual (not yet having been optimized). The model shown is designed for a test rocket, flying to an apogee of only 100ft and meeting FAR 101 regulations. This rocket is intended for repeated launch on private land outside pullman, clear of restricted airspace. The data collected in this set of launches will be used to compare four primary frameworks for airbrake control, and two logical methods for approaching onboard data processing. The team’s goal is to minimize the amount of time necessary to run one computational cycle while maximizing altitude control accuracy. These combinations are listed below. Test rocket systems will be run off of a Teensy 3.6.



“Bang-Bang” ControllerSystem has binary states. In the case of the airbrake, if overshoot is detected the brake will extend fully.


P-ControllerSystem state depends on error. The larger the error, the stronger the corrective action.


PI-ControllerSystem state depends on error and duration of error. The larger the error, the stronger the corrective action. In addition, if the error is too small for corrective action but is sustained, proportional corrective action will take place.


PID-ControllerSystem state depends on error, duration of error, and change in error. The larger the error, the stronger the corrective action. Small but sustained error will also result in a proportional corrective action. In addition, corrective action will be taken proportional to the current change in error.


Logical Methods:

Acceleration to AccelerationThe movement of the air brake is determined only by a direction and proximity to desired acceleration. The control system operates only based on acceleration response to movement, without prior knowledge of expected drag.


Acceleration to Brake PositionDesired brake position is determined as a function of needed acceleration to reach apogee. This function is created using the linear interpolation of a two dimensional coefficient of drag matrix. This dataset is generated in terms of velocity and brake position, but reparametrized in terms of average acceleration at a point.




There is very little time to accomplish this goal, and so a strict timeline will be adhered to.

A test rocket will begin launching the weekend of March 9th, weather allowing. Design of the test rocket concludes on February 16th, with construction beginning that same weekend. The base code for the test rocket will be completed prior to the end of February. Ordering of materials for spaceport rocket fabrication will be completed upon approval, with construction beginning as soon as they arrive. The club’s updated layup technique sacrifices a small amount of weight and strength in exchange for dramatically faster turnaround and low rejection rate as compared to previous years. All rocket body tubes (and extras) should be completed within a week of construction beginning.

The full rocket will be completed in time for full scale test launch in late April, the weekend before finals. If less than three club members can be found to attend this launch, it will be postponed until late May.

Rocket and Associated Testing Infrastructure for Use in 2019 Spaceport America Cup Intercollegiate Competition

Washington State University Aerospace Club

Casey Adam Doan, Aerospace Club President

Noah Thompson, Rocket Team Lead

Bryson Jaipean, Rocket Team Vice-Lead



A preliminary design report submitted on January 18th, 2019

Executive Summary

The aerospace club was created to provide students at Washington State University opportunities to develop practical experience and promote collaboration in a multi-discipline engineering team.

The goal for this year is to design and launch a high-powered rocket carrying both an egg payload and a Doppler-effect based velocity measurement device to 10k ft this June as participants in Spaceport America Cup 2019. The project objectives include

  1. Design a centrifuge for large volume egg testing and a custom stability simulation based on MATLAB and SolidWorks.
  2. Create an egg-support structure with neutral buoyancy suspension of the egg aided by a 3D-printed, elastic mesh.
  3. Design and construct a rocket that could deliver both packages to a targeted altitude and safely recover it.



Spaceport America Cup (previously IREC) is an international collegiate sounding rocket competition with over 100 participant teams from all over the world. With a significant scoring emphasis on team dynamics and engineering exhibition, this event serves as an excellent platform to empower and educate our club’s members. The caliber of technology produced for this competition increases every year. Last year, the club launched an L-class motor rocket carry a 3U CubeSat. Unfortunately, due to an unforeseeable circumstance, the custom navigation system on board failed and resulted in the loss of the rocket.

This year WSU Aerospace Club is pushing hard to transition into a more capable, competitive team. In order to achieve this objective, we are using this year’s rocket as a platform to introduce our members to more difficult testing techniques and facilitate the development of custom stability and trajectory simulations. Having learned from the failures of previous years, we are also taking advantage of a newly renewed enthusiasm for rocketry to strengthen our club’s structure and institutionalize the transfer of knowledge from one year to the next. In the following months, we will finalize the design of and begin construction on a high powered rocket for entry into Spaceport America Cup 2019. It will carry both a raw egg payload and a siren which will be used to measure velocity during boost using the Doppler effect.

In order to perform the high volume of testing necessary to properly design an egg support structure (and quantify egg strength), we are designing a centrifuge capable of spinning a 1kg payload up to 23g’s of centrifugal acceleration. Additionally, the Doppler payload requires that there be sound ventilation in the body of the rocket. These small additions, having been determined to invalidate commonly used rocket design platforms, pushes us to develop our own stability and trajectory simulation models.


Overarching Rocket Design Criteria

  • Basic criteria (imposed by competition):
    • Reach an altitude as close to 10k ft as possible
    • Carry at least 8.8lb of payload
    • Utilize a commercial propulsion system
  • Payload Based Criteria:
  • Minimize acceleration experienced (for egg structure)
  • Minimize design introduced probability of flight failure
  • Remain in Subsonic Flight (for doppler structure)
  • Minimize aerodynamic effect of strange external geometries
  • Maximize audibility of Doppler siren
  • Minimize Cost


Weight is, of course, always a consideration in rocketry. Because early designs use a surplus of propulsive power in conjunction with higher than usual weight to keep acceleration low, it is not listed with the other core requirements above. More specific rocket criteria were omitted here for the sake of redundancy, instead discussed in the design section of this report.


Centrifuge Design Criteria

The design criteria for our centrifuge testing apparatus are:

  • Safety
    • Apparatus must be of robust construction
    • Must fail safe where possible
    • Must be remotely controlled
    • Must be transportable (for testing away from populated areas)
    • Catastrophic Failure of Device must be contained or directed in some way.
    • Payload bay must be sealed to fluids, able to withstand sudden dislocation of weight.
  • Consistency & Repeatability of Applied Acceleration
    • Roughly equivalent acceleration must be applied to all points of payload
    • Multiple trials must have approximately same acceleration (repeatable application)
  • Power Supplied to Payload & Data Transmission
    • The payload must have some form of power supply to the payload bay to support measurement devices.
    • Resulting from the need to be remotely controlled, any data collected by the arm must also be stored onboard or transmitted to a ground station at remote control location.
  • Versatility
    • To maximize the value of this expenditure to the club, we would want to design this device in such a way that other payloads could be tested with it.
  • As Always, Minimum Expense Meeting Other Requirements


Current Status of Project 

1/ Overall Design

Though specific external dimensions of the rocket itself are in flux, pending the completion of all internal structures, the approximate length of the rocket will be 100in. The layout of the rocket will be, from fore to aft: nose, drogue parachute, ejection coupler, main parachute, rigid coupler, egg payload bay, electronics bay, Doppler speaker, Doppler vent support structure, motor, tail.

Figure 1. An approximate model of Rocket, made in OpenRocket

We have selected, based on previous experience, the following ejection sequence:

  • E-bay triggers drogue ejection, which is primarily driven by the momentum of separating sections (as opposed to the force from ejection charge detonation).
    • The aft drogue shock cord will be connected to the lower section by an eye-bolt anchored between the coupler charge wells in a 0.5” bulkhead. The fore drogue shock cord will be attached to the coupler section by an eye-bolt anchored between the charge wells.
    • At this point the wiring through the main shock cord has not held an applied load, so risk of flight critical failure of those wires are relatively low. They only need to maintain contact until drogue deployment, after which failure has no effect on flight or data collection.
  • E-bay triggers the main ejection
    • Wiring failure here after ejection does not affect flight performance.


2/ Points of Concern

Main Parachute Connection:

The main chute connection was an immediate point of concern when considering this layout. Usually, the heavier lower section of the rocket is attached to the main chute by a bolt directly into the motor casing forward closure. To ensure that a coupler might be a sufficient replacement for normal attachment, a baseline static study was performed:

Historically, we have used plywood as bulkhead material where possible to save cost. Our preliminary model reflects this design choice. The material applied here is a custom defined generic plywood, with material properties in a direction normal to the surface. This cannot be assumed to give conclusive or accurate numerical values for stresses, but sufficiently functions as an early verification.

Figure 2. Setup & Materials for bulkhead static study.


The coupler end itself is made of ¼ inch plywood, with a lip cut around the edge (making lip thickness of 1/8th inch).

Threaded rods will insert through coupler, with external bolts on either side bracing against the force of ejection. The closures themselves are also epoxied to the outer edge of the coupler tube. The results of a quick simulation are shown below. The surfaces where bolts on threaded rods make contact are fixed, and the edges which will be epoxied are fixed. This fixture on the threaded rods is not completely accurate, as some translation will be allowed before failure. Because of the direction of applied force, this should still serve as a passable approximation. Curvature based mesh with mesh control around applied load and fixtures is used


Figure 3. Static study results

As the results to the in Figure 3 illustrate, all parts of the closure except for the concentrations of stress around sharp edges exhibit less than 24 MPa of Von Mises stress, which should be below the tensile strength for most plywood[1] normal to plane. This is with 247 lbs applied to the surface under the rear eye-bolt fixture. At 15k ft above sea level, density of air is 14.96×10^-4 slugs/cf.

Assuming a drogue coefficient of drag of 1.34 (typical main chute drag) and a solid diameter of 4 ft, the force exerted by the chute instantaneously at 140ft/s can be estimated by:

Therefore, for preliminary projected apogee speeds, the use of ¼” plywood as a chute contact should be sufficient. Since it is near failure at stress concentrations, a thicker bulkhead with composite reinforcement will likely be used and verified during design finalization.


Doppler Payload Acoustic Vents:

Perhaps the component over our design which commands the most attention is our Doppler acoustic venting section. Sandwiched between the motor and the rest of the rocket, cutting holes in a section which is usually incredibly load bearing reduces the overall strength of the rocket dramatically.

By comparing the CFD produced coefficients of drag at variable velocities for identical rockets with and without the vent structure, we determined that by adding the structure we significantly increased drag as well. We used a one-dimensional trajectory program in MATLAB (taking SolidWorks data as input) written by the previous rocket lead, Sean Journot, to verify that 10k feet were still attainable with the addition of such significant drag. The base equations for this code are simply one-dimensional kinematics equations, substituting values on each iteration for differing atmospheric conditions. Sample code is included in Appendix 1.

To allow for these holes to be cut, we are designing an internal support structure to bear the load of that section. It will be fixed rigidly to the body tube below, between, and above the holes. The internal structure forms a parabolic nosecone shape at its core, being tangent to the body tube at vent exit. The structure is being designed to support both thrust and aerodynamic forces in entirety with a safety factor of four. Our own version of this MATLAB simulation is currently being developed, adding a degree of rotational freedom and horizontal translation. When completed, though very computationally expensive, this simulation should be able to more accurately predict stability and trajectory of any ballistic projectile modellable in SolidWorks.

Figure 4 . Lower section of rocket

3/ Centrifuge Design

The decision to attempt to use a centrifuge for egg testing supplements the usual method of launching test rockets. When considering whether to proceed with designing the proposed centrifuge, the following qualitative comparisons were made.


Table 1. Qualitative comparison of test rockets and centrifuge

Test Rocket Pros Test Rocket Cons
·         Less Expensive Upfront

·         Can apply large accelerations for short periods of time

·         Easier to develop and gain approval

·         Easy to balance

·         Testing opportunities are infrequent

·         Has per launch cost that is quite high (motors are guaranteed expense)

·         High risk of rocket failure or loss during testing

·         Applied acceleration cannot be easily fine-tuned

·         Variation between trials would be larger (dependent on more factors)


Centrifuge Pros Centrifuge Cons
·         Very low operating cost

·         Ability to test on a continuous basis

·         Lower risk of device failure or loss (money spent is a greater investment)

·         Opens door to physical acceleration testing of any small structural device

·         Very high consistency of applied acceleration

·         Need to Develop Safety Protocols and Obtain Approval

·         Higher upfront cost

·         Will require proper balancing

·         Will not likely be able to apply large accelerations for short periods of time.


Because of the egg payload, we benefit greatly from the ability to test a large number of eggs with varying structures in different orientations. We plan to examine the responses of different farming methods, weights, sizes, and orientations of eggs under a consistent acceleration to determine those most suitable for launch. Because eggs are so fragile, we also would benefit from the ability to test each possible egg support structure multiple times. We plan to produce a probability distribution of egg failure for each structure considered. The ability to have such a test and revision process, in addition to the return on value in following years with its lower operating cost, leads us to favor the construction of a centrifuge over increasing our use of test rockets.Figure 6. Centrifuge Illustration (not to scale)

The centrifuge arm is selected to be 13 inches in radius. In selection, we wanted a relatively low kinetic energy (easier to contain failure), and relatively low error of applied acceleration (difference in centripetal acceleration over payload). Taking the payload to be 3” in length along the axis of the arm and 25 g’s as the maximum desirable average applied acceleration, the percent error of applied acceleration (using 25g’s as base) can be plotted as a function of centrifuge arm radius. This relationship was used to pick an optimal radius, just below 10% error (deemed an acceptable threshold). Its simple derivation is included in Appendix 2. The remaining centrifuge components are still being modeled but will be completed by mid-February.

The main shaft of the centrifuge will be supported by a thrust bearing and a machine bearing. It will be threaded at the base, with a nut attached during assembly to prevent the shaft from lifting. The arm itself is currently intended to be carbon fiber (its lightweight deemed more beneficial to risk minimization than a more robust but heavier arm). A brushed DC motor will be used to power the shaft, connected by a chain drive.


Centrifuge Electronics

The centrifuge will be controlled by a custom microcontroller with onboard MPU- 6050 accelerometer sensor and strain gauges. The controller will store data on micro SD card as well as relay real-time stream of data to nearby electronic devices via a Wi-Fi access point. The centrifuge will also be equipped with a safety feature to analyze and detect any odd behaviors, and immediately shut down the centrifuge.

An application to control the centrifuge is also in an early stage of development to help multiple electronic devices to establish a 2 ways communication with the centrifuge. The apps will allows the device to obtain data from the centrifuge, and have full control of the centrifuge. User will be able to change centrifuge rotational velocity, turn on and shutdown the machine from a safe distance.  Currently the centrifuge could transmit data to a mobile device with a range up to 200ft. This allows users to establish a safety radius of operation


4/ Gantt Chart