Collection and Analysis of Dynamic Data with an Instrumented Aluminum High Power Rocket

Home , Drogue parachute

UNIVERSITY OF OKLAHOMA

GRADUATE COLLEGE

COLLECTION AND ANALYSIS OF DYNAMIC DATA WITH AN INSTRUMENTED ALUMINUM HIGH POWER ROCKET

A THESIS

SUBMITTED TO THE GRADUATE FACULTY

in partial fulﬁllment of the requirements for the

degree of

MASTER OF SCIENCE

TIMOTHY S. HUNT

Norman, Oklahoma 2005

I would like to thank my wife, Ashley, for her continued support and patience throughout my Master’s coursework. In addition, I would like to thanks my parents, Edie and George, for always making every eﬀort to ensure that I would have the chance to succeed in my educational career, as well as every day life. I greatly appreciate the opportunity to work with Dr. David Miller. Semester after semester, we are exposed to challenges and rewards that far surpass what the everyday graduate student sees. I would like to thank my committee members, Dr. Alfred Striz and Donna Shirley, for their helpful critiques and insight throughout the course of my research and my writing. The enthusiasm brought to this project by our high power rocketeers, Mike Babb and Stewart Ohler, was one of the most encouraging aspects of this entire research project. I enjoyed and appreciated all the real-world information which they provided. I would also like to recognize the Oklahoma Space Industry Development Au- thority for funding this research. Without their initial contribution, this project would have never taken place. Throughout the course of this project, numerous students helped, too many to mention individually. However, I appreciate all of the help that got this rocket oﬀ the ground.

iv Contents

Acknowledgements iv

List Of Tables ix

List Of Figures x

Abstract xii

1 Introduction 1 1.1 Background on High Power Rocketry ...... 4 1.2 Overview of OU SuperSonic Rocket System ...... 5 1.3 Physical Phenomena ...... 10 1.3.1 Supersonic Speeds and the Resulting Shock Wave ...... 10 1.3.2 “Dynamic Overshoot” ...... 11 1.4 Project Management Structure ...... 12 1.5 Organization of Thesis ...... 13 1.5.1 Chapter 2 ...... 13 1.5.2 Chapter 3 ...... 14 1.5.3 Chapter 4 ...... 14 1.5.4 Chapter 5 ...... 15

2 Mechanical System 16 2.1 Rocket System ...... 17 2.1.1 Instrumentation Section ...... 20 2.1.1.1 Nosecone ...... 20 2.1.1.2 Pressure Board Cradle ...... 28

v 2.1.1.3 Accelerometer Cradle / Cable Strain-Relief . . . . . 29 2.1.1.4 Instrument Casing ...... 31 2.1.1.5 DAQ Cradle ...... 35 2.1.1.6 Parachute Plug ...... 37 2.1.2 Altimeter Section ...... 39 2.1.2.1 Altimeter Casing ...... 39 2.1.2.2 Altimeter Cradle ...... 41 2.1.2.3 Parachute Plugs ...... 43 2.1.2.4 Couplers ...... 44 2.1.2.5 Parachute Selection ...... 45 2.1.2.6 Ejection Charge Sizing ...... 48 2.1.3 Booster Section ...... 51 2.1.3.1 Booster Casing Sleeve ...... 51 2.1.3.2 Booster Casing ...... 51 2.1.3.3 Booster Pressure Bulkhead ...... 53 2.1.3.4 Fin-can ...... 54 2.1.3.5 Exhaust Nozzle ...... 55 2.2 Launch Pad ...... 56 2.2.1 Main Structure ...... 57 2.2.2 Rail System ...... 58

3 Electrical System 63 3.1 Telemetry and DAQ ...... 64 3.1.1 Pressure Sensors ...... 64 3.1.2 Acceleration Sensors ...... 68 3.1.3 Strain Sensors ...... 72 3.1.4 Data Loggers ...... 81 3.2 Parachute Deployment ...... 85 3.2.1 Homebuilt Altimeters ...... 88 3.2.2 Commercial Altimeters ...... 93 3.3 Recovery ...... 96 3.4 Printed Circuit Board Design ...... 99 3.5 Software ...... 101

vi 4 System Testing 103 4.1 Pressure ...... 104 4.2 Acceleration ...... 106 4.3 Strain Gages ...... 108 4.4 DAQ System ...... 112 4.5 Parachute Deployment ...... 113 4.6 Transponder Experiments ...... 115

5 Results and Lessons Learned 117 5.1 Results ...... 117 5.2 “Pressure-to-Launch” ...... 119 5.3 Lessons Learned ...... 122 5.3.1 Wind Is Critical ...... 122 5.3.2 Rocket Transponder Use Is An Art Form ...... 123 5.3.3 Put Your Name On It ...... 128 5.3.4 Failure Is An Excellent Learning Tool ...... 129 5.3.5 Idealized Schematics ...... 131 5.3.6 Parachute Color ...... 132

Reference List 132

Appendix A Mechanical Drawings ...... 136

Appendix B DAQ Software ...... 137 B.1 Master DAQ Code ...... 138 B.2 Slave DAQ Code ...... 143

Appendix C Electrical Schematics ...... 149 C.1 Mother-Daughter Schematic ...... 150 C.2 Strain Gage Ampliﬁer Schematic ...... 151 C.3 Pressure Schematic ...... 152 C.4 Accelerometer Schematic ...... 153

vii Appendix D PCB Layouts ...... 154 D.1 Mother-Daughter Board (Top Layer) ...... 155 D.2 Mother-Daughter Board (Bottom Layer) ...... 156 D.3 Strain Gage Ampliﬁer Board (Top Layer) ...... 157 D.4 Strain Gage Ampliﬁer Board (Bottom Layer) ...... 158 D.5 Pressure Board (Top & Bottom Layers) ...... 159 D.6 Accelerometer Board (Top & Bottom Layers) ...... 160

Appendix E Launch Day Log ...... 161

viii List Of Tables

2.1 Isentropic Flow Table for a Gas Having γ = 1.14 ...... 24 2.2 1976 Standard Atmospheric Table ...... 47

3.1 Acceptable Power Density, [19] ...... 75

ix List Of Figures

1.1 Cutaway of Typical Model Rocket ...... 6 1.2 Typical Flight Proﬁle for High Power Rocketry ...... 7 1.3 OU SuperSonic Rocket ...... 8 1.4 Instrumentation Section of the Rocket (1/4 Cutaway) ...... 8 1.5 Altimeter Section of the Rocket (1/4 Cutaway) ...... 9 1.6 Booster Section of the Rocket (1/4 Cutaway) ...... 10

2.1 Nosecone ...... 21 2.2 Pressure Hose Attachment Nipple ...... 23 2.3 Pressure Port Spiral Layout ...... 25 2.4 CNC Setup for Drilling Pressure Ports ...... 27 2.5 Pressure Cradle ...... 28 2.6 Accelerometer Cradle and Data Cable Strain Relief ...... 30 2.7 Half Wheatstone Bridge Arrangement ...... 34 2.8 DAQ Cradle ...... 36 2.9 Parachute Plug ...... 37 2.10 Altimeter Cradle ...... 42 2.11 Adjustable Altimeter End Plate ...... 43 2.12 Parachutes ...... 49 2.13 Booster Upper Bulkhead ...... 54 2.14 Fin Can Assembly ...... 55 2.15 Graphite Exhaust Nozzle ...... 56 2.16 Launch Tower ...... 58 2.17 Attachment of the Radio Tower ...... 59 2.18 Launch Rail (Extruded T-slot Proﬁle) ...... 60

x 2.19 Launch Lugs ...... 61 2.20 Pinch Block ...... 62

3.1 Internal Layout of Absolute Pressure Sensor ...... 65 3.2 Internal Layout of Diﬀerential Pressure Sensor ...... 66 3.3 Variation in Pressure Sensor Porting ...... 67 3.4 Pressure Board ...... 68 3.5 Visual Representation of Motorola’s g-cell ...... 69 3.6 Motorola Accelerometer Package Layout ...... 71 3.7 Accelerometer Board ...... 72 3.8 Strain Gage Selection Chart, [19] ...... 79 3.9 Onset Tattletale Data Logger ...... 82 3.10 Master/Slave DAQ (Seen During Final Testing) ...... 83 3.11 Homebuilt Altimeter ...... 89 3.12 Adept Rocketry Altimeter ...... 94 3.13 Walston Retrieval System (Transmitter Not Shown) ...... 97 3.14 Front Panel of Receiver ...... 98 3.15 Electronics Design Process ...... 100

4.1 Pressure Testing Apparatus ...... 105

5.1 Weather Vaning Flight Proﬁle ...... 118

xi Abstract

The primary motivation for this thesis was to design, construct, and fly an instrumented N-class high power rocket with the attempt to answer some fundamental questions regarding rocket behavior during supersonic flight. In order to conduct this research, a custom rocket system had to be designed for the specific research needs.

This design process culminated with the construction and assembly of numerous custom electronic and mechanical components. The rocket system was then flown on November 29, 2003 at the Sayre Municipal Airport with the intent to explore two scientific phenomena: x The location of the shock wave which results from supersonic flight

y The structural behavior of the rocket when exposed to “dynamic overloading” on liftoﬀ.

These two topics prove to be of interest due to the fact that they are so highly debated in the engineering and rocketry communities. The rocket system had a successful launch but encountered an anomaly during ﬂight which ultimately resulted

xii in a loss of the entire modular rocket system. This thesis concludes with the documentation of the lessons learned from both the launch day events and overall project engineering.

xiii Chapter 1

Introduction

High Power Rocketry can provide many exciting and stimulating opportunities to explore topics in Systems Engineering and Project Management. This thesis covers the design, manufacturing, assembly, and instrumentation of an 11 ft tall minimum- diameter aluminum rocket. This aluminum rocket system was powered by an N-

Class rocket motor and was expected to climb to an altitude of 22,000 ft after reaching a maximum velocity of Mach 1.5. The instrumentation of the rocket system featured a custom sensor suite and a Data Acquisition (DAQ) device; together, these components would be capable of sensing and recording acceleration, strain, and pressure at various points along the instrumentation section of the rocket. This

1 project culminated with the launch of the completed rocket system on November

29, 2003.

Aside from the exciting aspects related to this extreme hobby, high power rockets provide a platform for studying the effects of various phenomena which will be observed during rocket flight. Specifically, this high power rocket system was designed to determine:

x The location of the shock wave which results from supersonic ﬂight

y The structural behavior of the rocket when exposed to “dynamic overloading” on liftoﬀ.

Design, manufacturing, and assembly of the supersonic rocket system required the use of an assortment of skills. Computer Aided Design (CAD) software was utilized extensively in the design of both the mechanical and electrical components.

Individual part designs were then optimized while they were still in the computer environment. In the case of the custom mechanical components, all of the parts were either “CNCed” or manually machined in-house. Space and size limitations inside the rocket dictated that the majority of the electronics would have to be custom fabricated. In a manner similar to the mechanical components, the custom electronics were designed to meet these restrictions using a CAD package specially

2 intended for Printed Circuit Board (PCB) design and layout. These raw PCBs were then manufactured by a supplier. Once the boards arrived from the supplier, they were populated in-house with the necessary electronic components. This iterative design and construction process, both mechanical and electric, requires that great attention is placed not only on part level design but also subsystem and system level integration.

This project also presented challenges in terms of project management. As the conceptual rocket system was reﬁned to echo the emerging design requirements, the scope of the project quickly outgrew the capabilities and knowledge of a single person. This led to the formation of a team which was composed of consultants (high power rocket hobbyists), co-investigators (University Professors), and students (both

Graduate and Undergraduate).

The main goal of this research was to design, construct, and ﬂy an instrumented rocket system which was capable of supersonic ﬂight. The data returned from these

ﬂights would then be used to answer question pertaining to the formation of the shock wave while traveling at supersonic speeds and to the severity of the “dynamic overshoot” on lift oﬀ.

3 The remainder of this chapter will:

x Provide background on high power rocketry

y Present a quick overview of OU SuperSonic Rocket System

z Explain the physical phenomena on which we want to gather data

{ Present the organization of the remainder of this thesis.

1.1 Background on High Power Rocketry

Model rocket hobbyists, as well as high power rocket hobbyists, come together hun- dreds of times each year for both sporting and competitive launch events. These events are usually organized, at the club level, by one of two separate organizations:

The National Association of Rocketry (NAR) [8] or the Tripoli Rocketry Association

[10]. These organizations also regulate the hobby through a set of guidelines and rules which insure that the hobby continues to remain safe and exciting.

High power rocketry is a subset of the model rocket community. This distinction is based upon the size of the rocket motor; model rockets with motors of H-Class or higher (total motor impulse > 160 Ns) are considered to be high power rockets

[8]. Despite the diﬀerence in the various motor sizes, all model rockets are usually constructed in the same manner, refer to Figure 1.1, which was adapted from [9].

4 Three to four fins are securely attached to a body tube. The body tubes are usually composed of glass-fiber, carbon-fiber, or cardboard; of the three types, glass-fiber and cardboard body tubes are the most common due to their low cost. The internal motor mounts, which are usually designed to hold a particular class of rocket motor, are then affixed at the end of the body tube closest to the fins. The opposite free end of the body tube houses the parachute. A plastic nosecone is attached to the body tube and parachute via a shock cord; then, the nosecone simply slides into the body tube. Some hobbyists increase the length of the body tube to create chambers capable of carrying light weight payloads; one common example is an electronic payload containing a video camera and TV transmitter which sends images back during flight [32].

1.2 Overview of OU SuperSonic Rocket System

The rocket system was designed to follow a typical high power rocket flight profile, see Figure 1.2. The first stage of flight is a Powered Ascent. During this stage, the rocket will lift off of the launch pad and accelerate into supersonic speed. After the motor burns out the rocket will coast until it reaches apogee. When the rocket

5 Figure 1.1: Cutaway of Typical Model Rocket senses apogee, it will deploy the ﬁrst of two parachutes, a drogue. This will allow the rocket to make a controlled high speed descent to a lower altitude. Upon reaching a ceiling of 400 ft, the main parachute will be deployed.

Our rocket system was slightly diﬀerent from the normal rockets created by hobbyists; one of the main distinctions was the sheer size of the assembled rocket.

The 10 ft 8 in (3.25 m) tall rocket was only 3.8 in (98 mm) in diameter, giving it a tall slender appearance when on the launch pad. In addition, hobbyists usually use cardboard or glass-ﬁber to construct the body tubes of rockets; however, due to requirements imposed by the use of strain gages, the casing of our rocket was

6 Figure 1.2: Typical Flight Proﬁle for High Power Rocketry constructed from aluminum tubing. The nosecone had a standard parabolic proﬁle; but, instead of being plastic like most manufactured nosecones, ours had been CNC machined from a solid billet of aluminum. Figure 1.3 shows that the outer casing of the rocket has been divided to create three separate sections, respectively from the top down: Instrumentation, Altimeter, and Booster sections.

7 Figure 1.3: OU SuperSonic Rocket

The Instrumentation section, seen in Figure 1.4, housed all of the custom sensors and the DAQ device. These electronics were mounted to cradles which protected them while inside the rocket. The nosecone was rigidly mounted to one end of the casing while a parachute plug was mounted to the other. The entire assembly would then sit atop the altimeter section.

Figure 1.4: Instrumentation Section of the Rocket (1/4 Cutaway)

The Altimeter section, seen in Figure 1.5, had two sliding couplers; these served as the interfaces between the altimeter section and the adjacent instrumentation and booster sections, respectively. Two more parachute plugs were mounted inside the

8 altimeter section. The altimeters, two homebuilt and one commercial, were then mounted to a cradle which spanned the distance between the two internal parachute plugs.

Figure 1.5: Altimeter Section of the Rocket (1/4 Cutaway)

The Booster section, seen in Figure 1.6, was the simplest section of the three.

This section had an upper bulkhead which prevented the motor’s pressure and heat from leaking into the upper sections of the rocket; this pressure bulkhead also served as the parachute plug. The motor slid into the casing from the aft of the rocket and was held in place by the exhaust nozzle and retaining ring. The fin-can assembly, which is composed of a fin-can and four mating fins, is slid down the booster section casing, from the top, and held in place by the threads on the aft end of the booster casing.

9 Figure 1.6: Booster Section of the Rocket (1/4 Cutaway)

1.3 Physical Phenomena

Countless experiments can be performed with both model and high power rockets.

However, two speciﬁc phenomena caught our attention above all others.

1.3.1 Supersonic Speeds and the Resulting Shock Wave

Airfoils moving through air cause tiny disturbance called pressure waves; these waves are the result of the airfoil bumping molecules into one another as it moves through the air [31]. As the airfoil reaches the speed of sound, and faster, these pressure waves combine through a process called constructive interference to produce one strong pressure wave called a shockwave [35]. On a typical airfoil, these shockwaves appear at both the leading and trailing edges of the proﬁle. However, the location

10 of the shockwave on a parabolic nosecone and a slender rocket body continues to be an ongoing dispute amongst high power rocketry enthusiasts. It was the intention of this research to determine the location of the shockwave in respect to the nosecone and shed some insight into its behavior during ﬂight.

1.3.2 “Dynamic Overshoot”

“Dynamic Overshoot” is the transient response of a system which results from a sudden unit-step input. Ali F. AbuTaha uses the example of the oscillating dial of an old bathroom scale to illustrate the commonly misunderstood phenomenon [12].

However, for the purposes of this paper, we will use the example of a rocket system at liftoﬀ. At liftoﬀ, the engines of the rocket come to full throttle within a fraction of a second; this quick throttle-up mimics the required unit-step input which causes this phenomenon. The inertia of the rocket initially resists this sudden force. In the case of rocket systems which are massy, the large force input and the rocket’s inertia then causes excessive compression in the internal structure. This compression generates a potential energy which is capable of subjecting the structure to an additional response on the order of 100% of the original loading [16]. This transient response

11 is present until the system damps out the excess forces and returns to a steady state condition. In the case of a rocket, the extremities and internal structure of the rocket have to have a chance to react and stabilize to the given sudden engine thrust.

These forces then produce dynamic stresses in the system that is experiencing the transient response. All components in these systems must be engineered to account for this excess loading. The sensor suite and DAQ equipment where designed speciﬁcally to catch this phenomenon such that conclusions could be drawn on the severity of the transient response, given a speciﬁc rocket motor thrust.

For reasons addressed in Chapter 5, the fundamental questions we hoped to answer by conducting this research are still unanswered.

1.4 Project Management Structure

Leadership for the project came in two forms: two professors served as the Co -

Principal Investigators on the research grant and two OU Alumni provided guidance through their own high power rocketry experience.

Serving as a Graduate Assistant (GA), I took on the task of detailed management of the team’s day-to-day activities. Project updates were continually discussed

12 with the co-investigators and new action items were generated. In addition, I was responsible for generating top-level constraints for systems, and ﬁnding components that would be utilized to complete the required electronic systems. I also served as the integrator for pieces that were farmed out to undergrad team members. In numerous cases, I walked undergrads through the design process such that they could achieve a better understanding of the overall project goals [17].

1.5 Organization of Thesis

1.5.1 Chapter 2

In Chapter 2, I will discuss the design and construction of the various mechanical systems of both the rocket system and the launch tower. The mechanical systems of the rocket are presented in a top-down manner beginning with the nosecone, which is located in the instrumentation section of the rocket; the presentation then works back to the exhaust nozzle, which is located in the booster section. The latter pages of Chapter 2 present the design methodology behind the launch tower. Throughout

13 the discussion of both the rocket system and the launch tower, part level design considerations are presented as the various components are discussed.

1.5.2 Chapter 3

The presentation of Chapter 3 is very similar to that of Chapter 2. However, the subject matter completely focuses on the assorted electronic systems of the rocket.

The beginning of Chapter 3 discusses the design, selection, and construction of the

DAQ and recovery systems, while the latter pages brieﬂy discuss the development of the software which controlled the DAQ devices.

1.5.3 Chapter 4

The individual systems of the rocket required testing to validate their functionality.

In addition to individual testing, the sub systems also underwent testing to validate their performance as a whole system. These tests are outlined in Chapter 4.

14 1.5.4 Chapter 5

In Chapter 5, I will discuss the events and factors that preceded the launch of the rocket on November 29, 2003. Following this discussion, Chapter 5 will present the results from the launch and the lessons learned from this experiment.

15 Chapter 2

Mechanical System

A great deal of importance was placed upon the mechanical design of both the rocket system and the launch pad. The primary design considerations for both of these systems were derived from the phenomena that we hoped to record during lift-off and supersonic flight. The lift-off event would take place quickly and last for, what we believed to be, durations on the order of fractions of a second. In addition, our experimental platform needs were generating a conceptual rocket system which was anything but small. This presented numerous safety concerns which had to be addressed in order to conduct a successful and safe flight. These abbreviated windows for data collection and obvious safety requirements ultimately created system and part level design requirements with little tolerance for error.

16 2.1 Rocket System

During the conceptual design stage, three key goals were identiﬁed as the guiding criteria for the rocket’s development.

x Minimize the number of parts

y Minimize the turn-around time needed to conduct a second ﬂight

z Construct in a modular manner.

The first goal was to minimize the number of parts. This was important because, as a system becomes complicated with unnecessary components, the respective reliability for the system decreases. Thus, design efforts would focus on the simplification of necessary components and on the minimization of the number of parts used. To further aid in accomplishing this goal, off-the-shelf components were sought after to ease the design load. This was due to the fact that it was realized early on that time spent on design and construction could become costly in terms of money and scheduling; therefore, minimizing the time spent in both design and construction of new components would be beneficial to the project.

The second goal, minimized turn-around time, suggested the use of a re-loadable motor system. Following recovery from the ﬁrst ﬂight, the spent motor shell could

17 be removed and exchanged for another motor which was ready for flight. This would allow the rocket to make multiple test runs within a single test day. Such repetitive sampling should minimize variation in data for a particular phenomenon due to varying weather and environmental conditions over extended launch windows like days, weeks, or months. Another benefit of the re-loadable system is the ability to install various propellants with different thrust vs. time characteristics. This would allow analysis to explore the effects of motor thrusts on the rocket structure.

The third goal, construction in a modular manner, would allow the rocket system to be broken down into pieces and easily transported. A single rocket system, which is 11 feet in length and made entirely from a single section of body tube, can be difficult to transport. However, three rocket sections, each approximately 40 inches in length, could easily be placed in a protective case and hauled/shipped to a given launch site. In addition, the modular design would allow new sections to be added or exchanged with the original system. Different instrumentation sections, equipped with various sensor suites, could be assembled to the rocket for the study of other phenomena; likewise, different booster sections could be used to accommodate a huge assortment of re-loadable rocket motors. In a case where the rocket was a single

18 structure, the instrumentation, altimeters, and other hardware can be buried deep inside the rocket where quick changes would be impossible to implement. However, a module system allows individual sections to be disassembled without aﬀecting the rest of the rocket.

A quick survey of on-line retailers revealed an oﬀ-the-shelf modular rocket system from Dr. Rocket which appeared to have elements that would meet the three primary goals established for the mechanical aspects of the rocket system. Thus, two orders were placed with the retailer to secure the necessary parts for the exterior of the rocket. The initial purchase was for an entire “98 mm re-loadable modular rocket system”. This purchase consisted of eight core parts: Nosecone, Nosecone Bulk- head, Payload Casing, Forward Booster Bulkhead, Rear Booster Bulkhead, Booster

Casing, Fins, and Fin-can. However, as the conceptual design progressed, it was realized that the purchased payload casing was not long enough to produce both the instrumentation and the altimeter casings. For this reason, an additional payload casing was purchased from Dr. Rocket. This purchase ensured that both the instrumentation and altimeter casings could be cut to the required lengths without running out of raw material. The creation of these two casings and various other

19 components began to unveil the ensuing three modular sections: Instrumentation,

Altimeter, and Booster Sections.

2.1.1 Instrumentation Section

The instrumentation section provided a self contained housing for all of the DAQ equipment as well as the test section of the rocket. The main interface between this section and the adjacent altimeter section was via the sliding coupler which was rigidly mounted to the altimeter section. Once the parachutes were deployed the instrumentation section would rely upon Kevlar tethers between the adjacent sections to return it safely to the ground.

2.1.1.1 Nosecone

Inspection of the purchased nosecone revealed that it had been machined from a single piece of aluminum billet. Several diametral measurements were taken at various lengths from the rear base of the nosecone. The resulting values revealed that the cross sectional proﬁle of the nosecone was parabolic in shape [37]. The natural tangential surface of this contour will blend smoothly into the trunk of the rocket. To enhance the transition from nosecone to trunk, our purchased nosecone

20 contained an area which protruded straight back, approximately 6 inches, from the end of the parabolic curvature. This protrusion helps streamline the surface by placing the assembly seam at the juncture of two cylindrical surfaces of the same diameter.

(a) CAD Model (b) Purchased Nosecone

Figure 2.1: Nosecone

Placement of Pressure Ports

To allow for the reading of exterior pressure, the nosecone had to be set up with pressure ports to allow sensors to read the pressure at various places along the outer contour. Each of these ports was composed of two blind holes which intersected each other. Only a single 0.062 inch diameter hole was visible on the outer surface for each port. This outer hole was then met by a larger 0.250 inch diameter blind hole drilled from the inside of the nosecone. These ports would guide air to the individual sensors

21 from small specific spots on the nosecone. This complicated interface was important for the following two reasons. The smaller outer hole would allow us to sense pressure on a small localized area. If a larger hole had been used then the sensed pressure would ultimately reflect the average pressure over a larger area on the surface. It is important to minimize the area in which the reading is taken so that the uncertainty of true pressure at a specific location can be minimized. Ultimately, the reduction of the outer hole diameter increases the accuracy of measuring the true location of the shockwave at supersonic speed. Secondly, the larger inner bore served as a seat into which hose nipples were pressed. Each aluminum nipple was manufactured with an o-ring groove and a seating shoulder; these features were included as a means to ensure a good seal between the hose nipples and the respective inner bores.

The hose nipples also possessed a barb which would prevent the hose from slipping oﬀ. Once a hose was installed onto a hose nipple, the hose was secured to the nipple with a “Zipties”; these “Zipties” were used only to further ensure that the hose would not pull oﬀ of the nipple once the hose was pressurized.

22 (a) CAD Model (b) Hose Nipple Installed in Nosecone

Figure 2.2: Pressure Hose Attachment Nipple

In a case where the density (γ) of a ﬂuid, velocity in terms of Mach (M), and

the ambient pressure (p) are know, the stagnation pressure (pt) at the tip of the nosecone can be found with Equation 2.1 [31].

p γ − 1 γ/(γ−1) = 1 + M 2 (2.1) pt 2

However, the exact flight profile was unknown which in turn makes it difficult to

predict the exact values for these terms. Therefore, the ratio p/pt was simpliﬁed by evaluating the stagnation pressure under a worst case scenario, maximum velocity

at sea level. Table 2.1 gives the ratio p/pt for a given Mach number when the ﬂuid has a density of γ = 1.4. In our case, Mach 1.5 was the absolute maximum expected

23 speed possible. Table 2.1 shows a value of 0.2724 for this speed. For the ambient pressure (p) we again use a maximum condition which is 14.7 psi at sea level. Solving

for the stagnation pressure (pt), Equation 2.1 yields a maximum pressure of 54psi at the tip of the nosecone.

Table 2.1: Isentropic Flow Table for a Gas Having γ = 1.14

The placement of the pressure ports also held some importance. It was uncertain as to what extent an upstream hole would disrupt airﬂow under supersonic ﬂight.

24 Therefore, it was decided that the pressure ports would lie on a spiraling path that tailed oﬀ from the tip of the nosecone. This layout would ensure that the hole at the

Figure 2.3: Pressure Port Spiral Layout

tip of the nosecone was the only upstream disruption for each sensor. An arbitrary distance of 1 inch was measured back from the tip of the nose cone; this marked the location of the 2nd pressure port. In a similar manner, another arbitrary distance of 3 inch was measured from the back shoulder of the nosecone; this was selected so that the 6th diﬀerential pressure port would lie in the cylindrical area of the nosecone. It was also necessary to rotate the 6th port 72◦ about the center axis of

25 the nosecone to maintain the spiral pattern. This rotation angle was based upon the fact that there 5 ports which were to be spaced equally around the circumference of the nosecone; therefore, 360◦ divided by 5 ports is 72◦. After placement of ports 2 and 6, a linear distance along the central axis of the nosecone was measured between the two ports. This distance would help determine the spacing of the remainder of the ports. Simply, the distance remaining was divided by the number of desired ports and the result served as the spacing. This distance and the required rotation of 72◦ allowed us to determine the intersection of a point on the surface of the nosecone for ports 3, 4, and 5. To increase the accuracy of placing the holes in the nosecone, both the nosecone and the desired placement of the ports were modeled in a CAD environment. Once clearances were checked to verify that ports were located in an area where seats could be milled, the nosecone was placed in a HAAS CNC machining center. The CNC machine was instructed to drill the outer 0.062 inch diameter holes in the nosecone. At the same time, an alignment mark was placed on the coupling surface of the nosecone such that inner holes could be aligned easily with the outer holes. Afterwards, using a CAD environment, a drilling jig was designed so that a hand drill could be guided to properly meet each blind hole when aligned with

26 (a) CNC Setup (b) Drilling Cycle

Figure 2.4: CNC Setup for Drilling Pressure Ports the outer alignment marker. The CNC machine was instructed to cut the desired jig from a piece of aluminum rough stock. The jig was placed inside the nosecone and secured with an alignment pin. Given the conﬁning space of the interior of the nosecone, a special 12 inch long 0.250 inch diameter twist drill was purchased. This tool gave us the ability to reach the internal shoulders that were deep inside the cone. The hole depth was carefully measured in the CAD environment, and this distance was marked on the twist drill before each mating blind hole was drilled.

After drilling all the holes, the alignment pin and drill jig were removed.

27 2.1.1.2 Pressure Board Cradle

It was desired that the electronics for sensing the exterior pressure of the nosecone be mounted close to the source of their measurement. This reasoning was based upon the fact that the readings would be taken from pressures ducted by Tygon tubing from speciﬁc ports. To minimize the errors in the pressure readings, the amount of tubing would be minimized. In addition, shortening the tubes reduces the chance of a tube vibrating loose due to oscillations during ﬂight. To achieve this assurance, the pressure sensors would have to be rigidly mounted inside of nosecone.

A fairly simple design was fabricated. The electronics would be securely mounted

(a) CAD Model (b) Installed Pressure Cradle

Figure 2.5: Pressure Cradle

to a thin aluminum plate with nylon bolts. This plate was milled in such a way that

28 excess material was removed leaving a web pattern for stability. In addition, the plate had tabs that protruded from the long edge of the plate. These tabs would then slide into a slot that had been milled into two 0.250 inch diameter rods. This left the electronics mounted to a structure that resembled a stretcher. Using the same pressure port drilling jig, two more holes were added to the first shoulder on the interior of the nosecone. These two holes would hold one end of the rods which supported the aluminum plate. A fourth piece was fabricated to hold the other end of the two rods. This small aluminum fixture slid onto the ends of the two rods and was bolted to the interior of the nosecone with flathead cap screws.

2.1.1.3 Accelerometer Cradle / Cable Strain-Relief

The interior of the nosecone was threaded to accept a bulkhead which was supplied by the manufacturer. The manufacturer’s original design of the rocket had called for the only separation to occur at the nosecone. After the motor burns out and the rocket begins to decelerate, an ejection charge would ﬁre, causing the parachute to deploy by pushing the nosecone oﬀ. The nosecone would then dangle from the

29 rocket and parachute via a Kevlar shock cord. This cord would have been attached to the bulkhead that threaded into the bottom of the nosecone.

Our design called for the nosecone to stay rigidly mounted to the instrument casing at all times. Therefore, we decide to use this bulkhead as the interface for mounting the accelerometer cradle to the instrumentation section. A small alu-

(a) CAD Model (b) Mounted In Place

Figure 2.6: Accelerometer Cradle and Data Cable Strain Relief

minum plate was designed and machined to serve as the secure support for the accelerometers. The accelerometers were mounted to this plate. The combination of the electronics and the plate was then mounted to the bulkhead via 0.250 inch aluminum standoﬀs. The ribbon cable from the pressure sensors was then passed through the bulkhead, and the bulkhead was threaded onto the nosecone. A set

30 screw was used to prevent the bulkhead from backing out while in flight. In addition to providing a secure mounting point, the accelerometer cradle was also designed to help protect the wiring which extended down from the nosecone. Three slots were milled into the accelerometer cradle. The ribbon cables from the electronic sensors were woven through these slots. This prevented lift-off g-forces from pulling the cables loose from their respective devices while in flight.

2.1.1.4 Instrument Casing

As noted before, the instrumentation casing was cut from one of the original purchased payload sections from Dr. Rocket. This section was marked at a length of

27 inches and cut on a horizontal band saw. This method cannot be relied upon when the new cut surface and the outer wall must be orthogonal to each other.

Misplacement of the tube in the horizontal band saw can result in an uneven cut.

Our rocket system required that the cut surface and the wall must be orthogonal to each other. This requirement was based upon the fact that two edges from mating sections would meet when assembled to one another via the couplers. If this assembly seam contains a gap produced by uneven cut lines, then a disruption is created

31 in the smooth outer surface of the rocket. Any disruptions in the airﬂow upstream from the altimeter vent holes can cause false altitude readings by the altimeters.

Therefore, after using the horizontal band saw, the edges of the casing had to be milled true. This was done by milling two wooden jigs in a Haas CNC machine.

Then, without removing the jigs, the casing was placed inside and the edges were trued to the outer wall of the casing.

The nosecone slid into the end of the instrumentation casing. The manufacturer,

Dr. Rocket, had designed this interface to be extremely loose. Therefore, a shim, approximately 0.062 inch thick, was rolled and placed between the inner diameter of the instrumentation casing and the mating surface of the nosecone. Then, the nosecone and the shim were rigidly mounted to the instrumentation casing via 12 ﬂat head cap screws. The holes were drilled in such a way that they created a “zigzag” pattern around the instrumentation casing.

Strain Gage Placement

The instrument casing served as the only surface for which strain data would be taken. Therefore, strain gages would have to be mounted to it. It was unclear, which surface would be best for mounting strain gages, outer or inner surface. Therefore,

32 a trade-off study was conducted to help us find the best solution. The first option required the mounting of the gages on the interior of the instrument casing. This method would shield the strain gages from frictional air flow on the exterior of the rocket and eliminate any false strain readings which may be induced by changes in air pressure over the gages. However, mounting the gages inside the instrument casing proved to be a challenging task. The diameter of the casing was small and left minimal room for one to squeeze in an arm or even instruments for the installation of the gages. This restricted space would prevent the proper application of the strain gages. The assurance of a good bond with the testing surface was crucial to the validity of the recorded data. In addition, it was nearly impossible to install the gages in the desired alignment. Misalignment of gages would have required much testing and calibration before any calculations could be performed on the data collected during flight [22, 23]. Mounting the gages on the exterior of the rocket would allow the gages to be installed properly ensuring their performance.

However, the gages would now be exposed to frictional airﬂow. Therefore, after application of the gages, wax ﬁlm would be placed over the gages and the surface would be protected by rubber electro-plating tape[28, 20]. This method meant that

33 numerous holes would have to be drilled in order to pass the wires from the gages to the DAQ electronics inside the casing. Both the exposure to ﬁctional air ﬂow and the numerous drilled holes were considered to be adequate trades for the assurance of a good gage bond and proper alignment [22, 23, 27, 30]. The gages were laid out in three rings of 8 gages. Each ring would consist of 2 half Wheatstone bridges and one whole Wheatstone bridge. The half Wheatstone bridges would be aligned

90◦ orthogonal to each other, see Figure 2.7, inspired by [18]. Together, the half

Figure 2.7: Half Wheatstone Bridge Arrangement

Wheatstone bridges would capture any axial compression/elongation along with the components of any beam bending in the rocket system. The full Wheatstone bridge was installed as a strain gage rosette. The rosette would capture any strain based upon torque in the rocket system. To ensure the exact placement of each strain

34 gage, each gage and its respective ring placement was modeled in CAD, and the

HAAS CNC machining center was used to drill the through-holes for the wires, which connected the strain gages to the ampliﬁer and DAQ electronics.

2.1.1.5 DAQ Cradle

The DAQ cradle had to house two important pieces of electronics, the DAQ electronics and the strain gage amplifiers. In a manner similar to the other electronic cradles, this cradle was machined from a thin aluminum plate. The plate had 6 thread holes in which standoffs would be placed. Both electronic boards would ride piggy back on the cradle via the threaded standoffs. The DAQ Cradle design called for 6 plastic 9 volt battery holders. Each of these battery holders would simply be bolted to the cradle via 4 threaded holes. However, as testing progressed it was realized that 6 batteries was extremely redundant in terms of power supply and consumption. Therefore, the design was reworked to only include 2 batteries for the onboard DAQ and sensor suite. The mounting location for the battery holders was also moved to the same side of the cradle as the electronics; this change helped provide some extra maneuvering room for when the cradle was being installed into

35 the casing. Extra material was removed from the plate to help reduce the cradle

(a) Original Design (Top (b) Original Design ISO View) (Bottom ISO View)

Figure 2.8: DAQ Cradle

weight. After assembling the electronics to the cradle, the cradle was assembled to the instrument casing at the same time as the nosecone. In order to ease this

36 assembly process, the DAQ cradle was suspended under the accelerometer cradle and loosely mounted via heavy duty “zipties”.

2.1.1.6 Parachute Plug

The purchased parachute and drogue parachute had maximum shock ratings of 100

lbf per line. Both of these parachutes had 16 lines for connecting the canopy to the rocket. Since the parachutes were purchased items, it was decided that all mechanical parts designed and built should exceed a safety factor of 2 based on the forces that would result under the maximum loading of the parachute. The Parachute Plug

(a) CAD Model (b) Raw Machined Part (c) Part with Parachute Ring

Figure 2.9: Parachute Plug

was designed to handle this loading with minimal deﬂection. Overall, the parachute plug was cylindrical with a large chamfer on one end. A single threaded thru hole

37 was machined concentric to the cylinder. This hole would hold a large shouldered eye-bolt. On the backside of the parachute plug and concentric to the threaded through hole was a counterbore which allowed a flat washer, lock washer, and nylon lock nut to securely hold the shouldered eye-bolt without protruding into the cavity housing the DAQ electronics. A web pattern was machined into the bottom of the parachute plug to a depth of 1.00 inch. A matching pattern was milled into the top of the parachute plug however; it was rotated 30◦ about the central axis of the plug. Both pocket milling operations left a solid wall that separated the bottom and top pockets. This wall would prevent any debris from entering the instrumentation section. The parachutes are deployed by using a deflagration charge to suddenly increase the air pressure to a point where it forces the sections to split apart at the coupler. Therefore, an o-ring groove was machined on the outer cylindrical surface to prevent an unnecessary increase in pressure in the instrumentation section of the rocket. Twelve flat head cap screws were used to securely fasten the parachute plug to the instrumentation casing. The hole pattern utilized a “zigzag” layout such that each bolt passed through the outer casing into the wall of the Parachute Plug.

38 2.1.2 Altimeter Section

The altimeter section made up the second modular piece of the rocket system. The primary function of the altimeter section was to house the rocket’s parachutes and their respective deploying devices. In addition, the altimeter section housed the sliding couplers for the rocket system.

2.1.2.1 Altimeter Casing

The altimeter casing was cut from the second payload section purchased from Dr.

Rocket. The length of the casing, 36 5/8 inches, was dictated by the components that were housed inside the altimeter section. In a manner similar to the fabrication of the instrumentation casing, the tube was marked, cut, and machined true. After the location of all internal components had been decided, the center of the area housing the altimeters was marked on the outer surface of the altimeter casing. It was necessary to drill static pressure ports for the altimeters. The standard rule for sizing these holes is based on the internal volume of the area housing the altimeters.

39 This general guideline recommends that a 1/4 inch diameter hole be used for each

100 cubic inches of internal volume housing the altimeters.

! ARef AHole = VChamber (2.2) VRef

Using Equation 2.2, one can quickly determine the size of the static pressure ports needed based upon the previous guideline. For example, our rocket had an internal volume of 120 cubic inches housing the altimeters; this suggests that our housing will require a hole approximately 0.274 inch in diameter. However, this static pressure port doesn’t have to be achieved as a single hole. In the case of our rocket system, four holes were desired such that the chamber could easily respond to quick pressure changes associated with supersonic ﬂight. When multiple holes are utilized, then an equivalent area should be utilized.

AHole = AMultHoles (2.3)

40 Using this guideline and Equation 2.4, it was determined that each of the 4 static pressure ports would have to be 1/8 inch in diameter and equidistant from each other around the circumference of the altimeter casing.

π A = N D2 (2.4) MultipleHoles 4 SingleHole

2.1.2.2 Altimeter Cradle

Our rocket system used a combination of two homebuilt altimeters and a single commercial altimeter. These altimeters all mounted to a single supporting structure.

This cradle was composed of 5 individual parts: 2 round endplates, 2 supporting ribs, and a mounting plate. The mounting plate held the altimeters via threaded standoﬀs. The two homebuilt altimeters were bolted on one side while the commercial altimeter was bolted to the opposite side. Modiﬁcations to the homebuilt altimeters called for an additional circuit board which would drive blinking LEDs showing the altimeters’ current status. This circuit board was mounted at the foot of the homebuilt altimeters. Small slots were milled at both ends of the mounting plate for the supporting ribs. These supporting ribs were placed into the slots and

41 (a) CAD Model

(b) Machined Parts

Figure 2.10: Altimeter Cradle bolted into place. The endplates had slots milled into them that corresponded to the T-shape produced by the assembly of the mounting plate and the supporting ribs. After fitting the endplates to the mounting plate and the supporting ribs, they were affixed using flat head cap screws. The endplates had three additional slots milled into them which allowed a floating grommet to slide in each slot. These slots were curved in such a way that they would allow the entire altimeter cradle to spin

± 30◦ about the concentric axis of the altimeter section.

42 Figure 2.11: Adjustable Altimeter End Plate

2.1.2.3 Parachute Plugs

The altimeter section required two additional Parachute Plugs which were similar in shape to the parachute plug in the instrumentation section of the rocket. The main diﬀerence between the two types of plugs was in their outer diameters. The parachute plug in the instrumentation section was designed to ﬁt snuggly inside the instrumentation casing which had an inner diameter of 3.625 inch. The parachute plugs in the altimeter section were designed to seat inside the couplers which had an inner diameter of 3.375 inch. The altimeter section parachute plugs had one more additional feature. Three small holes were drilled on the backside of each plug.

These holes were threaded and served as the mounting point for the altimeter cradle.

Bolts secured the cradle by passing through the grommets and threading into the

43 parachute plugs. When both parachute plugs were assembled to the altimeter cradle, the curved slots would allow the plugs to rotate a total of ± 60 degrees, with respect to each other, about the concentric axis of the altimeter section. This ability to rotate helped alleviate bolt hole alignment problems between the parachute plugs, instrumentation casing, and the couplers when conducting ﬁnal assembly. The bolts allowing this rotation could be tightened to various levels depending upon the level of resistance that was desired in the twisting motion.

2.1.2.4 Couplers

The couplers served as the mating components for assembling the instrumentation section to the altimeter section and similarly, connecting the altimeter section to the booster section. Mating sections would simply slide onto the couplers, creating a butt joint with the altimeter section. The couplers were machined from purchased aluminum tubing. The only critical dimension was the outer diameter, since they had to ﬁt snuggly into the altimeter section while providing a sliding ﬁt for the mating sections with minimal interference. In order to reduce the risk of a parachute or drogue parachute snagging on an interior surface while being deployed, it was

44 decided that the parachutes would reside entirely inside the couplers. This meant that the couplers would have to extend deep into the altimeter casing. In addition, the drogue parachute and the main parachute were different sizes. This difference meant that the couplers would have to be different lengths in order to accommodate the parachute which would reside in the respective coupler. Due to the long length of the coupler housing the main parachute, it was installed as a permanent item.

This would leave the area housing the altimeters accessible only from the end of the altimeter casing housing the drogue parachute coupler. For this reason, the drogue coupler was machined to be removable from the altimeter casing. When assembling the altimeter section, the parachute plugs and the altimeter cradle would be lowered into the altimeter casing from the drogue parachute end. After securely fastening the parachute plug to the parachute coupler, the drogue coupler was slid into place and fastened with screws just as the parachute plug and coupler.

2.1.2.5 Parachute Selection

The drogue parachute was chosen arbitrarily. One of the consultants happened to have a 40 inch diameter parachute which he commonly used as a main parachute

45 for smaller rockets. In our case, the hemispherical parachute was found to provide adequate drag to suﬃce as a drogue for our projected “dry” rocket mass. The actually descent velocity can be found with Equation 2.5 from [11],

v u u 8 · m · g VDrogueOnly = t 2 (2.5) ρ · Cd · π · DDrogue

where m is the dry mass of the rocket, g is the gravity constant, ρ is the density of

the air, Cd is the coeﬃcient of drag for the respective type of parachute, and D is the

diameter of the parachute. For a hemispherically shaped parachute, Cd is 1.5. The density of air changes approximately 2.777% with every 1000 m change in altitude, see Table 2.2. Since we are primarily concerned with the velocity of the rocket when it touches the ground, we use the density at launch elevation, ρ is approximately

1.22 kg/m3. Using a “dry” mass of 22 kg and a parachute diameter of 40 inches, the equation yields a descent velocity of 56.9 ft/sec.

The main parachute was sized based on the desired descent velocity of 15 ft/sec, respectively 4.57 m/sec. Using Equation 2.6, from [11], in conjunction with the

46 Table 2.2: 1976 Standard Atmospheric Table constants stated above, a hemispherical parachute with a diameter of 149 inch would be required to slow the rocket to the desired descent velocity.

s 8 · m · g DChute = 2 (2.6) ρ · Cd · π · VDescent

47 This posed a slight problem considering that the largest parachute available from retailers was only 120 inches in diameter. To evaluate the eﬀectiveness of this parachute we use a slightly modiﬁed version of Equation 2.5.

v u 8 · m · g u VBothChutes = (2.7) t 2 2 ρ · Cd · π DDrogue + DMain

The main diﬀerence between the two equations is that Equation 2.7 takes into account that the drogue parachute is still deployed. Therefore, when using both a 40 inch drogue parachute in conjunction with a 120 inch main parachute, the rocket will have a descent velocity of 18 ft/sec. This deviation from the desired descent velocity was discussed, and it was decided to be an acceptable alternative to ﬁnding a custom sized parachute.

2.1.2.6 Ejection Charge Sizing

The deﬂagration charges used to separate the rocket sections were sized according to two factors: the volume of the respective parachute chamber and the desired

force of separation. High power rocketry enthusiasts recommend 150 - 200 lbf as the separation force for rockets with an internal casing diameter of 3.375 inch [11, 3].

48 (a) Drogue Parachute (b) Main Parachute

Figure 2.12: Parachutes

Using Equation 2.8, this would suggest that a pressure change on the order of 16.8

- 22.4 psi would be required.

4 · F P = 2 (2.8) π · DCasing

For simplicity sake, an even value of 20 psi was arbitrarily chosen between these upper and lower bounds. Equation 2.10 takes into account a constant C which is

49 derived from the properties of FFFFg (4F) black powder. This constant has been simpliﬁed to the following equation.

C = 0.0004 · PDesired psi (2.9)

Replacing P with 20 psi gives a C constant of 0.008. Finally, the internal casing diameter and the respective length of both the drogue and main parachute chambers are combined with the C constant in Equation 2.10.

Bp = C · D2 · L (2.10)

The drogue parachute chamber was approximately 9.875 inches in length and required a 1.0 gram charge of black powder. The main parachute chamber was substantially longer at 35 inches and required a 3.19 gram charge of black powder for separation.

50 2.1.3 Booster Section

The third modular component was the booster section. The primary function of the booster section is to house the motor which provides thrust once the propellant is ignited.

2.1.3.1 Booster Casing Sleeve

The remaining piece of purchased payload section, which was leftover after cutting the necessary piece for the altimeter casing, was cut to form the booster casing sleeve. The sleeve was 11 inches long and threaded internally on one end. The threads fastened the sleeve to the booster casing. When in place, the booster casing sleeve and the coupler on the altimeter section served as the mating components which held the booster section and the altimeter section of the rocket together until a separation charge was ﬁred.

2.1.3.2 Booster Casing

The original purchased rocket system came with a stock N-class re-loadable motor casing. This would have suited our needs quite adequately. However, a problem arose when the only commercial facility in the United States producing these motors

51 caught ﬁre. This made it impossible to ﬁnd a commercial motor since the majority of the retailers dealing in this class of experimental motors generally order only what they need instead of readily stocking such an item. This development forced us to

ﬁnd an experienced rocketeer who could mix a custom N-class motor to meet our needs. It was known that our third party motor vendor could mix a motor to the physical size that we needed, however, it was unclear if the vendor could actually produce a motor that would be within the maximum pressure ratings for our original purchased booster casing. Therefore, to reduce the risk of a CATO (Catastrophic failure on take oﬀ) due to motor overpressure and booster casing failure, our motor vendor was asked to produce both the N-class motor and a casing which would handle his experimental motor pressures. The introduction of a new booster casing caused many of the parts which accompanied our original casing to become obsolete. A new booster bulkhead, exhaust nozzle, and relief ring would all have to be designed to match the new motor and booster casing. The new booster casing also had two threaded surfaces, one at the top of the casing and one at the aft end of the casing.

The threads at the top of the casing served as the mounting point for the booster sleeve. The sleeve would then slide over the parachute coupler when assembling the

52 rocket on the launch pad. The threads at the aft end of the casing securely fastened the ﬁn-can to the booster casing.

2.1.3.3 Booster Pressure Bulkhead

At the top of the booster casing, a bulkhead prevented the internal motor pressure for leaking into the trunk of the rocket system. An increase in pressure in the trunk of the rocket could cause the rocket to separate prematurely and deploy the main parachute. In addition, the gases internal to the booster casing are very hot; leaking these gases into the trunk could also destroy electronics and parachutes.

This bulkhead was machined from a cylindrical piece of aluminum round stock. The outer diameter was the most important dimension of the part. It was designed to

ﬁt snuggly into the booster casing. A large O-ring groove was placed in the outer diameter of the bulkhead to minimize any chance of pressure leaking by the bulkhead.

The bulkhead was held in place by a stiﬀ internal E-Clip ring. Once the E-Clip ring was installed into the booster casing, the bulkhead was securely assembled to the booster casing. The bulkhead also served as the mounting point for the tether which tied the booster section to the altimeter section upon separation. A 1/2 - 16 threaded

53 (a) Original Design (b) Original Design (c) Simpliﬁed Final De- (Cavity for Smoke Trail sign Cartridge)

Figure 2.13: Booster Upper Bulkhead hole was machined concentric to the center axis of the bulkhead. A shouldered eye bolt was then threaded into the hole such that the tether could be securely fastened.

2.1.3.4 Fin-can

The fin-can was machined from a thin walled aluminum pipe. The top of the fin-can was chamfered with a steep slope to help alleviate any disruption in air flow over the

fins during flight. The fins slid into dove tail joints which were spread equidistant around the circumference of the fin can. Each of these joints was accessible from the aft of the enclosure. Once each of the fins had been fully inserted, they were secured with two set screws. The fin can sub assembly slid down over the booster casing and fastened to the threads at the aft end of the booster casing.

54 (a) Individual Fin (b) Typical Dovetail (c) Fin Can Assembly Joint

Figure 2.14: Fin Can Assembly

2.1.3.5 Exhaust Nozzle

Internal booster pressure and external thrust are directly dependant upon the geometry of the exhaust nozzle. Small changes in the conical angles that make up the nozzle’s inlet and outlet can greatly change the pressure [15]. In addition, the nozzle will behave diﬀerently when subjected to motors that have variations in their performance. Therefore, we elected to have this part made for us by the same vendor who was mixing our custom motor. Exhaust nozzles also experience wear induced by high velocity exhaust gases. To reduce this wear, the exhaust nozzle was machined from a solid piece of graphite. Much like the bulkhead, it was critical that the nozzle

ﬁt snuggly into the booster casing. The exhaust nozzle also used an external o-ring,

55 (a) Exhaust Nozzle (b) Cutaway Show- ing Convergent and Divergent Zones

Figure 2.15: Graphite Exhaust Nozzle the same size as the o-ring used in the bulkhead, to seal itself to the booster casing.

The nozzle also had an additional notch machined into the aft end of the part. This notch allowed an aluminum ring to be ﬁtted ﬂush to the outer diameter and the aft surface. The ring would bear the stresses of the internal motor pressure and prevent any unwanted chipping or cracking of the nozzle upon motor operation. Both, the nozzle and the relief ring were held in place by an addition internal e-clip ring.

2.2 Launch Pad

The physical size and weight of the rocket, along with the thrust generated by the booster on lift oﬀ, dictated that the launch pad would have to be a very stable

56 structure to ensure a safe launch. We searched the internet for an “out-of-the-box” solution to our launch pad needs. This search left us with two hard truths. First,

“out-of-the-box” launch pads were only produced for rocket systems much smaller than our system. Secondly, there was no one single method, which high power rocketry hobbyists adhered to when constructing their custom launch pads. However, this vast variation allowed us to design a launch pad for our rocket which was loosely based upon the positive design elements of the numerous solutions implemented by other enthusiasts.

2.2.1 Main Structure

While studying various designs, it was noted that larger rocket systems utilized large frameworks to serve as the main structure of the launch pads. However, we simply did not have the time and resources to develop such a custom framed structure. To solve this issue, we decided to use scaffolding as the main structure for our launch pad. This provided a stable and economical solution to our structural needs. The scaffolding also fulfilled our need for a main structure that was portable and easy to erect or disassemble. In addition, the scaffolding can be found at almost any rental

57 outlet, making it a readably available resource. Given a launch site far from home,

Figure 2.16: Launch Tower

the rocket and launch rail system could be transported to the launch site where it could be assembled with scaﬀolding from a rental shop in a near-by town. This would cut down on the number of large parts that the team would have to transport to the launch site.

2.2.2 Rail System

The rail system was designed with the intent that it could be attached to any available scaﬀolding. The main backbone of the rail system was two pieces of triangular radio tower. When these pieces were joined end to end, the total tower length was

58 20 feet. Using eight U-bolts, the tower was bolted to one end of the scaﬀolding at

Figure 2.17: Attachment of the Radio Tower

a height of approximately 3 feet off the ground. The attachment of the radio tower to the scaffolding greatly increased the overall stability of the launch pad. When studying the launch pads utilized by other hobbyists, most used a rod or a grooved rail to guide the rocket on-lift off. Considering the mass of our rocket, we quickly ruled out the use of a rod. The rod would have to have a fairly large diameter to guide our heavy rocket. This would become a problem when attaching the corresponding launch lug for a large diameter rod to the exterior of the rocket. We required surface protrusion along the trunk of the rocket to be as minimal as possible; a guiding launch lug for a large diameter rod would protrude from the trunk of the rocket and would increase the drag on the rocket; in extreme cases, the parasitic drag could potentially prevent the rocket from reaching supersonic speeds while in

59 flight. Therefore, we opted to use a grooved rail system instead. This would allow us to integrate a small t-shaped launch lug which would fit snuggly to the trunk of the rocket. Upon lift-off, the t-shaped lug would slide inside of a grooved channel on the rail which would be affixed to the launch tower. In case a stronger launch lug was needed, the lug simply would be made longer. Longer launch lugs would gain strength by extending further down the trunk of the rocket instead of extending further out from the trunk as in the case of the rod systems. To achieve this design, we purchased three grooved rails and the corresponding launch lugs, which would attach to our rocket trunk and slide within the grooved rail. However, the

Figure 2.18: Launch Rail (Extruded T-slot Proﬁle)

purchased launch lugs would only work in one of two areas along the trunk where launch lugs were needed. The use of a ﬁn can, as compared to the attachment of the

ﬁns directly to the booster casing, meant that the two required launch lugs would

60 not be attaching to the same cylindrical surface. This required that one launch lug would have to be slightly taller than the purchased lugs. To our surprise, taller launch lugs were only produced for casings the next size smaller than our casing and less. This meant that we would have to design and machine a launch lug that

(a) Standard ACME Launch Lug (b) Modiﬁed ACME Launch Lug

Figure 2.19: Launch Lugs

could be attached in the desired location and protrude out to the same plane as the purchased component when it was attached to the ﬁn can. To accommodate these launch lugs, the three grooved rails were aﬃxed end to end and were attached to the protruding edge of the radio tower. Four pinch blocks were designed and machined to rigidly hold the grooved rail to the radio tower. These pinch blocks attached to the back side of the grooved rail with socket head cap screws and sliding t-nuts.

Then, the pinch blocks would wrap around and clamp to the outer tube of the radio tower using two additional socket head cap screws.

61 Figure 2.20: Pinch Block

62 Chapter 3

Electrical System

The electrical system of most high power rockets is comprised of an electronic altimeter and a recovery device of some nature. Occasionally, hobbyists will add a video camera to their rocket; however, one must remember that the true thrill of this sport comes from flying the rocket as compared to just developing the electronic gizmos for it. This limited need for electronic extras has placed a low demand on new developments in electronics, designed specifically for high power rocketry. However, the electrical system for this rocket system would be substantially different from that of the average hobbyist; it would require an elaborate mixture of altimeters, recovery transponders, telemetry electronics, and DAQ.

63 3.1 Telemetry and DAQ

A signiﬁcant amount of time was spent searching for “oﬀ-the-shelf” telemetry and

DAQ components. However, most of the components which were small enough to be housed inside the rocket exceeded the budget of this project. To compound the problem, most components, which were aﬀordable, were also designed for in- dustrial applications. The design requirements of these applications usually allow bulking/awkward packaging or additional features which our DAQ scheme did not tolerate. These problems spawned a need for custom ﬂight hardware, which was designed for the individual requirements of each separate system.

3.1.1 Pressure Sensors

Given the physical layout of the nosecone assembly and the required proximity of the sensing electronics, it was decided that the use of two types of pressure sensors, diﬀerential and absolute, would best suit the geometry of the ports on the nosecone.

An alumnus, who was helping with the development of the rocket system, suggested that the team look at Motorola Semiconductors for the needed sensors. It was found that Motorola oﬀered a wide variety of small package sensors which were “signal

64 conditioned on-chip” [2]. The signal conditioning meant that Motorola had already developed the necessary support electronics which would create a strong analog signal in the appropriate range for our selected data loggers. This ultimately meant that the use of Motorola pressure sensors would simplify the schematics down to three primary elements: the power conditioning for the sensors, the sensors themselves, and the data lines returning to the loggers. In addition, Motorola’s diﬀerential and absolute pressure sensors function in a very similar manner and have very similar packaging. The sensing element of these sensors is a small ﬂexible wafer which is instrumented with a strain gage. In the case of the absolute pressure sensor,

Figure 3.1: Internal Layout of Absolute Pressure Sensor

this wafer is mounted so that it bridges across the small sealed internal air pocket within the sensor, see Figure 3.1, which is adapted from [6]. The diﬀerential pressure

65 Figure 3.2: Internal Layout of Differential Pressure Sensor sensor, in turn, is ported on both sides of the wafer so that it may flex based upon the ambient pressure that the sensor is being used in, see Figure 3.2, which is adapted from [7]. When the sensor is exposed to changes in air pressure, the wafer is flexed to one side; it should be noted that Motorola requires the sensors, both absolute and differential, to be installed in such a way that they are always exposed to flexure in one predetermined direction. In the case of the differential pressure sensor, both positive and negative pressures can be measured; however, both cannot be measured in the same sensor without rearranging the porting. The flexure of the wafer causes a change in the resistance of the strain gage. The “on-chip” signal conditioning changes this physical phenomenon to an amplified analog signal, which the data logger can decode and record.

66 (a) Absolute Pressure Sensor [6] (b) Diﬀerential Pressure Sensor [7]

Figure 3.3: Variation in Pressure Sensor Porting

For our specific sensor formation, the front side of each differential pressure sensor was connected via Tygon tubing to a specific nosecone port. The back side of each differential pressure sensor was ducted, along with the inlet, to the absolute pressure sensor. This duct was composed of six stainless steel fish tank T’s and additional

Tygon tubing. The design of this duct left one end floating free while the other end terminated at the first differential pressure sensor. Leaving a single end of this duct open would allow all of the sensors to be exposed to a consistent changing pressure

67 Figure 3.4: Pressure Board inside the nosecone. The absolute pressure sensor would then capture this changing pressure during ﬂight. This would allow calculations to determine the pressures at the various locations on the nosecone’s surface due to the fact that pressure on the back side of each diﬀerential sensor was capture by the absolute pressure sensing instrument.

3.1.2 Acceleration Sensors

Seeing the advantage of using “conditioned on-chip” sensors, it was quickly noted that the accelerometers should have the same built-in function. Motorola’s sensor products guide was searched again, and it was found that Motorola supplies numerous “conditioned on-chip” accelerometers in addition to the pressure sensors.

Motorola’s accelerometers functioned similarly to the selected pressure sensors. The

68 sensing element can be modeled as two stationary plates with a third floating plate between them. This arrangement effectively creates two capacitors whose capacitance is based upon the distance from each stationary plate to the floating middle plate. Motorola refers to this arrangement as a “g-cell”, see Figure 3.5, which is adapted from [4, 5]. When the “g-cell” is exposed to some outside acceleration; the

Figure 3.5: Visual Representation of Motorola’s g-cell

gaps between each of the stationary plates and the floating plate change. This effectively changes the capacitance between each stationary plate and the floating plate.

The ratio of the capacitance is then used by the internal circuitry to determine the respective g-force experienced.

For simplicity, the accelerometers in our rocket system were arranged in a formation to capture the three cardinal vectors of the Cartesian coordinate system. Of

69 the three vectors, the z-axis (the axis which is coaxial to the trunk of the rocket system) was deemed to be the most vital acceleration to record. This importance was due to the thrust of the motor acting on the rocket system along the z-axis. It was expected that the majority of stresses in the rocket trunk would be induced by this motor thrust and the resulting nosecone drag. As for the x-axis and y-axis, it was required that they maintain their orientation to the z-axis to preserve the desired

Cartesian System; however, there would be no active guidance on board the rocket system; thus, the rocket would be free to rotate and pitch while in flight. These recorded values would only reflect phenomena experienced at an undeterminable location in the flight trajectory. Therefore, the z-axis and the magnitude of expected lift off g-forces became the driving factors in the final selection of accelerometers and PCB layout. In addition, all of the accelerometers needed a rigid mounting to the rocket system; securely fastening the accelerometers would minimize errors in acceleration measurements due to damping of forces through soft mounting. The most rigid location in the Instrumentation Section was at the base of the nosecone on the threaded nosecone bulkhead. A circular PCB was designed to lie parallel to the threaded bulkhead. To achieve the desired rigid installation, the PCB was

70 (a) z-Axis Accelerometer (Typi- (b) xy-Axis Accelerometer cal DIP Layout) (Typical SOIC Layout)

Figure 3.6: Motorola Accelerometer Package Layout fastened to the bulkhead with aluminum threaded standoﬀs and socket head cap screws. This method and orientation of mounting allowed the selection of an accelerometer which was designed to sense acceleration in the z-axis. To complete the desired sensing arrangement, a single chip, capable of sensing both x-axis and y-axis accelerations, was selected to complement the z-axis accelerometer. Thus, the desired Cartesian coordinate sensing arrangement was preserved by mounting two speciﬁc accelerometers coplanar to each other.

71 Figure 3.7: Accelerometer Board

3.1.3 Strain Sensors

Thrust forces generated by the motor will naturally create strain within the rocket casing when resisted by the inertia of the rocket system and by aerodynamic drag.

This strain is sensed by an arrangement of thin foil elements called strain gages; these devices are designed to sense the subtle changes in a parts shape when it is subjected to strain. However, unlike pressure sensors and accelerometers, strain gages are not produced in an “integrated on-chip” format. Instead, strain gages are simple foil resistors which change in value as the gage is strained. Since these gages have such a simple function, they require additional components to produce a meaningful signal, which the DAQ can record.

72 One of the simplest methods for operating strain gages is through the implementation of a Wheatstone bridge. In a bridge configuration, the strain gage is balanced by resistors of equal resistance. These bridges require an external voltage excitation source. As the gage is strained, a potential voltage difference is generated across the bridge as the strain gage changes resistance. This voltage differences is only a very slight change; at times, the change may appear to be of the same magnitude as system noise. Therefore, amplifiers are used to convert these slight changes into stronger signals, which the DAQ can interpret [39].

Strain Gage Selection

Strain gage selection takes into account three key factors: the excitation source of the bridge, allowable power density, and initial resistance of the strain gage.

After determining these three features, a strain gage selection chart from Vishay was consulted for the correct strain gage to suit the expected conditions [19, 21, 25].

Excitation Source

Vishay strain gage selection charts cover excitation sources from 0.2 - 100 volts.

This can seem like an overwhelming range from which to choose. However, the choice to use 5 volts as the excitation source was arrived at fairly easily. The ﬁrst

73 factor is the ease at which 5 volts could be generated. The use of a 7805 voltage regulator provides an ample 5 volt source which will drive all of the Wheatstone bridges without worry of overload. In addition, all of the custom circuitry had been designed around the logic level of 0 - 5 volts; thus, utilizing these limits would prevent the need for providing several different voltage levels for different components. The second factor involved the design of the strain gage amplifier. The amplifier design was being based on an idealized schematic from Analog Devices, which incorporated the use of a five volt excitation source.

Power Density

Vishay provides a table for deriving the power density based upon the Heat-Sink

Conditions of the application. The first step is to specify whether the expected strain is static or dynamic in nature. This rocket should expect various changes in forces during flight, and these dynamic forces would naturally create dynamic strain. The second step requires selecting the measurement accuracy level: low, moderate, or high. An accuracy level of high was selected such that minimal information would be lost while recording data during flight. The third step requires defining the heat-sink conditions of the application. This is important because the heat sink conditions will

74 directly aﬀect the amount of error within the strain gage readings. One must realize that strain gages are nothing more than variable resistors. In addition, resistors naturally create heat as a by-product. Since the gages are made from foil, they are prone to thermal expansion and contraction, which will directly aﬀect their resistance. Thus, if the strain gage is excited and produces an excessive amount of heat, which cannot be dissipated into the environment, the strain gage may actually be reporting an incorrect strain value due to a change in its own geometry imposed by internal thermal strain. Thin-wall aluminum tubing is not listed in the Vishay

Table 3.1: Acceptable Power Density, [19]

table; therefore, deﬁning the heat-sink condition requires some subjective reasoning.

The testing material is aluminum and has good thermal conduction characteristics.

The main draw back to the testing surface is the fact that it is thin walled; this

75 geometry will limit its ability to sink heat away from the strain gage. However, the thermal characteristics of the thin-wall tube should be offset by the fact that the testing surface is exposed to open fluid flow. In addition, this fluid flow should be cool considering the decreasing temperature with respect to increasing altitude.

These considerations combined with the initial classiﬁcation of heavy aluminum as

“excellent” in Table 3.1, led to deﬁning the heat-sink conditions of this speciﬁc application as “excellent”. Then, reading from the table, a strain gage should be selected according to a power density of 5 - 10 watt/in2.

Initial Resistance of the Strain Gage

Vishay manufactures strain gages with gage resistances of 120 ohms, 350 ohms, or 1000 ohms. Strain gages with an initial gage resistance of 120 ohms have been considered as “general use” or “all purpose” strain gages; they are commonly used where heat dissipation is not a concern. Strain gages with an initial gage resistance of 1000 ohms are used in high accuracy applications. This is due to the fact that the higher initial resistance will allow the use of higher excitation voltages, which improves the signal-to-noise ratio of the Wheatstone bridge. In addition, higher initial gage resistances reduce the concerns for strain gage heating in situations where

76 the same excitation voltage is being utilized. This can be conﬁrmed by looking at

Equations 3.1 and 3.2 [36].

P = V · I (3.1)

V I = (3.2) R

In Equation 3.1, the power (P ) generated by a resistor is based on the voltage drop

(V ) and the current (I) ﬂowing through the resistor. In addition, Equations 3.2 shows Ohm’s Law and the dependency of the current ﬂow (I) with respect to the restance (R) in the path. Substituting Equation 3.2 back into Equation 3.1 yields

Equation 3.3,

V 2 P = (3.3) R which shows that shows that when voltage is held constant the power generated within a resistive element is reduced as the resistance is increased. Strain gages with an initial gage resistance of 350 ohms provide a excellent mixture of the beneﬁts associated with both 120 ohm and 1000 ohm strain gages. When compared to strain gages with initial gage resistance of 120 ohms, a 350 ohm strain gage’s heat generation rate is reduced by a factor of three. The signal-to-noise ratio is still

77 increased when operating on the same excitation voltage source. These beneﬁts make the 350 ohm strain gage readily available in a wide assortment of shapes and sizes. To further reinforce the selection of a 350 ohm strain gage, Vishay oﬀers bridge completion modules for half Wheatstone bridges. These modules house two resistors, which are balanced to within ± 0.005% of each other.

Strain Gages

To ﬁnd the proper strain gage, excitation voltage and power density are plotted on the 350 ohm strain gage selection chart, see Figure 3.8. The left axis shows excitation voltage and a horizontal grid line of 5 volts is selected. Running diagonally across the graph are trend lines representing power density. Both, the 5 and the 10 watt/in2 trend lines are marked. Minimum and maximum acceptable strain gage areas will result at the points where the power density trend lines intersect the horizontal excitation voltage. The boxes inset in the graph contain part numbers which match grid area sizes. The minimum and maximum acceptable strain gage areas are simply compared to the part numbers to retrieve the optimum strain gages for the speciﬁed conditions.

Strain Gage Bridge Ampliﬁers

78 Figure 3.8: Strain Gage Selection Chart, [19]

The ﬁrst feature built into the ampliﬁer was a precision 20 ohm potentiometer.

This potentiometer allows the strain gage leg of the Wheatstone bridge to be balanced. If a bridge is not balanced, the hardware or software must account for the initial presence of voltage which is produced by an un-balanced bridge. The potential voltage difference generated when a strain gage undergoes some deformation is on the order of microvolts [39]. Therefore, the signal must be amplified to the same range as the DAQ system, in this case 5 volts. However, setting the amplifier’s gain

79 requires some insight as to what level of strain the system will be exposed to. A corresponding voltage can then be calculated by using the maximum expected strain and the initial strain gage resistances. This calculated expected voltage represents the maximum expected strain in the structure. A safety factor can be applied to the expected voltage creating what should be a voltage representation of the maximum expected strain plus a safety margin. The majority of the strains expected in this rocket system consist of longitudinal strains. Therefore, this new voltage was multiplied by two to create a voltage range which would cover the maximum expected strain in both the compression and elongation directions. However, in a case where the system is known to have different maximum elongation and compression levels, those respective limits can be used instead of simply multiplying a single known magnitude by two. To determine the gain level for the amplifier, the voltage range for the DAQ inputs is divided by the voltage range for the expected swing from maximum elongation to maximum compression. After scaling the signal, the amplifier must complete one additional task before the signal is sent to the DAQ.

The potential voltage diﬀerence generated at the Wheatstone bridge is centered on the ground reference of the excitation source, 0 volts. Thus, when a bridge has

80 been properly balanced, a compressive strain will produce a negative voltage and an elongation will produce a positive voltage. If the signal is not biased to 2.5 volts then DAQ’s ground reference of 0 volts will clip all negative values it receives.

3.1.4 Data Loggers

Three-axis acceleration measurements, seven pressure sensors, three thermocouples, and nine Wheatstone bridges meant that the data logger would have to have a total of twenty-two Analog to Digital Converter (ADC) input channels. A quick search of available data loggers revealed that manufacturers simply don’t produce small form factor DAQ equipment with the number of ADC ports which were required. Brain- storming provided two viable solutions; either multiplex the data lines to allow more than one line to talk to each port or use more than one data logger. Multiplexing would have required additional electronics to serve as the gate keepers, directing data

ﬂow to speciﬁc ADC lines. With the addition of more electronics to achieve multiplexing, it was feared that the circuitry and the required additional programming would generate undesirable latency in the data logging. In addition, multiplexing looked to be just outside of both the current scope of electronics knowledge and

81 an aﬀordable learning curve. Therefore, we opted to ﬁnd a small stand-alone DAQ board, which was easily programmable and stackable. The Onset Computer Corpo-

Figure 3.9: Onset Tattletale Data Logger

ration produces the Tattletale data logger; this small device has eleven twelve-bit and eight eight-bit ADC channels. To achieve the needed twenty-two ADC lines, two

Tattletales were programmed to work together in data collection to cover all incom- ing data lines, excluding one thermocouple channel, with the twelve-bit ADC lines; the excluded thermocouple would be accounted for on a free eight-bit ADC line.

Using two DAQ boards did present one hurdle, which had to be addressed. Both boards would have to be synchronized in order for data analysis to be meaningful.

Conclusions could not be drawn if the loggers couldn’t show the same phenomenon

82 occurring at the same point in time. In order to overcome this synchronizing challenge, the two DAQ boards were arranged in a master and slave conﬁguration. The

Figure 3.10: Master/Slave DAQ (Seen During Final Testing)

master would serve two basic roles. First, it would be responsible for logging eleven data lines at a ﬁxed sampling frequency. Secondly, the Master would send pulses to the slave so that it knew when to take a paired reading. It is important that both the master and the slave complete measurements faster then the required sampling rate; in this scenario, the master will always be ready to send the next pulse, and the slave will always be awaiting this pulse. This method allows the time lag between measurements taken on the master and slave DAQ boards to be minimized to an error on the order of microseconds. To improve the robustness of the master and

83 slave conﬁguration, similar data lines would be split between the master and slave data loggers. This would allow both data loggers to be responsible for approximately half of the data lines streaming from each sensor type formation. The only exception from this division of work was that both the master and the slave recorded the acceleration in the z-direction. This redundancy was to serve as a marker in matching the paired data sets from the two data loggers. One additional marker was added to the data sets for alignment. Both, the master and the slave data loggers added a blank byte in the ﬂash memory after ever sequential 1000 measurements. With all three safety markers in place, the recorded data should easily be matched in paired groupings for analysis. A single circuit board was constructed to hold both the master and the slave data loggers. This circuit board also served as the primary collection point for all data lines returning from the onboard rocket system sensors.

The returning data signals were carried through ribbon cables to the master/slave circuit board. The distribution of these signals was then hardwired to the respective data logger via traces “burnt” on the circuit board. In addition to collecting the data signals, the master/slave circuit board also served as the only source of power distribution. Each of the ribbon cables, which carried data signals, also carried power to

84 the respective sensors. The master/slave circuit board contained eight LEDs (Light

Emitting Diode) with driving circuitry, four LEDs for each Tattletale data logger, which could be utilized as status indicators during testing and before launch. Each

LED driver was wired directly to a speciﬁc output pin on the Tattletale data loggers. The drivers consisted of simple transistor and resistor circuits commonly used for this function. During programming, these pins could be set high to excite their respective transistors. The transistor would then allow current to ﬂow through the

LED, causing it to glow. Ultimately, this feature would be used to indicate that the logging software had entered a speciﬁc routine in the program ﬂow.

3.2 Parachute Deployment

The two most common methods for determining altitude during ﬂight require either inertial measurements or atmospheric pressure measurements. Of these two methods, deriving altitude based upon atmospheric pressure is considered to be the most fundamentally sound methodology. Such altimeters determine altitude simply by sensing the pressure gradient associated with the Earth’s atmosphere and applying that reading to the US Standard Atmospheric Model. This type of altitude sensing

85 can be conducted with a single absolute pressure sensor. In contrast, inertial altimeters for rockets of this scale require three roll rate sensors plus an accelerometer.

This complexity stems from the fact that high power rockets are not actively guided and can change orientation while in ﬂight. Therefore, to capture these ﬂuctuations in trajectory, inertial altimeters are composed of a z-axis accelerometer, which is supplemented by yaw, pitch, and roll rate sensors. In addition, inertial altimeters must be mounted at the center of gravity for the rocket system. This ensures that the rocket’s trajectory can be calculated in real world coordinates. If inertial altimeters are not designed with this level of complexity, they run the risk of improperly determining altitude and prematurely deploying a parachute. For example, an inertial altimeter with only a z-axis accelerometer would correctly deploy the parachute if the rocket followed the idealized trajectory. However, the same altimeter could incorrectly deploy the parachute if the altimeter was ever turned over on to its side as when rockets “wind seek” in high winds. This “weather vaning” motion would give the accelerometer a false reading along the z-axis. Overall, this false reading could trick the altimeter into believing that the rocket was at apogee with a minimal velocity, as in the case of the idealized trajectory, instead of the true projectile

86 path at relatively high speed. Thus, the altimeter would deploy the parachute and possibly cause catastrophic damage. Regardless of the type utilized, it can be seen that altitude, velocity, and the ﬂight trajectory are critical in safe parachute deployment. However, of these three components, altitude is the most critical decision maker. Once altitude is correctly calculated, parachute deployment controllers can be programmed to take desired action.

Deployment of the parachutes was considered to be one of the largest safety concerns of the entire rocket system. This concern was generated by the relationship between the center of gravity and the center of pressure. The locations of these centers, with respect to each other, indicated that our rocket system would be very stable during ﬂight. If the rocket didn’t separate at apogee, the system could return to the launch site as a stable projectile. Impact of our rocket system at terminal velocity would destroy the main structure including all instrumentation inside and could possibly cause damage to people, livestock, or infrastructure. Therefore, the design called for a triple redundant system. This system would utilize both commercial and homebuilt altimeters. A commercial altimeter would serve as the primary deployment device, while two homebuilt altimeters served as redundant backups.

87 Each altimeter would be wired to its own set of charges for rocket separation and parachute/drogue deployment. This would create three parallel systems, which were completely independent of each other. The introduction of multiple parallel deployment devices greatly increased the reliability of the separation and parachute deployment sub-system. Multiple altimeters also carried another distinct advantage.

They ensured a correct altitude reading for apogee. If only a single altimeter was used and it reported a false altitude it would be diﬃcult to prove. However, the use of multiple altimeters allowed for comparison maximum altitude readings once the rocket system was recovered.

3.2.1 Homebuilt Altimeters

The homebuilt altimeters of our rocket system were designed and constructed by one of the high power rocket enthusiasts who were helping with the project development.

These pressure altimeters were designed with onboard microprocessor, supplemental

ADC IC, and an absolute pressure sensor; both, the microprocessor and the pressure sensor were manufactured by Motorola while the supplemental ADC was manufactured by Linear Technology. Together, these components form a fairly simple circuit.

88 Figure 3.11: Homebuilt Altimeter

The signal from the absolute pressure sensor was routed to the supplemental 12bit

ADC. The 68HC11 microprocessor uses a single line of port C to read the digitized pressure from the supplemental ADC. The 68HC11’s embedded program dictates the actions required based upon this pressure reading. The use of the homebuilt altimeter and its 68HC11 offered a wide range of flexibility. The program could literally be re-written to perform actions at different altitudes. Ultimately, these altimeters gave great flexibility which only required minutes to change.

A problem that altimeters have to tolerate is the presence of pressure irregularities due to transition from subsonic to supersonic speeds and vice versa. During these transitions, the altimeters can experience a sudden increase or decrease in pressure as the shockwave retreats or advances from the nosecone to the aft of the rocket.

Since the fundamental design of an altimeter is to sense the point of increasing pressure (pressure increases as the rocket falls from apogee to launch site), altimeters

89 without any safety features can cause premature rocket separation and parachute deployment. To overcome this problem, our homebuilt altimeters employed a two point backward difference method to calculate the current ascent rate [14, 13]. Then, a three point finite difference was calculated by comparing the current ascent rate to the previous ascent rate to show any discontinuities in the ascension acceleration.

These calculated pressures and accelerations were then used to sense ﬁve speciﬁc events.

x Determine when an “arm” altitude had been reached

y Determine if the rocket had gone supersonic

z Determine if the rocket had returned to subsonic

{ Determine if apogee had been reached

| Determine if 400 ft AGL had been reached

Determine when an “arm” altitude had been reached - As the rocket lifts away from the launch pad, the altimeter is “locked out” until such a time that the rocket passes the safety deck of 400 feet. This safety precaution is in place to prevent anything from triggering the altimeter and the parachute deployment charges while the rocket is on the launch pad. Once the rocket passes the safety deck, the altimeter

90 gains full functionality and shifts focus to apogee detection (this includes supersonic detection).

Determine if the rocket had gone supersonic - The ascent acceleration is constantly being monitored at a rate of two hertz by the altimeter. When a large negative acceleration rate is discovered, it is compared to a preset value of -800 f/s2.

If the sensed acceleration is greater in magnitude, then the altimeter has experienced the eﬀects of a pressure shockwave as it moves toward the nosecone of the rocket.

Under this circumstance, a supersonic inhibit bit is set in the microcontroller; this will prevent the microprocessor from deploying the parachutes at supersonic speeds.

However, if the acceleration is not greater than -800 f/s2, it is assumed that the rocket is merely slowing down due to aerodynamic drag; this condition would take no action on the supersonic inhibit bit.

Determine if the rocket had returned to subsonic - Given the mass of the rocket and the expected thrust from the motor, it would be impossible for this rocket system to achieve true acceleration rates of 50 g’s. Therefore, if ascension acceleration is calculated to be above 1,600 f/s2, this must be indicating the retreat of the pressure

91 shockwave from the nosecone to the tail of the rocket. This condition would clear the supersonic inhibit bit in the microprocessor and allow the parachutes to be deployed.

Determine if apogee had been reached - Apogee is declared when two condition are met: the ascent rate must be less then 30 fps and the supersonic inhibit bit must be cleared. When these conditions are met, the altimeter makes a ﬁnal reading and stores that value in memory. The altimeter then deploys the drogue parachute.

Determine when 400 feet was reached during decent - After apogee, the altimeter still reads pressure at the rate of two hertz. When an altitude of 400 feet is reached, the main parachute is deployed. Also, the apogee altitude is displayed on a seven segment LED. The output reads as a series of numbers on a single seven segment

LED. The output begins with a long pause and is then followed by each digit of the recorded altitude. This output string, along with the pauses, is repeated until the altimeter has been powered down. An example output of “long pause - 1 - 5 - 4 - 8 -

5 - (repeated)” would indicate a maximum altitude of 15,485 feet above the launch deck (AGL).

The original design for the homebuilt altimeters didn’t include any remote status indicators. Thus, verifying that the altimeters were functioning properly would be

92 diﬃcult due to the fact that they were securely fastened deep within the altimeter section of the rocket system. This prompted a change for the original homebuilt altimeter circuit board. It was modiﬁed to allow the remote mounting of a blinking

LED. This LED could be mounted close to the outer skin of the rocket where it would be visible through a small hole. The reasoning for the use of a blinking

LED, as compared to a non-blinking LED, comes from the environment in which the rocket would be launched in. High power launches require minimal cloud cover and the absence of strong winds, in other words - an average day or better. Trying to determine whether or not a solid LED is on or oﬀ can be diﬃcult in daylight.

However, when the LED is blinking it becomes a slightly more obvious indicator. In order to activate this remote LED, it would be connected to the altimeter with two small wires.

3.2.2 Commercial Altimeters

Adept Rocketry produces numerous altimeters and carries a strong rapport with high power rocket enthusiast. For our rocket system, we chose to use a recording

93 altimeter, ALTS2-50K, with the capability to deploy two parachutes. The ALTS2-

50K altimeter is an upgraded version of the ALTS2 altimeter from Adept Rocketry; its main advantage is the ability to recorded maximum altitudes (apogee) on the order of 50,000 feet above see level. Adept Rocketry states that the resolution of

Figure 3.12: Adept Rocketry Altimeter

their altitude reading is on the order of one foot [1]. This resolution is accomplished with an absolute pressure sensor and a sixteen-bit ADC. During flight, Adept’s altimeters are constantly differentiating altitude in real time. Upon reaching apogee, the altimeter is designed to automatically deploy the first of two parachutes. This deployment on apogee is a feature which cannot be changed. In a similar manner, the deployment of the second parachute is fixed to a specific altitude as well. However,

Adept Rocketry includes a jumper setting on the altimeter which would allow the user to select whether the second parachute was deployed at an altitude of 750 feet or 250 feet.

94 Even though the altimeters produced by Adept Rocketry have limited flexibility in use, they do carry a very important safety feature, which deals with the pitfalls associated with “lockout timers”. The primary problem with “lockout timers” begins with the fact that they are simply designed to sleep and wait for a set amount of time to pass after lift off in an attempt to ignore the expected irregular pressures associated with supersonic flight. Adept Rocketry illustrates the failing of this method with the following example:

For instance, if a “lockout” timer is set to ten seconds and the motor has a failure, the rocket may very well crash into the ground before the “lockout timer” allows the parachute to deploy and save everything.

Instead of employing “lockout timers”, Adept Rocketry programs their altimeters to constantly monitor pressure readings and simply dump the irregular readings.

This allows the altimeter to retain full functionality as long as two conditions are met: the altimeter must sense that it is 300 feet above the launch deck and has slowed to an ascent rate of less then 500 feet per second. Adept Rocketry states:

Problems with premature deployment due to sound barrier eﬀects cannot occur in Adept altimeters as they can in others. And there is none of the problems associated with “lockout timers.”

95 In addition to being capable of deploying two parachutes, the ALTS2-50k also records the maximum altitude reached. This altitude is then relayed to the user via audible tones instead of LEDs. The beneﬁt here is the fact that the tones can be heard from outside of the rocket casing. This means that, upon recovery of the rocket, maximum altitude is known without even picking up the rocket or using any tools. This audible indicator begins with a long pause. Then, each digital of the recorded altitude is beeped out (one - beep, two - beep, beep, three - beep, beep, beep ). Once a digit has been beeped out it is followed by a short pause before continuing to the next digit. In a manner similar to the homebuilt altimeters, the cycle is repeated until the altimeter has been powered down.

3.3 Recovery

The tried and true method of locating “misplaced” high power rockets after launch is with the use of a radio transponder [3]. This method is very similar to the way that

“tagged” wild animals are tracked in the great outdoors. To aid in recovery of our rocket system, a radio transmitter, radio receiver, and antenna were purchased as a set from Walston Retrieval Systems. At the base of the instrumentation section, the

96 Figure 3.13: Walston Retrieval System (Transmitter Not Shown) rocket was outﬁtted with a CA MODA 3750 MVS-HI radio transmitter. This speciﬁc radio transmitter had been designed to operate on a preset frequency of 217.394

MHz. The CA MODA 3750 MVS-HI is considered to be an upgrade from standard transmitters and is rated for situations where long range sensing was required. Our speciﬁc radio transmitter was rated for 90 miles line-of-sight or 9 miles on the ground

[38].

To accompany the radio transmitter, we used a TRX-3S three channel radio receiver which was speciﬁcally designed for use with our CA MODA 3750 MVS-HI radio transmitter. One of the three available channels was designed to work with frequencies ranging between 217.390 - 217.400 MHz. This band would allow us

97 Figure 3.14: Front Panel of Receiver to hear our purchased transmitter. The other two channels were set to arbitrary frequency ranges of 217.310 - 217.320 and 217.350 - 217.360 MHz. To use these arbitrary ranges, transmitters with frequencies within the ranges would have to be purchased.

A quick connect directional antenna was included for use with the TRX-3S receiver. It required only minor assembly before it could be used. Three elements slid into a main shaft and were held in place with thumbscrews. This hand held antenna could then be connected to the receiver with a coax cable which hangs from the middle element of the antenna.

98 Use of the receiver and transmitter takes some practice before a ﬁrm understanding of the system is gained The shape of this antenna gives it a very narrow beam width for sensing the transmitter, on the order of 30◦. This is further limited by the fact that the antenna is polarized; thus, not only is it sensitive to being pointed in the right direction but it is also sensitive to whether or not it is orientated in the same direction as the transmitting antenna. However, the limits on orientation can be viewed as positives when hunting for a rocket. If the signal is stronger when the antenna is held vertically then one might expect the rocket to be hanging from a tree, stuck nose ﬁrst in the ground, or possibly still be making a descent to the ground.

3.4 Printed Circuit Board Design

Several sequential steps were taken to develop each of the custom circuit boards.

The ﬁrst step was to consult each of the data sheets for the chosen sensors to ﬁnd schematics depicting the recommend supporting circuitry. Then, a computer software package called Protel SE was used to create digital libraries containing pin

99 layouts and printed circuit board footprints for each of the sensors and their respective supporting circuitry. These digital libraries were then used to create a series

Figure 3.15: Electronics Design Process

of schematics which integrated all of the sensors, respective supporting electronics, and the DAQ. After at least two people analyzed each of the resulting schematic for errors, the schematics were reconstructed on circuit prototyping bread boards.

These prototyped circuits were tested for proper functionality. The next phase was to create the actual custom printed circuit boards, to which all components would be soldered. Again, Protel was used to graphically arrange each component within a boundary, which represented the edges of the desired board. The traces connecting each of the components on the printed circuit board were formed using a built-in op- timization feature of Protel called “Autoroute”. These digital printed circuit boards were analyzed for errors just as done with the digital schematics. Protel was used to generate a set of Gerber ﬁles. This process usually generates more ﬁles than is

100 necessary for manufacturing simple two sided printed circuit boards. In addition to the file containing the hole drill pattern, only the 4 files containing the traces and solder mask for the top and bottom layers were required for processing. These files were then zipped together into a compressed computer file (.zip) and emailed to a manufacturing company name Express PCB. Once the board had been fabricated, it’s returned via mail. The components were then placed on the board and soldered into place. The final step of the process was testing the completed boards for proper function.

3.5 Software

The Onset Computer Co. provides a programming language called TFBASIC for the purpose of programming their Tattletale DAQ equipment. This proprietary syntax is nothing more than a set of helper functions which are designed to make programming the onboard 68HC11 easier [34]. However, one is not required to use the supplied TFBASIC Software; in fact, the Tattletale can be programmed in nothing but Assembly code if one so chooses. Our code used a mixture of both. See

Appendix B for the actual code, which was used on the day of the launch.

101 Onset did not provide ﬂexibility for downloading the code to the Tattletale DAQ.

A proprietary communication program called TFTools was used to send compiled programs to the Tattletale via the host computer’s parallel port.

102 Chapter 4

System Testing

Only two “ready-for-ﬂight” electronic components were purchased for use on this rocket system: the Walston transponder and the commercial altimeter from Adept

Rocketry. This meant that the entire custom DAQ system and the recovery system would have to undergo numerous tests to ensure that the complete electrical system was functioning properly before the maiden ﬂight. This validation process would begin by testing each custom electronic board separately; after verifying proper functionality separately, the entire electronics package would be assembled on a bench top and tested as a complete system. The ﬁnal stage of testing would require assembling the electronics with the rocket system and performing a mock launch and recovery.

103 4.1 Pressure

One of the primary reasons for selecting the Motorola MPX4250 and MPX5500 pressure sensors was the fact that they produced an output signal which was linear to the pressure reading of the sensor. This linear relationship was given to be true throughout the entire operating range of the sensor. However, this linearity was dependent upon the utilization of the prescribed decoupling circuit illustrated in the products’ data sheet. In addition to the recommended circuit, the sensor data sheets also documented the transfer functions which illustrate the expected output

voltage (Vout) when using the sensor in the approved manner. For each sensor, these

functions were given in respect to two variables: the supply voltage (VS) for the sensor and the pressure reading (P ) in kilopascals. The absolute pressure transfer function is shown in Figure 4.1 [6] and the diﬀerential pressure transfer function is shown in Figure 4.2 [7].

Vout = VS · (0.004 · P − 0.04) ± Error (4.1)

Vout = VS · (0.0018 · P + 0.04) ± Error (4.2)

104 A short series of experiments was performed to verify that the custom fabricated circuit board was behaving according to the expected transfer functions. Each sensor, in turn, was ﬁtted with a small testing apparatus. This apparatus consisted

Figure 4.1: Pressure Testing Apparatus

of a valve stem, hose nipple, and an analog pressure gauge, which were all housed within an aluminum block. A piece of Tygon tubing was attached to the hose nipple.

The other end of the tubing was attached to a stainless steal T-junction, which was placed between the nosecone ports and their respective sensor. This apparatus was charged with air via the valve stem. While pressure was stored within the setup, the dial gauge was watched to verify that the system was stable and not leaking. An

105 oscilloscope was used to measure the source voltage and the output voltage of the sensor. These voltage values and the reading from the pressure gauge were recorded.

The observed pressure reading and the source voltage were then used in the supplied transfer functions to calculate an expected output voltage. A comparison of the expected output voltage and the recorded output voltage showed that the circuit was operating as predicted by the published transfer functions.

4.2 Acceleration

Similar linear relationships exist with the Motorola MMA3201D and MMA1201P accelerometers, although, upon examination of the respective data sheets, one can see that Motorola does not explicitly state the typical transfer function when used in the recommended decoupled circuit. Instead, Motorola states the zero g voltage reading and the sensitivity of the sensor in mV/g. However, using these points a transfer function, see Figure 4.3 where g-force is (g), was derived. Both sensors utilize the same transfer function due to the fact that they have the same measurement range [4, 5].

Vout = VS · (0.0125 · g + 0.5) ± Error (4.3)

106 Validation of the accelerometers began with some simple experiments. The custom acceleration circuit board was mounted to the bulkhead of the nosecone as it would be mounted during flight. The bulkhead was then laid flat on a table; such an orientation would impose 1 g on the z-axis and 0 g’s on the x and y-axes. The oscilloscope was used to measure the supply voltage and each of the three respective output voltages. In an attempt to collect a variety of samples, this process was conducted an additional five times in different orientations. This was done by setting the bulkhead on its side and rotating it such that the x and y-axes were exposed to the following combinations of g force: 1 g in the +X, 1 g in the +Y, 1 g in the

-X, 1 g in the -Y, and 1 g 45◦ between +X and +Y. After recording the respective voltage readings, an analysis similar to the one conducted on the pressure gages was preformed. The measured supply voltage and the experienced g force were used to calculate an expected output voltage for each of the three axes in all six scenarios. These expected values were then compared to the actual recorded values. The comparisons showed that the circuit was behaving as predicted by the data sheets.

However, it would be impossible to verify that circuit was functioning exactly as

107 portrayed by the transfer function. The sensitivity of the accelerometers is approximately 50 mV/g, thus a swing from -1 g to +1 g is on the order of 100 mV. When this is compared to an operational output range of 0.5 - 4.5 volts, the measured range of 100 mV is not adequate enough to validate the linearity or overall accuracy of the accelerometers. A capstone group continued testing the accelerometers by placing them inside a drop module, which was released from a height of approximately 50 ft. They took the recorded data and compared it to expected values. They reported back that there was no noticeable error in the accelerometers. With this being the case, it was assumed that the sensors were functioning according to the transfer function since the output observed was behaving as expected.

4.3 Strain Gages

The ﬁrst step to validating the strain measurement system was to verify the proper installation of each individual strain gage. One must understand that the mounting surface of a test section undergoes a signiﬁcant amount of preparation before a gage can be bonded to the part under investigation. This needed preparation is to ensure that the strain gage accurately senses the strain in the surface which it is bonded

108 to. During the bonding process, the strain gage is clamped to the test surface; this method insures that a good bond is created while preserving the desired orientation.

However, this step of the bonding process does produce a side effect. The strain gage, once bonded, naturally stabilizes with some residual strain. This residual strain can be of any magnitude of micro-strain from infinitesimal on up. In addition, the residual strain can result from uncontrollable factors like the geometry of the test surface. Whatever the cause may be, the level of residual strain in a gage becomes a critical design factor when taking two strain gages and placing them together in a Wheatstone bridge configuration. The magnitude of the difference in residual strain between a pair of strain gages will determine the required calibration shunt necessary to allow the bridge to function properly. Special equipment is required to accurately measure the residual strain in a bonded gage. A Vishay Model 1300

Gage Installation Tester was used to make the necessary residual strain readings for each individual gage. In addition, the Gage Installation Tester can also determine whether or not the strain gage is creating an undesirable short across the gage when mounted to an electrically conductive material.

109 The next step required the use of DAQ equipment and software from National

Instruments. Each gage, in turn, was connected to a quarter-Wheatstone bridge ampliﬁer, and the software was calibrated for the respective residual strain. Then, while the instrumentation casing was resting on two wooden blocks, a weight was place in the center of the section. This added weight induced a strain on the casing.

The Labview VI (Virtual Interface) was used to display the resulting strain on a computer screen. The results from most of the strain gages produced values with no deviation from the expected results, although, a couple of strain gages had slight deviations in their micro-strain readings. These deviations could be attributed to numerous factors: motion of the suspended weight, a slight diﬀerence in the distance between the wooden blocks, or a possible misalignment in the position of the strain gage, either in respect to the test surface, in respect to the testing apparatus, or both. To validate these results, each strain reading was compared to calculations, which were based on simple beam bending theory.

110 After verifying that the individual strain gages were functioning properly, the

National Instruments DAQ equipment was replaced with the custom strain gage amplifier, and the experiments were repeated. Longitudinal strain gages were paired together to form half-Wheatstone bridge configurations while the rosettes were paired to create full-Wheatstone bridges. All lead wires connecting the strain gages to the circuit board were soldered instead of using some type of disconnects; this was done to minimize any change in resistance trough the circuit due to irregularities in manufacturing or wear and tear [29, 27, 26]. Each channel of the amplifier was tuned to the specific pair of strain gages which were connected. The instrumentation casing was then placed in the same experiment setup as used in the first round of experimentation. The amplifier was turned on, and three complete channels (six total strain gages) could be tested at once due to the fact that all the strain gages were excited by the custom amplifier. After the amplifier was calibrated, the output was measured with an oscilloscope [24]. The instrumentation casing was then rotated

90◦ such that the remaining three channels of longitudinal strain could be tested.

The values showed the sensitivity of the ampliﬁer to a known strain.

111 4.4 DAQ System

DAQ testing began with writing simple software routines to record data values from the ADC ports of the processor and to store the respective values in the ﬂash memory of the DAQ Board. Then, after recovery of the rocket system, the data could be oﬄoaded to a PC via a function within the Tattletale development kit software from

Onset Computers. The next step of development required recording the amount of time necessary for the routines to loop through all of the ADC ports. The software was written to loop through the “record and store” routines in the most optimized and controlled manner possible. This method of looping was controlled by an output compare timer built into the DAQ processor. The robustness of this “read and store” method was tested numerous times by setting the DAQ to record data until the flash memory was full. These data sets were offloaded to a PC and examined for any strange readings. One of the purchased DAQ boards continually recorded strange data into a specific bank of the flash memory. This malfunctioning board was replaced with a new board and retested. The third step of development required syncing two Tattletales together such that all the data lines could be recorded in

112 known regular time intervals. This was accomplished by establishing a Master and

Slave (Mother/Daughter) setup where one DAQ’s timing was dependant upon the others. To implement the required timing, an output pin on the master was pulsed low when a reading was required. The slave then waited for this pulse and initiated the reading when signaled. This method required only a small modification of the software. After the altered software was downloaded to the master/slave DAQ setup, the configuration was tested by simulating a launch condition that would start the record and store routines until the flash memory was full.

4.5 Parachute Deployment

The two electronic parachute deployment devices were purchased from separate ven- dors who guaranteed their functionality; thus, no major testing was done before the components were attached to the altimeter cradle and installed in the rocket. How- ever, there were many concerns about whether or not the rocket would separate once the deﬂagration charges had been deployed. One of the primary concerns was determining the correct size of the charges needed to separate the rocket. Calculating the size of these charges would take into account the volume of the parachute chambers

113 and the severity of the desired separation, see Section 2.1.2.6. Another concern was the sliding tolerance between the couplers and the casings of the adjoining sections.

Initially, it was thought that too much air would escape the parachute chambers between the mating surfaces before complete separation had been achieved. This prompted an experiment consisting of four trials. The first two trails tested the ease of separation while only using one single charge in both the drogue and main parachute chambers. The rocket was prepped indoors and taken outside where it could be laid on the ground. Deploying the deflagration charge in the main parachute chamber was the first trial. A 2 gram charge successfully separated the rocket’s booster and altimeter sections as planned. However, when the single one gram charge was deployed in the drogue parachute chamber, the rocket’s altimeter and instrumentation sections barely separated. The next stage was to simulate the ignition of three charges at the same time within a single parachute chamber; this would simulate a condition where the triple redundant altimeters all fired simultaneously.

After cleaning the rocket of the residue left by the black powder, the rocket was reloaded with three separate two gram charges and three separate one gram charges.

However, without realizing the mistake, the charges had been placed in the wrong

114 chambers meaning that there was now 6 grams of black powder in the smaller drogue parachute chamber and only 3 grams in the main parachute chamber. This mistake led to an interesting discovery. The three grams in the main parachute chamber separated the rocket sections with ease; but, the use of three 2 gram charges in the smaller drogue parachute chamber clearly separated the sections with considerable force. Therefore, after the trials, it was decided that all separation charges would be two grams; this would safely deploy both the main and drogue parachutes in either of the two scenarios: only one charge ﬁring or all three at once.

4.6 Transponder Experiments

Experimentation with the transponder system began as soon as the electronics were received from Walston Retrieval System. Specially shaped batteries, which came with the equipment, were installed in the transmitter and receiver. The transmitter was then turned on and placed just outside the doors exiting the lab. While standing approximately 25 yards away, the antenna was swept back and forth while listening through the headphones for the locating chirp from the transmitter. This ﬁrst round of testing was an attempt to gain some idea of what a transmitter chirp sounded

115 like; it was quickly learned that a great deal of tuning was required based upon your distance to the transmitter. This prompted the notion to increase distance by a significant amount and slowly decrease the distance until a signal could be detected. To accomplish this, the transmitter was once again placed outside and the receiver was used from a distance of approximately 2.5 miles. This proved to be unsuccessful since the test was conducted within city limits. Even though the transmitter was rated for distances up to nine miles on the ground, the transmitter was not strong enough to pass a signal through buildings. Thus, it was noted that a significant amount of testing would be required to properly learn how to use this system effectively, and a member of the team was appointed to master the retrieval system.

116 Chapter 5

Results and Lessons Learned

5.1 Results

This project culminated with the launch of the rocket on November 29, 2003. The rocket appeared to have a flawless powered flight phase. However, trouble began once the rocket entered the coasting phase of the flight. After much brainstorming, we believe that the rocket encountered wind shear during flight that was much higher than the speeds which were measured on the ground. It is believed that this excessive force would have caused the rocket to weather vane into the wind even though the rocket was still traveling at a relatively high speed. Once the trajectory had taken a horizontal path, the altimeters would have been incorrectly triggered to fire the drogue parachute. Deploying either of the parachutes at high speed could cause

117 Figure 5.1: Weather Vaning Flight Profile countless damage to the rocket. However, without finding the rocket debris, it is difficult to assess whether or not our failure evaluation is correct.

118 5.2 “Pressure-to-Launch”

During the design and construction of this rocket system, the team took part in weekly status meetings. This gave the team a convenient means of communicating goals, requirements, and concerns. One such topic that arose on numerous occasions concerned deﬁning the conditions that would constitute a “go-for-launch” decision the morning before the launch. Consistently the team concurred that it would not succumb to the anticipated pressure to launch the rocket. However, with all of our best intentions telling us not to, we found ourselves making a risky decision just before the launch for the following reasons:

x The research grant had “ran out”

y The project was approaching 2 yrs, original goal was 3 months

z We drove all the way to Sayre

{ The sponsors were there for the launch

| The wind “isn’t too bad”

The research grant had “ran out” - As the project drew on, money quickly became an issue. It could be seen that the monetary amount allotted for parts and equipment was insuﬃcient. The purchased modular rocket system did save construction

119 time but was traded for a premium price. Likewise, the strain gage ampliﬁer took two costly design iterations before a solution was found. These premiums and development costs ultimately required us to supplement the budgetary needs through other sources of funding.

The project was approaching 2 yrs, original goal was 3 months - Originally, the team thought that the project would take, at most, 3 months to design and construct the rocket system. However, parts like the strain gage amp took months to complete, leaving the team constantly guessing at how long development would take.

We drove all the way to Sayre - After driving all the way to Sayre, Oklahoma, the team felt a sense of urgency to launch the rocket and ﬁnish the project. In addition, the time spent prepping the launch site and rocket system only added to the “pressure-to-launch”. By the time the rocket was on the launch pad, it was diﬃcult to even consider that it should be taken back down

The sponsors were there for the launch - Having the sponsors in attendance presents a duel dilemma. On one hand, the team was ecstatic to show oﬀ all of the hard work. However, their presence also proved to be intimidating. The last thing

120 the team wanted was to appear incompetent by not being able to get the rocket oﬀ of the launch pad.

The wind “isn’t too bad” - It has always been known that wind, or simply weather, was a critical aspect of a successful high power rocket launch. However, it was unclear what effect these factors, at any level of severity, would have on the rocket during flight and recovery. For instance, the wind speed just before launch was steady at approximately 18-20 mph and had intermittent gusts on the order of 20-22 mph. At first inspection, the wind speed seemed high but it was still well within the NAR limitations of 25 mph at the launch site. When combining the observed wind conditions with the fact that the sky was clear blue, it didn’t seem like a bad day for a launch. Of course, one noticeable down side was that the rocket would drift a substantial amount due to the high wind and the slow decent rate during recovery, but this was considered to be an acceptable side effect of the observed weather conditions.

121 5.3 Lessons Learned

In the minutes, hours, days, and weeks that followed the launch of the rocket, endless discussions took place concerning what went wrong and what should be done next.

These discussions left us with the following hard “lessons learned”.

5.3.1 Wind Is Critical

In reality, the severity of the wind and the high stability rating of the rocket system were not a good combination. It could be seen that parachute deployment happened much too soon in the ﬂight time line. A coast stage of 20-30 seconds had been expected. Instead, the baby powder clouds were seen around ﬁve-ten seconds after motor burn out. In addition, the clouds appeared a disturbing distance up-wind.

This led us to believe that the rocket had drastically weather vaned into the wind and possibly deployed the parachutes while still coasting a high velocity. In retrospect, launching a balloon before actually launching the rocket would have served as a more eﬀective visual aid in judging the severity of the wind.

122 5.3.2 Rocket Transponder Use Is An Art Form

At several points throughout the project, the transponder and the tracking equipment were taken outside for practice sessions. It was found early on that the performance of the system was directly coupled to the severity of the obstructions in the line of sight between transponder and receiver. Therefore, it was noted that the system truly needed to be tested in an environment which was similar to the area surrounding the launch site, in other words - a rural area as compared to a building ﬁlled urban environment. In addition, the practice sessions were quickly becoming time consuming due to the excessive amount of slow-paced walking over distances on the order of a few miles. After realizing the complexity of this system and the subsequent time requirements, the task of mastering this recovery system was assigned to a team member who, in turn, would lead the recovery eﬀort at the launch site.

On the day of the launch, several false positives were identiﬁed during the recovery eﬀort. The receiver continually picked up interference coming from electrical transformers in the surrounding area. These strong false positives, combined with

123 the limited operational knowledge of the recovery system, presented a big challenge.

In addition, it was unknown whether or not the transponder was working correctly after the parachutes deployed prematurely up-wind.

In the days following the launch, the World Wide Web was consulted for a tutorial on the use of the Walston Retrieval System or one similar in nature. This search turned up only a couple of sites with subject matter close to the desired topic.

However, of the limited results, a high power rocketry web page by Vern Kwoles had a link to a PDF document by Sue McMurray which discussed “Walston Tips and

Tricks” [33]. Her tutorial proved to be invaluable and highlighted several ﬂaws in our recovery process.

The mounting location of the transponder and its antenna was continually debated during the design and construction of the rocket. At the end of this debate, it was decided that the transponder would be placed at the bottom of the instrumentation section. The transponder would be rigidly mounted to a bulkhead and the curled up antenna would rest on top of the parachute until it was deployed. Our mounting method basically placed the transponder inside a Faraday Cage. This

124 phenomenon eﬀectively blocked any signal from leaving the metal tube, which ren- dered it undetectable by the receiver at any range. Sue McMurray argues that this is an unnecessary risk. She states that there are 4 basic rules which apply to the installation of a SuperXmtr:

x Protect the transponder

y The antenna should always be straight

z The transponder should never be installed in a Carbon-Fiber or Metal tube

{ The transponder and antenna should always be inside the airframe

When following this credo, she stated that she had recovered rockets up to 26.8 miles away. In addition, she spoke of one particular instance where a rocket had been recovered even though the antenna had broken oﬀ. Regardless of her success, it was plain to see how our chances of a successful recovery were greatly reduced when we violated two of the four basic rules.

Sue McMurray also states that a critical calibration reading should be taken from the launch control site while the rocket is on the ground just before launch.

In our rocket system, that would have been impossible due to the fact that the antenna was coiled up inside a metal airframe. If the antenna had been outside the rocket, the receiver could have been used to verify the proper function of the

125 system before launch. In addition, the tone heard while conducting this reading would have represented the distance from the launch control site to the launch tower. Then, during recovery eﬀorts, once the recorded tone was heard again, the user could safely assume that the rocket was within that predetermined distance from the user’s current position.

Sue McMurray also states that she had recovered over $ 500k in “lost” rockets and payloads with this technology. She said the biggest key to success with this system, outside of practice, is to turn your eyes off and only use your ears to find the “lost” rocket. Looking back on our practice sessions, the user had always known where the transponder had been placed prior to the exercise. This leads me to believe that on some psychological level the decision process to chase down certain blips over others was predetermined or affected by the user’s previous knowledge of the testing environment. Her searching method illustrates a scanning method where the user only listens to the tones coming through on the headset. While slowly sweeping one direction, the user listens for maximum intensity in the tone. At this point the eyes are opened and a point on the horizon is marked. Then the process is repeated while spinning slowly in the opposite direction. Then using the two points on the horizon

126 a third point is generated in the middle. This then serves as the heading which one should walk towards. After several hundred yards have been traversed, then a new heading should be determined. In the event that the two points on the horizon are too widely separated, the gain should be lowered until two narrow points on the horizon are representative of the maximum tone intensity.

Another realization taken from her tips and tricks was a discussion concerning the mounting of the antenna to the exterior of a carbon-glass tube. She recalled that the signal was very “one sided”, meaning that when the airframe landed in an awkward orientation, the signal was only detectable from one side of the airframe.

This left us with the thought that the implementation of multiple antennas on our system might have allowed us to embed the antenna in the exterior of the metal airframe while generating an omni-directional signal. However, this option was not researched, and it is unclear as too what ill side aﬀects this antenna array would have created. This array or another viable solution would have to be explored before another rocket launch was attempted.

127 5.3.3 Put Your Name On It

Endless discussions took place concerning design and construction of the rocket system as a whole. However, never once while we were discussing these engineering topics did it occur to us that it would be a good idea to put our names or even the university’s name on the rocket system as a precautionary measure in the event the system was not recovered.

On our rocket system, names only existed in two instances. The first instance was a laser etching in the original booster casing, which simply read “The University of Oklahoma” along with some additional use and safety warnings. Since this casing was supplied by Dr. Rocket and had not been qualified for use with our experimental custom motor, it had been set aside as a spare part and was not flown during the actual flight. The second instance was in the commercial altimeter which was purchased for the flight. There were two major problems with this. First, the altimeters were buried deep in the altimeter section of the rocket system. One would have to know that all 24 bolts needed to be removed in order to remove the altimeters and see name on the altimeter. The second problem was the fact that

128 the name wasn’t even visible on the outside of the altimeter. As a requirement for the purchase of a live detonating device, Adept Rocketry encodes the name of the customer in the ﬁrmware which is loaded on the ICs of the altimeter. This is done so that if the altimeter is used for ill intent and is recovered, it can be examined for evidence in criminal and civil litigation.

Regardless of what measures were taken, they weren’t enough. The solution to this is simple. At the very least, the university’s name and/or logo needs to be inscribed on everything possible. This would then solve the problem of someone

ﬁnding a lost rocket system and not knowing who to contact for its return. In addition, handling instructions should be etched on the exterior of the rocket, instructing people to not put their hands inside the parachute chambers where live charges may still be present.

5.3.4 Failure Is An Excellent Learning Tool

In the moments that followed the realization that a failure had occurred, sponta- neous brainstorming began to produce countless possible scenarios. The discussion

129 surrounding the validity of these scenarios was by far the most passionate, yet an- alytical discussion that the group had as a whole. Each scenario was dissected in detail and either classiﬁed as possible or impossible. In the days that followed, more discussions took place, facts were rehashed, and conclusions were drawn as to what actually happened. Given a choice, I cannot imagine any engineer would want to experience a failure for the sake of learning. By human nature, it seems counter- productive to seek out a failure purely for the opportunity to learn and grow. In addition, not many companies would openly spend money to put their employees in the position to learn from failures. This is not to say that failures don’t naturally occur; but failures, by nature, can be costly in terms of money and time. The loss on this experiment is arguably not acceptable. In all reality, approximately twenty thousand dollars in equipment and two years of construction eﬀorts were lost in a mere seventeen seconds. However, regardless of the loss, the most important thing you can do is to accept the failure and plot a course to resolve the issue. On the day of the launch, after discussing the possible failure modes, our attention quickly

130 turned to notifying the proper authorities in the surrounding area. To help the recovery eﬀorts, ads were placed in newspapers, and radio personalities passed along contact information.

5.3.5 Idealized Schematics

The design of the sensor suite and DAQ recorder for this project was a major under- taking. Initially it was thought that sensors and DAQ equipment would not be difficult to find and acquire. However, product searches quickly showed that ready made devices would be sizable and expensive. This left only one choice, pick components and build sensors which would be custom to our application. For the most part, the boards were easily designed, burnt, and populated. The exception to this statement concerns the design of the strain gage amplifier.

It was thought, early on, that a custom strain gage ampliﬁer design, which would suit our application perfectly, had been found in the product documentation of an

Analog Devices Operation Amplifier. However, it was not known that “idealized” schematics were simple representations of the actual circuits needed to recreate the publicized results. Hence, the first version of the strain gage amplifier was based

131 entirely on a schematic which didn’t contain all the components to function properly.

This left us to search out a complete circuit which functioned exactly as required.

The complexity of this circuit was the one item that pushed this project beyond the original scope of 3 months.

5.3.6 Parachute Color

Both parachutes were vibrant colors. However, it was difficult to distinguish the two shades, pink and red, from one another while the rocket was at altitude. In retrospect, the use of two different colored parachutes would have helped us determine which stage of the flight profile the altimeter thought it was experiencing. This would have assisted us in determining the root failure that led to the loss of the rocket system.

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[25] Vishay Measurements Group Inc. Tech note tn-515: Strain gage rosettes - selection, application, and data reduction. Technical report, 1990.

[26] Vishay Measurements Group Inc. Tech tip tt-602: Silver soldering technique for attachment of leads to strain gages. Technical report, 1983.

[27] Vishay Measurements Group Inc. Tech tip tt-603: The proper use of bondable terminals in strain gage application. Technical report, 1983.

134 [28] Vishay Measurements Group Inc. Tech tip tt-607: Strain gage installation and protection in ﬁeld environments. Technical report, 1983.

[29] Vishay Measurements Group Inc. Tech tip tt-609: Strain gage soldering tech- niques. Technical report, 1986.

[30] Vishay Measurements Group Inc. Tech tip tt-610: Strain gage clamping tech- niques. Technical report, 1992.

[31] William S. Janna. Introduction to Fluid Mechanics. PWS Publishing Company, Third Edition, 1993.

[32] Vern Knowles. Vern’s rocketry web site. URL http://www.vernk.com/.

[33] Sue McMurray. Walston Tips and Tricks, 1999.

[34] TFX-11 Remote Datalogger and Controller Engine User Guide. Onset Com- puter Corporation, 2000.

[35] Ray Preston. Formation of normal shockwave. URL http://142.26.194.131/aerodynamics1/High-Speed/Page2c.html.

[36] Ralph J. Smith and Richard C. Dorf. Circuits, Devices, and Systems. John Wiley & Sons, INC., Fifth Edition, 1992.

[37] Gary A. Crowell Sr. The Descriptive Geometry of Nose Cones, 1996.

[38] Jim Walston. Increase your odds in model retrieval. URL http://texastimers.com/helpful hints/walston/walston retrieval.htm.

[39] Anthony J. Wheeler and Ahmad R. Ganji. Introduction to Engineering Exper- imentation. Prentice Hall, First Edition, 1996.

135 Appendix A

Mechanical Drawings

All mechanical components used within this rocket system can be found on the enclosed CD-rom. The format for all part and assembly ﬁles is Pro/E Wildﬁre II.

136 Appendix B

DAQ Software

137 B.1 Master DAQ Code

//Master Pin-Out:

//List of Pins: //I/O 0: (Pin A18): LED#1 //I/O 1: (Pin A19): Sync Pin #2 //I/O 2: (Pin A20): LED#2 //I/O 3: (Pin A21): Sync Pin #1 //I/O 4: (Pin A22): LED#3 //I/O 5: (Pin A23): Launch Pin #1 - Set Low //I/O 6: (Pin A24): LED#4 //I/O 7: (Pin A25): Launch Pin #2 - Set High

//Global Variables ’n’ Stuff

//Input Channels //Channel 0 - (Pin A36): Strain Gauge - X1 //Channel 1 - (Pin A37): Strain Gauge - T1 //Channel 2 - (Pin A38): Strain Gauge - X2 //Channel 3 - (Pin A39): Strain Gauge - T2 //Channel 4 - (Pin A40): Accel - X //Channel 5 - (Pin A41): Accel - Z //Channel 6 - (Pin A42): Pressure - AB //Channel 7 - (Pin A43): Pressure - P2 //Channel 8 - (Pin A44): Pressure - P4 //Channel 9 - (Pin A45): Pressure - P6 //Channel 10 - (Pin A46): Temperature - TMP2

//------PRE LAUNCH------

//Use moment button to wake up if no main power

Backup: if BBPWR = 1 //If running on back up power, sleep Picint 1 //Set pin B20 to work as wake up hyb 65535 //Sleep for just over 18 hours if BBPWR = 1 //If still no power goto Backup //Go back to sleep endif endif

138 pset 0 //Turn on LED #1

//Create large array for storing data dim testArray(10300) arrayPointer = varptr (testArray)

//Store values of Z accel for 3G load plus or minus GLoadHigh = chan(5) + 1966 GLoadLow = chan(5) - 1966

//Adjust pins for prelaunch pclr 1 pclr 3 pclr 5 pset 7

//Store values in worthless variable to make launch connection work f = pin(5) g = pin(7)

//Wait for pin to be pulled pset 2 //Turn on LED #2 repeat sleep 0 sleep 200 pclr 5 pset 7 until( pin(5) = 0 )

//------STRAIN LOG------

//Sleep for 3 minutes pset 4 //Turn on LED #3 hyb 0 hyb 180

139 //----Sync Boards--- pset 3

//----Log Strain---- asm $

ADCTL equ $0030 ;A/D Control Register ADR1 equ $0031 ; Results Register 1 ADR2 equ $0032 ; 2 ADR3 equ $0033 ; 3 ADR4 equ $0034 ; 4 OPTION equ $0039 ; Power up the A/D STRBSR equ $fda6 ;Store to bank switched ram (@ array) pointer equ arrayPointer array equ testArray

;initialize pointer to array ldx pointer+2 stx 0,x ldd 0,x addd #D’10296 ;number = size of array - 4 std 2,x fcb H’18 ;force assembly of exchange d with y fcb H’8f inx inx inx inx

; power up the A/D Converter ldaa #H’80 staa OPTION

;wait for A/D power up ldaa #H’40 wait deca bne wait

; -----Wait for wire to burn through------

140 ldaa #21 staa H’00 test ldaa H’00 cmpa #21 bls test

;* set A/D to read first value convert ldaa #H’10 staa ADCTL

;wait for conversion ADL00 LDAA ADCTL ; Get the status BPL ADL00 ; Wait until complete, bit 7 is set

;save values ldaa ADR1 staa 0,x inx ldaa ADR2 staa 0,x inx ldaa ADR3 staa 0,x inx fcb H’18 ;force assembly of exchange d with y fcb H’8f std 0,x fcb H’18 fcb H’8f cpx 0,x bmi convert ;Branch back if x index is less than top of array end

Print "WE HAVE LIFT OFF!!!!!!!!!" Print "WE HAVE LIFT OFF!!!!!!!!" Print "WE HAVE LIFT OFF!!!!!!!"

//------LIFT OFF------

141 //----Resync boards----

//Let other board catch up sleep 0 sleep 10 //Wait 100ms ???? Too long? pset 1

//Goto Full Log repeat store #2,chan(0) store #2,chan(1) store #2,chan(2) store #2,chan(3) store #2,chan(4) store #2,chan(5) store #2,chan(6) store #2,chan(7) store #2,chan(8) store #2,chan(9) store #2,chan(10) until dfpnt >= 915750 //Until Memory full

//Store data logged at start //Store some junk between end of eeprom and ram for r = 0 to 99 store 59904 next r // print "Wrote Junk 1" sleep 0 sleep 10 // print "Writing RAM" //Store fastest log for a = 0 to 10299 //Array size - 1 sleep 0 //print peek(arrayPointer +a )," "; store peek(arrayPointer + a) sleep 4 next a stop

142 B.2 Slave DAQ Code

//SlavePinOut

//List of Pins: //I/O 0: (Pin A18): LED#1 //I/O 1: (Pin A19): Launch Pin #1 - Set High //I/O 2: (Pin A20): LED#2 //I/O 3: (Pin A21): Launch Pin #2 - Set Low //I/O 4: (Pin A22): LED#3 //I/O 5: (Pin A23): Sync Pin #1 //I/O 6: (Pin A24): LED#4 //I/O 7: (Pin A25): Sync Pin #2

//Use LED#3 and LED#4 to connect launch wire?

//Global Variables ’n’ Stuff

//Input Channels //Channel 0 - (Pin A36): Strain Gauge - Y1 //Channel 1 - (Pin A37): Strain Gauge - Y2 //Channel 2 - (Pin A38): Strain Gauge - X3 //Channel 3 - (Pin A39): Strain Gauge - Y3 //Channel 4 - (Pin A40): Strain Gauge - T3 //Channel 5 - (Pin A41): Accel - Y //Channel 6 - (Pin A42): Pressure - P1 //Channel 7 - (Pin A43): Pressure - P3 //Channel 8 - (Pin A44): Pressure - P5 //Channel 9 - (Pin A45): Temperature - TMP3 //Channel 10 - (Pin A46): Temperature - TMP1

//------PRE LAUNCH------

//Run on backup battery until main power comes on pclr 0 pclr 2 pclr 4 pclr 6

Backup: if BBPWR = 1 //If no main power

143 picint 1 //Set button to work hyb 65535 //And go to sleep if BBPWR = 1 goto Backup //if Still no power, go back to sleep endif endif

//Store some data at the start of the file for warmup for d = 0 to 25 store 000 next d

//Wakes up on button push //Show power on pset 0 //Turn on LED #1

//Create large array for storing data dim testArray(10300) arrayPointer = varptr (testArray)

//Adjust pins for prelaunch pin pull pset 1 pclr 3 pclr 5 pclr 7

//In order for pin pull to work, values must //be stored in dummy variables f = pin(1) g = pin(3)

//Wait for pin to be pulled pset 2 //Turn on LED #2 repeat sleep 0 sleep 200 pset 1 pclr 3

144 print "stuck" until( pin(3) = 0 )

//------STRAIN LOG------

//Sleep for 3 minutes to allow time to exit area pset 4 //Turn on LED #3 hyb 0 hyb 170 //Make sure this board wakes up first to look for sync signal

//----Sync Boards---- repeat print "Syncing!" until( pin(5) = 32 ) //32 is the value of pin 5 when it is set high

//----Log Strain---- asm $

;initialize pointer to array ldx pointer+2 stx 0,x ldd 0,x addd #D’10296 ;number = size of array - 4 std 2,x fcb H’18 ;force assembly of exchange d with y fcb H’8f inx inx

145 inx inx

; power up the A/D Converter ldaa #H’80 staa OPTION

;wait for A/D power up ldaa #H’40 wait deca bne wait

; -----Wait for wire to burn through------; Checks Pin 4 and 6 ldaa #53 ;21 + bit 5 staa H’00 test ldaa H’00 cmpa #53 bne test

;* set A/D to read first value convert ldaa #H’10 staa ADCTL

;wait for conversion ADL00 LDAA ADCTL ; Get the status BPL ADL00 ; Wait until complete, bit 7 is set

;save values ldaa ADR1 staa 0,x inx ldaa ADR2 staa 0,x inx ldaa ADR3 staa 0,x inx

146 fcb H’18 ;force assembly of exchange d with y fcb H’8f std 0,x fcb H’18 fcb H’8f cpx 0,x bmi convert ;Branch back if x index is less than top of array end

Print "WE HAVE LIFT OFF!!!!!!!!!" Print "WE HAVE LIFT OFF!!!!!!!!" Print "WE HAVE LIFT OFF!!!!!!!" //------LIFT OFF------

//----Resync Boards---- repeat until( pin(7) = 128 ) //Loops until signal from other board

//Store data logged at start //Store some junk between end of eeprom and ram for r = 0 to 99 store 59904 next r // print "Wrote Junk 1" sleep 0

147 sleep 10 // print "Writing RAM" //Store fastest log for a = 0 to 10299 //Array size - 1 //sleep 0 //print peek(arrayPointer +a )," "; store peek(arrayPointer + a) //sleep 4 next a pclr 0 pclr 1 pclr 2 pclr 3 pclr 4 pclr 5 pclr 6 pclr 7 stop

148 Appendix C

Electrical Schematics

Digital copies of the following schematics can be found embedded within the Protel database which is on the enclosed CD-rom.

149 C.1 Mother-Daughter Schematic

150 C.2 Strain Gage Ampliﬁer Schematic

151 C.3 Pressure Schematic

152 C.4 Accelerometer Schematic

153 Appendix D

PCB Layouts

Digital copies of the following PCB layouts, and the accompanying Gerber produc- tion ﬁles, can be found embedded within the Protel database which is on the enclosed

CD-rom.

154 D.1 Mother-Daughter Board (Top Layer)

155 D.2 Mother-Daughter Board (Bottom Layer)

156 D.3 Strain Gage Ampliﬁer Board (Top Layer)

157 D.4 Strain Gage Ampliﬁer Board (Bottom Layer)

158 D.5 Pressure Board (Top & Bottom Layers)

159 D.6 Accelerometer Board (Top & Bottom Layers)

160 Appendix E

Launch Day Log

Months prior to the launch day, the necessary paperwork was ﬁled with the Federal

Aviation Administration (FAA) to secure a floating flight wavier. This wavier could then be activated with only two weeks notice before the desired launch date. Upon activation the wavier would effectively designated the airspace specifically above the

Sayre Municipal Airport as a no-ﬂy zone for all air traﬃc private or commercial from ground level to a predetermined altitude. After countless testing of the rocket system, this launch wavier was activated for the weekend of November 29 & 30,

2003.

The launch weekend began with the transportation of the rocket systems and the necessary support equipment to the launch site. To ease the stress of travel and transportation of the equipment, the team stopped to spend the night in Elk

161 City, Oklahoma on Friday, November 28, 2003. This stop placed the team within approximately 20 of the Sayre, Oklahoma destination.

Early in the morning, around sun rise on Saturday November 29, 2003, the team made its way to the Sayre Municipal Airport. Upon reaching the airport, the team was eﬀectively divided into two groups such that tasks could be completed in a concurrent manner. The ﬁrst group focused on the erection of the launch tower while the second group focused on the pre-launch preparation of the rocket.

The launch tower erection process was very simply in nature. First concert blocks and plywood were arranged on the ground to serve as footings for the scaﬀolding.

This proved to be necessary due to the fact that it had rained the previous two days and the excess precipitation had reduced the grounds ability to support the weight of the launch tower structure. This problem was further compounded by the fact that the weight of the structure was focused on the foot print of four small wheels which were at the ends of the first tier scaffolding legs. Once the first tier was erected, the tower was leveled by adjusting the jack screws at the bottom of each leg. The following tiers were then added to the structure. The two sections of radio tower were assembled while laying flat on the ground. This allowed the use of a hammer

162 to gently persuade the two sections to slide together. While the radio tower was still on the ground, the launch rail was assembled piece by piece, starting at the bottom section of the radio tower. Then, as a whole unit, the radio tower was then lifted into place on the scaffolding and held there with eight large u-bolts. After securing the radio tower, the launch rail was checked by visual inspection to make sure that each of the four sections created a straight path. After adjustments to the launch rail, a spare launch lug was dropped several times from the top of the rail to insure that there were no snags in its path during liftoff. Two guide ropes were attached to the structure to increase the overall stability of the system. At ground level, all combustible materials were removed within a fifty foot radius. In addition, water was poured on the ground to wet the grass and several water buckets were placed near to the launch site such that a fire could be put out in emergency situations.

The last item needed to complete the launch tower was the addition of a banner proudly displaying the crimson and cream logo of the University of Oklahoma.

The pre-launch preparation of the rocket was much more complex then the erection of the launch tower. The ﬁrst section prepped was the booster section. The upper bulkhead contained an o-ring. This o-ring and the upper internal surface of

163 the booster casing were lubed with white lithium grease. The bulkhead was then inserted into the booster casing and secured with an internal e-clip ring. To ease installation of the motor, the exterior cardboard casing of the motor was coated with white lithium grease. Then the motor assembly was then loaded from the open bottom end of the booster casing. The graphite exhaust nozzle also contained an o-ring groove. This o-ring was coated with white lithium grease and the exhaust nozzle was inserted into the booster casing. The aluminum relief ring was then slid into place and secured with another internal e-clip ring. The ﬁn can then slid down the exterior of the booster casing and was threaded into place. The booster section was then placed aside away from the current work area for safe keeping until ready for ﬁnal assembly on the launch pad.

Preparation of the altimeter section followed. Before installation of the altimeter cradle, the three altimeters were connected to the deflagration charges, one charge in each parachute chamber for each altimeter. The altimeters were then turned on to verify continuity of each firing squib. After verification, the altimeters were powered back down and the firing safeties were removed. The altimeters were then turned back on and given sixty seconds to obtain a base pressure altitude reading. White

164 lithium grease was applied to the o-ring grooves of the parachute plugs; these parachute plugs had previously been assembled to the altimeter cradle. This assembly was then installed in the altimeter casing and the main parachute plug was fastened with twelve screws. The remaining coupler was then slid into place and, together with the drogue parachute coupler, was attached with twelve more screws. Baby powder was then put inside each of the parachutes just before the chutes were rolled up and installed inside the altimeter casing. The altimeter section, ready for launch, was then place next to the booster casing for safe keeping.

The simplest section to prep was the instrumentation section. A pin was simple inserted into an access hole to toggle a power switch.

All three sections were then carefully transported to the launch tower in the bed of a pick up. Once at the tower, the altimeter section was slid onto the launch rail and held in place by hand. The booster section was then slid onto the launch rail just under the altimeter section. The block which would support the rockers weight was then installed at the bottom of the launch rail under the booster section.

The booster section and the altimeter section were then slid together and allowed to rest on the support block at the bottom of the launch rail. The instrumentation

165 section was then installed on the top of the altimeter section. At this point all the surrounding equipment was moved away from the launch tower. In addition, spectators were asked to remove themselves to the designated viewing areas. Once the area was clear, the igniter was installed in the motor. At this stage, only two people were allowed to be within the 1000 ft safety radius. The ignition system was attached to the igniter and the remaining team members retreated to the launch control area. At approximately 1:45 pm, after a ﬁnal safety check and a ten second count down, the rocket was launched.

The overall ascent of the launch went as was expected. The rocket left a white smoke trail for the duration of the motor burn. This trail helped in the visual tracking of the rocket. A few seconds after the smoke trail ended, a visible white cloud could be seen. The cloud was produced by the baby powder released during the deployment of the parachutes. At this point a couple of team members maintained visual contact with the parachute while a couple other members began to search the sky with the transponder tracking equipment. First attempts to track the rocket electronically proved to be misleading due to the fact that ”blips” heard on the tracker did not come from the same directions as the visible parachute. Therefore,

166 for the fear that something at the launch site was interfering with the electronics, the trackers began to move from roadside site to roadside site with the hopes that a new venue would produce a good signal. Visual contact with the parachute only lasted for about 45 minutes, at which point the parachute was simply no longer visible with the naked eye. Therefore, crew at the launch site quickly turned their focus to the disassembly of the launch tower even though the tower had been erected in a safe area of the Sayre Municipal Airport complex. This was prompted by the fact that the launch tower had no warning lights, nightfall was beginning to draw near, and the last thing the team needed was an accident involving a private plane. While the launch site crew began to disassemble the tower, the tracking crew continued their search for approximately 2 hours.

When the tracking crew returned to the launch site, a disappointing realization began to set in We had an unknown failure and the rocket was missing.

167