METEOR - EDGEedge.rit.edu/edge/P07109/public/Design I/Published... · 8.2.3.2 Pyrotechnic Ignition...

METEOR Microsystems Engineering and Technology

for the Exploration of Outer Regions

Senior Design Project 06006:

“Design and Testing of a Small Scale Rocket for Pico-Satellite Launching.”

Dr. Jeffrey Kozak Dr. Dorin Patru Project Advisor Project Advisor David Dale John Chambers Chris Hibbard Mechanical Engineer Mechanical Engineer Mechanical Engineer Project Manager Jeff Nielsen Jessica LaFond Anthony Fanitzi Mechanical Engineer Mechanical Engineer Mechanical Engineer Daniel Craig Brad Addona Mechanical Engineer Mechanical Engineer

Table of Contents

1. Introduction / Motivation 12. Organization / Team Breakdown 7 2.1 Senior Design Project 06006 Scope 7 2.2 Team Breakdown 8 2.3 Project Planning 83. Literature Review 10 3.1 Introduction 10 3.2 Similar Projects / Missions 10 3.2.1 Pegasus 10 3.2.2 Minotaur Designs 11 3.2.3 AspireSpace 12 3.2.4 University of Colorado at Boulder 12 3.2.5 Air Launched Flight Trajectories 13 3.3 Hybrid Rocket 14 3.3.1 Introduction / Benefits of Hybrid Rocket 14 3.3.2 Propellants 14 3.3.3 Performance Predictions 14 3.3.4 Fuel Chamber 17 3.4 Feed System 17 3.5 Ignition System 19 3.6 Injector 19 3.7 Exit Nozzle 20 3.8 Federal Specifications 204. Needs Assessment 21 4.1 Performance Goals 21 4.2 Design Goals 21 4.3 Secondary Goals 215. Specifications / Success Qualifiers 226. Concept Development / Feasibility 26 6.1 Hybrid Engine 26 6.1.1 Propellant Selection 26 6.1.1.1 Solid Fuel Selection 26 6.1.1.2 Liquid Oxidizer Selection 26 6.1.2 Fuel Grain / Propellant Sizing 28 6.1.2.1 Propellant Mass 28 6.1.2.2 Regression Rate 29 6.1.2.3 Mass Flow Rate / Fuel Grain Sizing 30 6.1.3 Fuel Chamber Intended for Ground Testing 31 6.1.3.1 Basic Configuration 31 6.1.3.2 Obtaining Required Data 32 6.1.4 Injector Design 34 6.1.4.1 Basic Concept 34 6.1.4.2 Machining / Assembly 36 6.1.5 Ignition System 37 6.1.6 Feed System 43 6.1.6.1 Basics 43 6.1.6.2 Oxidizer Tank Filling 43 6.1.6.3 Nitrogen Tank Loading 44

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6.1.6.4 Oxidizer Properties 45 6.1.6.5 Valves 45 6.1.7 Exit Nozzle 49 6.1.7.1 Nozzle Shaping 49 6.1.7.2 Nozzle Materials 50 6.1.7.3 Nozzle Attachment 50 6.2 Test Stand 51 6.2.1 Introduction 51 6.2.2 Basic Configuration 51 6.2.3 Safety Considerations 55 6.2.3.1 Built in Redundancies 56 6.3 Paper Design 57 6.3.1 Introduction 57 6.3.2 Aluminum Truss System 58 6.3.3 Satellite Base Plate 58 6.3.4 Thrust / Injector Plate 59 6.3.5 Satellite Release Mechanism 59 6.3.6 Satellite Containment 60 6.3.7 Hybrid Engine Configuration 637. Engineering Analysis / Design Validation 64 7.1 FEA Analysis 64 7.1.1 Introduction 64 7.1.2 Test Stand Analysis 64 7.1.2.1 Materials 64 7.1.2.2 Loading and Restraints 65 7.1.2.3 Study Properties 66 7.1.2.4 Stress Results 67 7.1.2.5 Strain Results 68 7.1.2.6 Displacement Results 69 7.1.2.7 Design Check Results 70 7.1.2.8 Conclusion 71 7.1.3 Combustion Chamber Analysis 72 7.1.3.1 Materials 72 7.1.3.2 Loading and Restraints 72 7.1.3.3 Study Properties 73 7.1.3.4 Stress Results 74 7.1.3.5 Strain Results 75 7.1.3.6 Displacement Results 76 7.1.3.7 Design Check Results 77 7.1.3.8 Conclusion 78 7.2 Rocket Calculations 79 7.2.1 Rocket Sizing and Thrust Calculations 79 7.2.1.1 Givens and Assumptions 79 7.2.1.2 Mass Estimation 79 7.2.1.3 Regression Rate 80 7.2.1.4 Sizing 80 7.2.1.5 Oxidizer Mass Flow Rate 81 7.2.2 Exit Nozzle Shaping 81 7.2.2.1 Given and Assumptions 81 7.2.2.2 Area Sizing 81

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7.2.2.3 Exit Velocity 838. Senior Design II 84 8.1 Deliverables 84 8.1.1 Primary Deliverables at End of SD II 84 8.1.2 Secondary Objectives 84 8.2 Future Plans 85 8.2.1 Time Table 85 8.2.2 Propellant Testing 85 8.2.2.1 Molding 85 8.2.2.2 Combustion 85 8.2.3 Ignition Testing 86 8.2.3.1 Glow Plug Ignition 86 8.2.3.2 Pyrotechnic Ignition 86 8.2.4 Data Acquisition Component Testing 86 8.2.5 Feed System / Oxidizer Flow 86 8.2.6 Combustion Chamber Tests 87 8.2.7 Test Stand Materials 87 8.3 Testing the Rocket 87 8.3.1 Test Results 87 8.4 Budget 88References 89Appendices 91 Appendix 1 - Drawing Package Appendix 2 - Bill of Materials Appendix 3 - Gantt Chart SD I Appendix 4 - Risk Assessment Appendix 5 - Objective Trees Appendix 6 - Timeline SD II Appendix 7 - Rocket Calculations Appendix 8 - Regression Rates Appendix 9 - Rocket Nozzle Calculations Appendix 10 - Feed System Schematic Appendix 11 - Ni-Chrome Wire Temperature Properties Appendix 12 - Safety Report Appendix 13 - MSDS Sheets

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List of Figures Figure 1 Launch System Block Diagram………………………………………………….4 Figure 2 Typical Mission Profile………………………………………………………….5 Figure 3 Work Breakdown Structure……………………………………………………...8 Figure 4 Typical Pegasus Mission Profile……………………………………………….11 Figure 5 Minuteman Launch Vehicle Configuration…………………………………….12 Figure 6 Cross-sectional View of Test Chamber………………………………………...31 Figure 7 Swirling Nozzle………………………………………………………………...34 Figure 8 Spray Pattern for Various Gasification Injection Gas Ratios…………………..35 Figure 9 Initial Injector Concept Drawing……………………………………………….36 Figure 10 Final Injector Concept Cross-section…………………………………………37 Figure 11 Pure Oxygen Ignition System Schematic……………………………………..38 Figure 12 Glow Plug……………………………………………………………………..39 Figure 13 Glow Plug Ignition System Schematic………………………………………..40 Figure 14 Pyrotechnic / Ni-Chrome Igniter……………………………………………...41 Figure 15 Pyrotechnic Ignition System Schematic………………………………………43 Figure 16 Oxidizer / Nitrogen Tank Fill Schematic……………………………………..44 Figure 17 Concept for Testing Within Test Cell………………………………………...52 Figure 18 Horizontal Test Stand Concept……………………………………………….54 Figure 19 Preliminary Drawing of Chosen Test Stand Design………………………….55 Figure 20 Front and Isometric View of Aluminum Truss System, Satellite Base Plate,

Thrust Plate and Satellite Release Mechanism Assembly……………………….58 Figure 21 Pressurized Satellite Release Mechanism…………………………………….60 Figure 22 Pico-Satellite Containment and Release Schematic…………………………..62 Figure 23 Hybrid Rocket Configuration…………………………………………………63 Figure 24 Stress Distribution on Test Beam……………………………………………..67 Figure 25 Strain Distribution on Test Beam……………………………………………..68 Figure 26 Displacement of Test Beam…………………………………………………...69 Figure 27 Yield Factor of Safety Distribution on Test Beam……………………………70 Figure 28 Ultimate Factor of Safety Distribution on Test Beam………………………...71 Figure 29 Stress Distribution on Test Chamber………………………………………….74 Figure 30 Strain Distribution on Test Chamber………………………………………….75 Figure 31 Dislocations of Test Chamber………………………………………………...76 Figure 32 Cross Sectional View of Yield Factors of Safety in Test Chamber…………..77 Figure 33 Cross Sectional View of Ultimate Factors of Safety in Test Chamber……….78

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List of Tables Table 1 Experimentally Determined Correlation Parameters……………………………16 Table 2 Weighted Method Analysis of Hybrid Engine………………………………….24 Table 3 QFD Analysis of Hybrid Engine………………………………………………..25 Table 4 Pugh Analysis of Solid Fuel Propellants………………………………………..27 Table 5 Correlation Parameter Table – Regression Rates……………………………….29 Table 6 Pugh Analysis of Oxidizer Injection Methods…………………………………..36 Table 7 Pugh Analysis of Ignition Systems……………………………………………...38 Table 8 Pressure of Nitrous Oxide Based on Temperature [15]…………………………45 Table 9 Control Valve Sizing Calculation……………………………………………….48 Table 10 Injector Orifice Sizing Calculation…………………………………………….49 Table 11 Pugh Analysis of Rocket Test stands………………………………………......53 Table 12 Test Stand Assembly Parts…………………………………………………….64 Table 13 Material Properties of AISI 1020 Steel………………………………………...65 Table 14 Material Properties of 4140 Annealed Steel…………………………………...65 Table 15 FEA Analysis Mesh Information………………………………………………66 Table 16 FEA Analysis Solver Information……………………………………………..66 Table 17 Location of Maximum and Minimum Stress on Test Beam……..……………67 Table 18 Location of Maximum and Minimum Strain on Test Beam…………………...68 Table 19 Location of Maximum and Minimum Displacement on Test Beam…………..69 Table 20 Materials Used in Test Chamber………………………………………………72 Table 21 Material Properties of AISI 304 Stainless Steel……………………………….72 Table 22 FEA Analysis Mesh Information of Test Chamber……………………………73 Table 23 FEA Analysis Solver Information of Test Chamber…………………………..73 Table 24 Location of Maximum and Minimum Stress on Test Chamber……………….74 Table 25 Location of Maximum and Minimum Strain on Test Chamber……………….75 Table 26 Location of Maximum and Minimum Displacements on Test Chamber……...76

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Nomenclature

Acronyms: METEOR: Microsystems Engineering & Technology for the Exploration of Outer Regions LEO: Low Earth Orbit DAQ: Data Acquisition MEMS: Micro-Electro-Mechanical Systems NASA: National Aeronautics and Space Administration OSCAR: Orbiting Satellite Carrying Amateur Radio RIT: Rochester Institute of Technology START: Strategic Arms Reduction Talks HTPB: Hydroxy-terminated Polybutadiene (Solid rocket fuel) PMMA: Polymethyl Methacrylate (Solid rocket fuel) AGI: Analytical Graphics Inc. STK: Satellite Tool Kit NC lacquer: Nitrocellulose lacquer FAA: Federal Aviation Administration MIL: Military standard prefix O2: Oxygen F2O: Oxygen Difluoride H2O2: Hydrogen Peroxide N2O: Nitrous Oxide SAE: Society of Automotive Engineers EDM: Electron Discharge Machining BP: Black Powder NH4ClO4: Ammonium Perchlorate C-D: Converging-diverging TODOR: A flow software program FOS: Factor of Safety FEA: Finite Element Analysis Variables: *All units are in Metric and can be converted to English if desired. Isp is specific impulse [sec] vtotal is total velocity required to achieve low earth orbit (LEO) [m/s] mL is payload mass (mass of the satellite) [kg] ms is structural mass of the rocket [kg] melec is mass of the electronics in the rocket [kg] di is the inner diameter of the circular fuel grain [m] d0 is the outside diameter of the fuel grain [m] L is the total length of the fuel grain [m] ρHTPB is density of solid fuel grain HTPB [kg/m3] g is the acceleration of gravity [m/s2]

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ve,Ve is exhaust velocity of the rocket [m/s] Δv is the change in velocity of the rocket [m/s] N is the number of stages of the rocket [--] m0 is the initial total mass of the rocket [kg] m1 is the final total mass of the rocket [kg] mp is the total propellant mass [kg] Ea is activation energy [kJ/mol] A is the Arrhenius preexponential constant [mm/s] Ts is the surface temperature of the fuel grain [K] R is the universal gas constant [J/(mol-K)] M is molecular weight of HTPB [kg/mol] r is the solid propellant regression rate [mm/s] VHTPB is the volume of HTPB burned in 1 second [m3]

HTPBm& is the approximate mass flow rate of HTPB [kg/s]

HTPBm is the total mass of HTPB [kg] t is the amount of time [sec] tb is the amount of time for the fuel to be completely burned [sec]

oxm& is the oxidizer mass flow rate [kg/s] Go is the oxidizer mass flux [kg/m2-s] x is axial location in the fuel grain [m] k is the gas absorption coefficient [(m-MPa)-1]

p is the pressure [MPa] h is the port height between fuel slabs [m] n,m,k,C1,C2 are all parameters developed by Chiavirini for HTPB m& is total mass flow rate of the propellant [kg/s] pe is exit pressure [Pa] pa is ambient pressure [Pa] Ae is exit area of the nozzle [m2] τ is thrust [N] Cd is Orifice Coefficient Ao is Orifice Area [m2] ρ is Fluid Density [kg/m3] ΔP is Pressure Drop Across Orifice [Pa] mf is total fuel burned [kg] T0 is stagnation temperature [K] P0 is stagnation pressure [Pa] γ is the ratio of specific heats Tt is nozzle temperature [K] Pt is nozzle pressure [Pa] Mgas is the molecular weight of the exiting gasses [kg/kmol] At is throat area [m2] Dt is throat diameter [m] Me is exit Mach number

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About the METEOR Rocket Preliminary Design Report

The following preliminary design report details the process our design team took

to develop, evaluate and select the overall design concepts to meet our objectives outlined

at the conclusion of this paper. The report begins with an introduction to the history of

and the motivation behind the METEOR (Microsystems Engineering & Technology for

the Exploration of Outer Regions) project here on RIT’s campus. A description of the

scope of our portion of this bold mission will follow describing the organization and

breakdown of our approach to solving our objectives. The next section of our paper is a

brief description of the elaborate literature review our team conducted in order to come

up to speed on current technologies and similar projects.

With a foundation of knowledge established, the report then discusses the needs

assessment and outlines the goals and desirables that our team will accomplish by the end

of Senior Design II. After establishing our design goals the paper goes on to discuss the

specifications and success qualifiers that we need to achieve in order to prove that we

have successfully completed our design and performance goals.

After discussing the team organization and breakdown of the approach to solving

the problems faced by our team the paper focuses on the heart of our concept

development and design work. In depth conversation on concept development and the

feasibility of each major design portion of our project is outlined in Section 6 of our

paper. After much discussion on concepts, a summary of our final design and validation

for our choices is given. This portion is then followed by a description of our plans for

Senior Design II and where we plan to take our project over the next three months.

1. Introduction / Motivation A brief scenario taking place in today’s world…

A small group of intelligent and innovative young engineers have been working

together for the past couple years putting in their limited time and hearty effort to develop

a small satellite with Micro-Electro-Mechanical Systems (MEMS) technology. The

intention of this satellite is to better understand the effects of pollution in our outer

atmosphere and potential risks that we may be headed for collectively as a human race.

Making the satellite on such a small scale allows the engineers to construct several

Senior Design P06006 Page 1

satellites with the intention to examine many different orbit inclinations at different

locations around the world. They have put in years of hard work to develop reliable

technology and are finally ready to test their theories and try to improve the

understanding of our world.

The team then takes their satellite to industry, to NASA, to anybody that has the

capability of reaching space and gives them a well thought out and thorough presentation

on the validation of their work. All the institutions are thoroughly impressed and are

onboard with their ideas, however, and unfortunately they do not have the capability to

launch a satellite so small; or it would cost the team upwards of $500,000 to 1 million

dollars to reserve a flight onboard a rocket designed to carry cargo 100 times its size.

Another option presented to them is to wait 2-3 years and they could possibly reserve a

spot on a prototype air force rocket, but of course there is no guarantee that it will be

delivered successfully to lower earth’s orbit. Until they finally discover a program that

has been running on RIT’s college campus in Western NY for a number of years lead by

generations of groups of undergraduate students with similar amateur ambitions. They

are able to launch their satellite to any inclination at about 1/10 the cost they would have

paid, and as frequently as once a month if they wish. Their aspirations are realized and

their research will go on to help scientists better understand exactly which chemicals are

causing damage to our ozone.

Historical Background

Barely four years after the first American satellite was launched, on December

12th 1961, the first Orbiting Satellite Carrying Amateur Radio, OSCAR I, was

successfully launched and orbited the earth for 21 days. OSCAR I was launched

piggyback, as a secondary payload on an Air Force rocket. It carried a 144 MHz beacon

and weighed only 4.5 kg [1].

Over the last half a century there has been unprecedented advances in technology,

in both space travel and miniaturization of technological devices. Contradictory to

expectations, the cost of space travel has essentially remained the same, effectively

making non-for profit launched satellites have to continue to hitch hike their way to space

as secondary missions aboard expensive government rockets. This is where project


METEOR is basing their reasoning for pursuing such a mission and providing a means

for the academic world to reach outer-space cheaply and efficiently.

What is a Pico-Satellite?

A Pico-Satellite is a satellite with a weight less than 1 kg or about 2 pounds. Such

a satellite can incorporate a beacon transmitter, or a transponder, and/or a video camera,

or any other miniaturized scientific instrument which would fit within the specified

weight limit. With current and future advances in MEMS technology it appears that

these small scale satellites may be capable of performing operations once designated for

much larger and more expensive satellites and launch missions.

Current Options for Pico-Satellite Launch

Current launch vehicles are designed to carry payloads from 100 to 6000 kg to

various orbits. Orbital Sciences Inc. offers an air-launched rocket, Pegasus, which is

designed to carry a payload with a minimum weight of 285 kg to lower earth’s orbit

(~160 km to 400 km) [2]. The project Cubesat [3], lead by CalPoly and Stanford

Universities, uses a mother satellite to carry several Pico Satellites. After the mother

satellite reaches the desired orbit it releases the individual Pico Satellites which then

function on their own. Although the satellites weigh less then 1 kg each the cost for such

a mission is $80,000. In addition the typical wait time to get aboard this mission is 3-5

years, based on the completion of each of the individual satellites. More than 40 different

high school and university teams worldwide are building satellites for this project and

that number is growing. In rare instances teams are lucky enough to reserve a piggy back

spot as a secondary payload aboard a governmental test flight. In these cases the flight

cost is actually free, but obtaining the rights to join these missions is along process and

almost never works out for small amateur groups.

Description of Project METEOR

The METEOR project is currently in its initial phase, which consists of designing

a platform capable of reaching an altitude of about 80,000 ft utilizing helium filled, high-

altitude balloons. This platform will eventually be the launching point for future space


bound missions. The design team is working on the program through RIT’s multi-

disciplinary capstone design project and is composed of 5th year Electrical and

Mechanical engineering students who are enrolled in senior design.

The proposed launch configuration is pictured in Figure 1 “Launch System Block

Diagram” (not to scale). Current balloons can rise above 30,000 m and float for extended

periods of time, from hours to days, with p

as heavy as 1000 kg [4]. The idea of an air

launched rocket materialized in the early ‘90s

with Orbital Sciences’ first aircraft launched

Pegasus rocket. The idea of a balloon based

launch is not new either, but to date has never

been successfully completed. This flight pattern

is not feasible for satellites or payloads on the

order of 100 kg or more, however launching a 2

– 400 kg rocket with a 1-5 kg satellite is withi

the realm of possibilities.

ayloads

00

n

A generic mission profile is illustrated in

Figure 2 “Typical Mission Profile”. The balloon

and its payload are launched and reach rocket

launch altitude after approximately one hour.

After the platform passively stabilizes, the rocket

is turned and oriented in the right direction. Once

this is achieved, ignition of the first stage occurs

and the rocket leaves the platform. Orientation

accuracy of the rocket at the time of ignition can

be within +/- 3 degrees, as the rocket guidance system should be able to correct later for

the difference. Thereafter, the sequence of events is similar to any conventional rocket

launch. The platform is recoverable via the parachute tether located between the balloon

and platform, while the rocket stages are expendable.

Fig. 1 Launch System Block Diagram: (1) Balloon (2) Tether (3) Parachute (4) Stabilization Tethers (5)

Launch Platform (6) Rocket Suspension Lines (7) Rocket -

Satellite


Why Choose METEOR Over Other Technologies?

The proposed launch system has the following advantages:

• No need for ground infrastructure, except for an approximately seven person

mobile launch control unit, which can be located in the back of a van

• Launch location is only limited by safety range issues over populated areas

• Launching from such altitudes virtually eliminates atmospheric drag during the

phase of powered flight; the density of air above 80,000 feet is less that 1% that

on the surface of the earth

• Maximum dynamic pressure will be very low, resulting in a more relaxed

structural design; further supported by the fact that the low weight of the payload

will cause less vibration and g-force issues

• Weather conditions do not affect the phase of powered flight to orbit

• Rocket motors operate in a virtual vacuum at all times; thus exit nozzles can be of

fixed optimized dimensions for highest efficiency

• Launch frequency can be as often as once per month

Fig. 2 Typical Mission Profile


Current Situation of Project METEOR at RIT

A past Senior Design team has completed the basic design, and built a prototype

for the platform designed to carry a rocket to 80,000 feet. Since that time two successful

launches have been completed with smaller balloons attached. The platform reached an

altitude of approximately 30,000 feet before it returned safely to the ground, the entire

time taking live video feed and keeping contact with the team on the ground.

Another Senior Design Team which completed Senior Design I in the fall 2005

and will be entering Senior Design II at the same time as our team in the spring 2006, has

updated and improved the platform. They have added several electronic devices and

intend to give the platform the capability to orient itself prior to launching the rocket.

Currently there are also two EE graduate students that are working on inertial and

navigation systems for the rocket and updating the technologies that will be placed on the

Pico Satellite. Both of these students are putting in extensive time and effort to ensure

that the rocket and satellite will perform correctly during and after the powered flight

stage.

Our team is intended to lay the ground work for future teams that will focus on

the rocket design portion of the METEOR project. We intend to gain experience with the

propellants used and provide enough information that future teams can accurately size

and predict the nature of the rocket launch vehicle.


2. Organization / Team Breakdown

2.1 Senior Design Project 06006 Scope Based on preliminary work done by Dr. Patru and other students in order for a rocket

to carrying a 1 kg satellite to lower earth’s orbit (LEO) and obtain a velocity of 7600 m/s

it is expected that it will need four stages making the following assumptions [5]:

• A redundant weight of 10%; structural weight is only 10% of the total weight of

the rocket

• Isp = 235 sec

• A loss of 1600 m/s due to drag loss, gravity loss, maneuvering and launch

window allowance

Over the next couple of years the METEOR project aspires to complete successful sub-

orbital launches with single stage rockets, then making the transition to putting a multi-

staged rocket into LEO.

The scope of our project includes ground testing hybrid rocket propellants and

obtaining as much experience and data as possible in order for future teams to properly

estimate performance and the size of future rockets. We are also responsible for a

conceptual paper design of the upper stage of this rocket, and develop a manner to

contain and release a Pico Satellite into orbit. Excluded from this paper design are

navigational controls and thrust vectoring which are being examined by current graduate

students on campus. Much of our paper design is dependant on the data we are able to

collect by testing a prototype hybrid rocket on the ground, therefore much of the material

selection and sizing procedures may be limited by the time frame allotted to our group.


2.2 Team Breakdown Understanding hybrid rocket engines and predicting there performance required a

lot of background research and investigation into similar projects. Also establishing a

safe and accurate testing procedure has required extensive investigation by our team. In

order to tackle a project with such a wide scope we have broken the team up into the

following focus areas:

METEOR Rocket Design

P06006

Hybrid Engine Lead / Exit

Nozzle Design

Propellant Selection / Data

Acquisition

Feed System / Test Set-up

Ignition / Materials

Project Manager /

Safety Issues

Test Stand / Test Procedure

FEA / Paper Design

John Chambers

Chris Hibbard & Jessica LaFond

Brad Addona Dan Craig Anthony Fanitzi

Jeff Nielsen David Dale

Fig. 3 Work Breakdown Structure

While this figure illustrates the leader in each area of research, decisions on

overall system design are made as a group and evaluated with the Pugh method.

Additionally it should be noted that while the above individuals are the “experts” on the

specific assignments, responsibilities will change throughout the progression of the

project as work load dictates. Furthermore, communication will remain open between all

members as to better facilitate the synergy of components in the overall system design.

2.3 Project Planning Based on work completed by David Dale in Design Project Management in the

fall of 2005, a tentative and malleable Gantt Chart was created in order to give a timeline

and short term goals for the team (Appendix 3 – Gantt Chart). The team has adhered to

the schedule laid before them as well as can be expected, which has allowed us to

confidently present to you, the reader, our design.


In addition to the Gantt Chart a Risk Assessment was performed in order for our

team to gauge exactly what could be accomplished in Senior Design I and II (Appendix 4

– Risk Assessment). The different aspects of our design were weighed against four main

criteria (1) Resource Feasibility, (2) Economic Feasibility, (3) Schedule Feasibility, and

(4) Technological Feasibility.


3. Literature Review

3.1 Introduction With the team structure established and a tentative timeline before us the first step

our team needed to take was to conduct a comprehensive literature search to better

understand current technologies and establish a good base of knowledge in the specific

areas of our design. The focus of our search can be broken down into the following

sections: similar projects, hybrid rocketry, test stands and data acquisition, feed system,

materials, and federal specifications. The following sections outline our findings.

3.2 Similar Projects / Missions 3.2.1 Pegasus

Considering that Orbital Sciences Inc. was the first to successfully complete an air

launched flight to orbit by way of the Pegasus rocket we decided to find out much

information on their chosen flight patterns and methods of achieving desirable orbital

Fig. 4 – Typical Pegasus Mission Profile


altitudes and velocities. The most useful document we could find on Pegasus was the

“Pegasus Users Guide” [2] which gave detailed descriptions of the capabilities and

typical missions that are completed by Pegasus for its customers. Figure 4 illustrates a

typical Pegasus mission profile.

In addition to reviewing the Pegasus User’s Guide we found several other papers

that reference the Pegasus rocket. One in particular introduced us to some possible

scaling factors for the Pegasus rocket based on smaller payload sizes [6]. In this

downsizing process, the fraction of mass for each stage to total vehicle mass is

maintained, and the mass of each component is determined by using a scale factor that

follows a cubic scaling law. This law infers that (for the same average density) the ratio

of scaled-down mass to full-size mass is given by the cube of the scale factor; (e.g. if the

scale factor is one-half, the mass ratio is one-eighth). But because it is unlikely that the

avionics and the attitude control hardware can be scaled in the same manner, there would

have to be some adjustments made to the downsized payload to compensate.

3.2.2 Minotaur Designs

Looking for similar projects in the academic field we came across a design team

from the University of Texas at Austin, which completed a design report titled

“Converting the Minuteman Missile into a Small Satellite Launch System” [7]. The

design team devised a method to convert 450 Minuteman II Intercontinental Ballistic

Missiles that had been recently taken out of service as part of the Strategic Arms

Reduction Talks (START) peace treaties between the United States and the Ex-Soviet

Union, into launch vehicles for small satellites. These missiles are still on a much larger

scale than we would like to achieve, but this paper gave us much insight into how to

contain and release satellites safely into orbit. I would strongly suggest for future teams

to refer to this paper, especially pertaining to staging release and structural

considerations. Figure 5 illustrates the configuration of the Minuteman Launch Vehicle.


Fig. 5 – Minuteman Launch Vehicle Configuration

3.2.3 AspireSpace

While conducting a search for hybrid engines our team came across a group of

engineers, scientists and entrepreneurs who, in their spare time, build and operate small

sounding rockets. The team calls themselves AspireSpace [8]. AspireSpace is based in

the United Kingdom, and their main goal is to revamp the English space program and

bring young students into their organization. The paper principally covers the

development of AspireSpace’s first hybrid rocket engine, the H2, operating on Nitrous

Oxide and Polyethylene it had an impulse of up to 1800Ns and a thrust of 600N. This

paper outlined some of the basics behind hybrid rocket technology and gave us some

insight into methods used to test hybrid rockets. Our team is pursuing connection to

AspireSpace and I would recommend future teams to attempt to keep a dialog with this

group and try to absorb as much information as possible from them, as they too began as

a novice group of engineers.

3.2.4 University of Colorado at Boulder

Near the end of Senior Design I our team came across a technical paper prepared

by Otto Krauss, a student at University of Colorado, who worked with a group of


aerospace engineers on a similar design project to ours [9]. Their project mission is

similar to our METEOR project, its title is MaCH-SR1 which aims at building a sounding

rocket to deliver a 10-lb payload to the edge of space, about 125 km above the earth’s

surface. They have chosen a single stage Nitrous Oxide / HTPB hybrid engine rocket to

achieve this goal. This paper focuses on the design and test firing of a lab-scale model of

their rocket and the results they obtained.

Many of the design components and materials they selected are coincidently

similar to many of the decisions we made, which will be explained more thoroughly in

the Concept Development portion of this paper. Going through this paper we have come

up with some good ideas to how to test various components of our design prior to

actually assembling and firing our rocket. Also, many of the references they used will be

examined more in depth.

3.2.5 Air Launched Flight Trajectories

One paper used by the team to get an understanding of some typical flight

trajectories and velocities associated with them was “Optimal Ascent Trajectory for

Efficient Air Launch into Orbit” by Frederick Boltz [10]. This paper discusses three

main types of trajectory to achieve LEO altitude and velocity; however, the two optimal

trajectories required so much travel that our mobile control unit may not be able to keep

radio contact with the rocket throughout the duration of powered flight.

The team was also referred to a computer program called STK (Satellite Tool

Kit), distributed by Analytical Graphics Inc. (AGI), a free version can be obtained by

visiting their website, www.agi.com [11]. Examination of this software package was

completed by Chris and Jessica, based on what the free version was capable of it was

determined that we needed to purchase a license for the professional version with a

missile trajectory package add on. When all was said and done we were looking at a

licensing fee over $25,000, which is far beyond the scope of our budget. Based on this

we decided to do preliminary calculations based on information we obtained from rocket

propulsion textbooks [12].


http://www.agi.com/

3.3 Hybrid Rocket 3.3.1 Introduction / Benefits of Hybrid Rocket

During the summer quarter of 2005 a fellow student at RIT, Jeffery Cappola,

conducted independent research on different methods to power a small scale rocket [13].

Based on his work and previous investigations completed by Dr. Patru [5] the obvious

choice for rocket propulsion based on safety and convenience is a hybrid rocket engine.

Other methods include pure solid or liquid propellants. Liquid propellant engines use

both liquid oxidizer and liquid fuel such as liquid oxygen and liquid hydrogen stored in

separate tanks and mixed together during the combustion process. Liquid rocket

propellants are extremely volatile and exposure to high heats can cause explosions and

dangerous situations. Solid rocket engines are sometimes unpredictable and do not have

the capability to be throttled or shut off. Choosing a hybrid engine allows us to

compensate for the negative safety issues associated with both solid and liquid rocket

engines.

3.3.2 Propellants

Our team has relied on the work done by Jeffery Cappola and several of his

references in determining which propellants to consider for our hybrid rocket; Dr. Kozak

was his faculty advisor for this work and supported its validity. Jeff came to the

conclusion the HTPB and PMMA are the two most promising solid fuels and he went as

far to suggest PMMA to possibly be a better option. These two propellants were

determined to be the most promising because of their availability, environmental

friendliness, ease of molding or manufacturing, and experimental data available from

similar projects. Our decision making process is elaborated more in the concept

development phase of the paper. For a liquid oxidizer it was obvious that Nitrous Oxide

is the clear choice because of safety and storability issues associated with more volatile

oxidizers such as liquid oxygen.

3.3.3 Performance Predictions

Thrust and mass equations are fairly basic and can be found in almost any

propulsion resource. To size a rocket, regression rate of the solid fuel must be known;


and this is no simple task. There are limited resources available in this particular area of

rocketry. Regression rate is difficult to estimate without experimental data; however,

initial estimates can be made based on equations and correlation parameters developed by

previous scientists.

The equation used to calculate thrust, seen below, was found in a book called

Mechanics and Thermodynamics of Propulsion [14].

( ) eaee Appvm −+= &τ m& is total mass flow rate of the propellant

ve is exit velocity from the nozzle pe is exit pressure

p is ambient pressure aAe is exit area of the nozzle

Tsiolvolsky’s rocket equation was found in the same book, and can be seen below:

⎟⎟⎠

⎞⎜⎜⎝

⎛=Δ

1

0lnmmvv e

Δv is the change in velocity of the rocket

ve is the exit velocity of the rocket m is the final total mass of the rocket, which includes the payload mass,

structural mass, and electrical mass component 1

m is the initial total mass of the rocket, which is the final total mass plus propellant mass

0

Knowing the change in velocity one wants to achieve, the exit velocity of the rocket, and

the total final mass of the rocket allows the propellant mass to be found.

To size the HTPB fuel grain, regression rate must be known. Experiments on

HTPB regression in a cylindrical fuel grain with a central circular port were conducted by

Chiaverini and published in a paper named Regression Rate Behavior of Hybrid Rocket

Solid Fuel [15]. Using this data, he developed a regression rate equation based on

correlation parameters, motor pressure, oxidizer mass flux, and axial location.

( ) ( )⎪⎭

⎪⎬⎫

⎪⎩

⎪⎨⎧

⎥⎦

⎤⎢⎣

⎡−

−+−= −

−−

−kph

nnkph

nnmn e

xGCe

xGCxGCr 1exp1 1

0

21

0

201

r is instantaneous regression rate of the solid fuel [mm/s]

G is the oxidizer mass flux [kg/m2-s] ox is axial location [m]


k is the gas absorption coefficient [(m-MPa)-1]

p is the pressure [MPa] h is the port height between fuel slabs [m]

n,m,k,C ,C1 2 are all parameters developed by Chiaverini whose values are given in the

table below, specifically for a mixture of 96% HTPB and 4% Ultra-Fine Activated Aluminum (UFAL)

Correlation Parameter Table

Parameter 96% HTPB / 4% UFAL Units C 0.0535 n/a 1

C 14.197 n/a 2

n 0.63 n/a m 0.122 n/a K 57.11 (m*MPa)-1

Table 1 - Experimentally Determined Correlation Parameters

In the “Regression-Rate and Heat-Transfer Correlations for Hybrid Rocket

Combustion” paper by Martin J. Chiaverini, Kenneth K. Kuo, Arie Peretz, and George

C. Harting at The Pennsylvania State University, Pennsylvania 1680 [16], another one of

the regression rate equations we used in our analysis was found. This paper showed an

analysis of a hybrid rocket motor that used solid propellant, and how the regression rate

was obtained for this engine. The regression equation that they experimentally found is

shown as:

)/exp( sua TREAr −=

The parameters used in this equation include Ea as the activation energy (kcal/mol), Ru as

the universal gas constant (kcal/kg-K), Ts as the surface temperature, and A as the

Arrhenius pre-exponential constant (mm/s).

Their test engine included two opposing fuel slabs with an oxidizer flow rate of

530 kg/m2. They used both real-time, X-ray radiography and ultrasonic pulse-echo

systems to deduce the local, instantaneous solid-fuel regression rates. Pressure

transducers provided the motor pressure history along the motor port. The regression

rates were found using video images of the pyrolysis process, with micro thermocouples

showing the surface temperature.

This group found the equation parameters using their test results by correlating

the regression rate and the surface temperature. They then plugged these parameters into


the pyrolysis law, which is the above equation, where the surface temperature was greater

than 722 K. This method was shown in this paper to be valid for HTPB hybrid engines.

3.3.4 Fuel Chamber

Some of the basics for the configuration and different components of the solid

fuel chamber were examined from multiple papers, one in particular, A Preliminary

Design Code for Hybrid Rockets, Werthman, W.L. was exceptionally useful in coming

up with a base design [17]. This paper introduced us to the basic idea of how the engine

works, along with the idea of pre and post combustion chambers. Using theses chambers

with no fuel grain allows for a more uniform flow and more predictable thrust. Teamed

up with papers of similar projects we were able to come up with the basic configuration

elaborated on in Section 6.1.3.1.

Several papers were used to design based on optimal conditions inside the

chamber [18] [19]. We discovered that, optimally, the hybrid rocket solid fuel chamber

should operate at around 550 psi, and the turbulent boundary layer that results over the

fuel grain, along with the solid fuel’s evaporation, will keep some of the high

temperatures away from the surface of the walls and fuel grain.

3.4 Feed System The basic schematic for our feed system was taken from a labscale hybrid rocket

tested at University of Arkansas at Little Rocket with similar intentions to our ground test

[20]. The main objectives of the feed system is to supply the fuel chamber with oxidizer

but also to ensure that no hot gases are able to flow back into the tanks. The schematic

we are modeling after also uses a pure nitrogen purge system to shut of the combustion

process by flushing out any oxygen that may be in the chamber.

To predict the flow rates and pressure drops through the system we referred to

Rocket Propulsion Elements: 7th Edition, Sutton [12]. This book was also used when

trying to predict the performance of our hybrid rocket engine combustion process.

Predicting the pressure drops across the different valves and components of the feed

system allowed us to ensure that we can obtain the desired pressure drop across the

injector to get a well atomized oxidizer flow into the chamber.


To locate many of the valves and components of our feed system we relied on

McMasterCarr.com for many of the basic components and GlobalSpecs.com to search for

some of the more specialized components. Brad Addona was primarily responsible for

the feed system, he has worked for a company, ValveTech, which specializes in

aerospace valves for some time and we have used that as a resource for common

information and guidance also. ValveTech was also generous enough to donate some

solenoid valves that will be used in our feed system.

Further Investigation

One important thing to note is that the feed system we have designed is intended

for ground testing only. For the feed system that will be used on the actual rocket much

lighter, smaller, more specialized, and much more expensive valves will need to be used

to cut back on redundant weight. In addition we have not been able to locate an

electronically controlled throttling valve which can change the flow rate over time, which

would be highly desirable for the actual rocket to obtain constant oxidizer to fuel ratios

throughout the combustion process.

List of Companies Used for Components:

• McMaster Carr o Needle Valves o Check Valves o Tubing o Pipe Connection Components o Mechanical Pressure Gauges

• Omega o Pressure Transducers

• CoAx o Solenoid Valve

• ValveTech o Solenoid Valve (donated)

• AeroCon o Oxidizer Tanks


3.5 Ignition System When performing an initial concept design as a team we agreed to use a propane

type of ignition system, filling the chamber with propane prior to firing a spark plug and

introducing the flow of Nitrous Oxide to the chamber. After further investigation it

became apparent to the team that because of the lack of oxygen in outer space this type of

system would not work at all. This sparked a more thorough literature search to

determine other more reliable methods to ignite our hybrid engine. Once again we

resorted to looking at previous work done [21, 22] and also looked into methods that are

used in basic solid propellant Estes model rocket engines [23].

One of the most common methods of ignition is similar to our propane idea

except that you would fill the chamber with pure oxygen rather than propane. Our team

has decided to shy away from using pure oxygen because of its potential hazards and

corrosive properties. Using pure oxygen would greatly increase the risk to the students

and faculty facilitating the ground test of our hybrid rocket.

One of the more promising methods but still requires further investigation is using

a glow plug or the like to pre-heat the Nitrous Oxide before it enters the chamber

effectively disassociation the nitrogen and oxygen for easy ignition. Another method that

is used for solid propellant motors is the use of pyrogens and Ni-Chrome wires.

Essentially you supply a current to the Ni-Chrome wire which heats up considerably

igniting NC lacquer or a similar pyrogen which has been coated around the wire. In

order to ensure that the pyrogen burns long enough to separate and burn the oxidizer we

would surround the NC lacquer with an Ammonium Perchlorate / HTPB combination.

Ammonium Perchlorate is a solid oxidizer that is able to burn in oxygen deprived

environments, its burn temperatures are high enough to disassociate the oxidizer and it

burns slow enough to ensure that there is the necessary amount of time to ignite our

rocket. Further investigations of these methods are outlined in Section 6.1.4.

3.6 Injector The injector serves multiple purposes, its main objective is to reliably deliver

oxidizer into the solid fuel chamber in order to stimulate the combustion process, but a

secondary purpose of the injector is to atomize the oxidizer flow as much as possible.


Making the particle size of the Nitrous Oxide as it enters the fuel chamber as small as

possible increases the rate of disassociation of the oxygen and nitrogen which ensures no

wasted oxidizer and more predictable thrust [24]. Based on designs that have been

experimentally proven to improve this asset of the injector we decided to look at three

different designs: a swirling nozzle [25], a shower head nozzle, and using a technique

called gasification [24].

3.7 Exit Nozzle Designing the exit nozzle is one of the most critical portions of a rocket design,

but in most respects it is a rather straight forward process. Information regarding this

process can easily be found in Fluid Mechanics and Aerodynamics textbooks [14, 26].

The basic concept is to run the hot gases through a converging subsonic nozzle until the

flow reaches mach 1 at the throat area. From there you send the fluid through a diverging

supersonic nozzle. Using the Method of Characteristics to ensure isentropic flow further

increases the effectiveness of your nozzle.

3.8 Federal Specifications In order to have the proper permission to test a hybrid rocket we need to consider

and abide by any local and federal regulations that pertain to this type of test. We have

located two different resources that we are basing this assessment on. The first is based

on FAA regulations [27] and the second according to MIL specs [28] which are

commonly used in industry. To further gain permission we have completed a safety

review report that we have provided to facilities management to gain permission to test

our rocket on campus, see Appendix 12.


4. Needs Assessment Our team being the first design group of the second phase of the METEOR

project, the rocket design phase, we needed to properly assess what should be done in

order for following teams to be successful. Our project being a very research intensive

project made it difficult to properly predict exactly what could be completed in 20 weeks,

for both the team and our mentors. Based on the original proposal and discussions we

had with our sponsors the team decided on the following goals and objectives.

4.1 Performance Goals

• Hybrid engine must meet predicted performance requirements to carry a 1 kg

satellite to lower earth’s orbit

• Upper stage must be able to safely house, transport and release a Pico Satellite

• Team must establish a safe means for ground testing the rocket

• Gain as much knowledge and experience with propellants as possible

4.2 Design Goals

• Hybrid engine must be reliable and predictable when firing

• Team must design a safe and useful test stand to collect data

• Secure a location on campus to complete ground testing

• Choose propellants and all components of a test fuel chamber

4.3 Secondary Goals

• Structural design paper on the final stage of the four stage rocket

• Choose a flight pattern and trajectory to achieve lower earth’s orbit

• Devise a means of thrust vectoring to control attitude

• Design stage separation mechanical device


5. Specifications / Success Qualifiers After conducting some research and meeting with our sponsors the team was able

to determine exactly what specifications we needed to achieve in order for our design to

be considered successful. The following is what we determined:

• The redundant structural weight (empty weight) needs to be less the 15% of the

total weight, for our paper design

• Rocket engine must be able to achieve a change in velocity of Δv = 2300 m/s

• Determine the optimum oxidizer to fuel (O:F) ratio for maximum thrust

• Obtain enough temperature and pressure data to properly design a fuel chamber

for the paper design

• Structural design of upper stage must be able to withstand 30 g’s

• Ignition system must properly work 100% of attempted fires

• Be able to measure thrust in ground test with a 1% resolution factor


6. Concept Development / Feasibility

6.1 Hybrid Engine Because the many of the design objectives behind our project are focused around

the hybrid rocket engine and its performance the team put in a considerable amount of

effort to analyze the importance of the different aspects of a hybrid engine. We chose to

use a weighted method analysis to determine which of the customer requirements carried

the most importance towards our design; the results are shown in Table 2. We used the

following customer requirements in our analysis:

Customer Requirements:

• Reach Lower Earth Orbit (LEO)

• Weight Requirements

• Satellite Safety (acceleration)

• Safety During Testing

• Cost

• Propellant Availability

Examining these results the team was able to determine that the safety during

testing was the most important requirement not only to our sponsors but also for

ourselves. Using this information we then applied it to a QFD analysis to get an

understanding of which specification or metric was the most important to ensure a

successful design. We used the following specifications to qualify a successful design.

The results from the QFD analysis can be seen in Table 3.

Hybrid Engine Specifications:

• Specific Impulse, Isp (sec)

• Mass Flow Rate (kg/sec)

• Empty Weight (kg)

• Chamber Pressure (Pa)

• Extra Solid Propellant for Insulation

• Force Produced (N)


• Total Impulse (N-s)

• Burn Time, tb (sec)

• Propellant Density

• Propellant Selection

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Design a hybrid rocket engine capable of containing and realeasing a 1kg satellite into low-earth orbit (LEO). Rocket will launch from a platform suspended at 100,000 ft above earth's surface.

2. CUSTOMER REQUIREMENTS 1 Reach Low Earth Orbit (LEO) 0.27 9 7 8 6 1 8 10 7 5 8 2 Weight Requirement 0.13 1 1 10 1 7 1 1 1 6 6 3 Satellite Safety (acceleration) 0.2 5 3 2 1 1 10 7 5 3 3 4 Safety During Testing 0.33 1 6 1 9 5 8 5 5 2 5 5 Cost 0 1 1 5 1 5 1 1 1 5 9 6 Propellant Availability 0.07 8 1 1 1 3 6 6 4 3 10

4.4 4.7 4.3 5.0 3.3 7.3 6.3 4.9 3.6 5.98. ABSOLUTE IMPORTANCE 0.6 0.6 0.6 0.7 0.4 1.0 0.9 0.7 0.5 0.89. RELATIVE IMPORTANCE 10. TARGET SPECIFICATIONS 11. RISK EVALUATION

12. ASSESSENT OF COMPETITOR'S OR EXISTING DESIGN

* 10 = ABSOLUTELY ESSENTIAL, 7 = VERY IMPORTANT, 5 = MODERATLY IMPORTANT, 3 = NOT VERY IMPORTANT, 1 = UNIMPORTANT

Table 3 - QFD Analysis of Hybrid Engine

Conclusions

Based on the QFD analysis completed on the hybrid rocket engine it became

apparent that the amount of thrust and total impulse were the most important

specifications to look at. Close behind them was the propellant selection process, which

in retrospect we realized ultimately determines what kind of thrust and impulse we are

able to create. With our team focused on propellant choices based on thrust and impulse,

and the customer requirement of safety for testing in the forefront of our minds we went

forward on the hybrid engine design.


6.1.1 Propellant Selection 6.1.1.1 Solid Fuel Selection

Based on our literature search the team was looking two different types of safe to

handle, readily available, and environmentally friendly solid propellants. One is

Hydroxyl Terminated Poly-Butadiene (HTPB), similar to tire rubber, and another being

Poly-Methyl Methacrylate (PMMA), which is essentially an acrylic Plexiglas material.

Both of these materials had negatives and benefits in different areas so we decided to use

a pugh analysis to determine which of the two to go forward with. Table 4 illustrates the

pugh analysis we conducted, based on this analysis we determined that HTPB is a better

option because of its cheap cost and its ability to be cured into any desired mold we

construct. It should be noted from Table 4 that the normalized scores are very close and

that future teams may want to look into testing PMMA verse HTPB experimentally. The

main advantage of PMMA is its higher density which would allow for a smaller fuel

chamber which in turn would lower the redundant weight of the rocket.

6.1.1.2 Liquid Oxidizer Selection

The biggest determining factor when selecting a liquid oxidizer is the safety and

storability of the material. Possible oxidizers include liquid Oxygen (O2), Oxygen

Difluoride (F O), Hydrogen Peroxide (H2 2O ), Nitrous Oxide (N2 2O) plus many, many

more [13]. When safety and environmental factors were incorporated the obvious choice

became Nitrous Oxide. Nitrous Oxide is considered a non-combustible (below 300oC),

non-hazardous material see Appendix 13 for the MSDS and Hazmat information for

Nitrous Oxide. The Nitrous Oxide and HTPB combination is non-combustible unless

temperatures above 300oC are introduced, at which point the nitrogen and oxygen

disassociate allowing for the oxygen to burn with the HTPB [29]. Essentially this means

that we could run Nitrous Oxide through our fuel chamber and would not have any

combustion unless we turn on our ignition system.



Evaluate each additional concept against the baseline, score each attribute as: 1 = much worse than baseline concept 2 = worse than baseline 3 = same as baseline 4 =

better than baseline 5= much better than baseline

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6.1.2 Fuel Grain / Propellant Sizing 6.1.2.1 Propellant Mass

In order for our four stage rocket to reach lower earth orbit, calculations needed to

be done regarding how much solid propellant was needed, and the velocity needed to

reach the proper altitude. The propellant diameter, thickness, and length were found

using the equations shown in Appendix 7. To begin these calculations an exit velocity

had to be found, along with a total velocity using Tsiolkovsky’s basic rocket equation

[14]:

⎟⎟⎠

⎞⎜⎜⎝

⎛=Δ

1

0lnmm

vv e

mo is the initial total mass of the rocket m is the final total mass of the rocket 1

ve is the rocket’s exhaust velocity ∆v is the rocket’s change in velocity

This essentially tells us the maximum rocket velocity change that can be achieved by

expelling a known amount of mass (m -m ) at a known velocity. o 1

The mass of the propellant was the first dimension of the propellant we found. By

using the specific heat and gravity we found the exit velocity of stage four of our rocket

to be 2,300 m/s. This velocity times each of our four stages gives a total velocity that

needs to be achieved of 7,600 m/s, which is adjusted for gravitational and drag effects,

resulting in a final total burnout velocity of 9,200 m/s. This number used in conjunction

with our four stages resulted in the rocket’s change in velocity to be 2,300 m/s per stage.

After finding the total velocity, the masses were calculated using three different

parts of the rocket to make up the total mass. These three parts included the payload

mass of 1 kg satellite, a structural mass of 0.75 kg (estimated using a redundant weight of

15%), and an electrical component mass of 0.3 kg. When theses masses were added up

they gave a total mass of 2.05 kg. By using the above rocket equation, plugging in the

exit velocity, velocity change, and the total mass, an initial mass of the rocket was found

to be 5.572 kg. By subtracting the initial mass of the rocket from the total mass, a

propellant mass was found to be 3.522 kg.


6.1.2.2 Regression Rate

The next step was to use regression rate analysis that will later be used to find the

diameter of the solid propellant. The regression rate equation (as shown below) that was

found from extensive research in a paper by Chiaverini [16], was used to find an estimate

of the regression rate for the last stage of our four stage rocket:

( ) ( )⎪⎭

⎪⎬⎫

⎪⎩

⎪⎨⎧

⎥⎦

⎤⎢⎣

⎡−

−+−= −

−−

−kph

nnkph

nnmn e

xGC

exG

CxGCr 1exp1 1

0

21

0

201

r is the regression rate of the solid fuel [mm/s]

G is the oxidizer mass flux [kg/m2-s] ox is axial location [m]

k is the gas absorption coefficient [(m-MPa)-1]

p is the pressure [MPa] h is the port height between fuel slabs [m]

n,m,k,C ,C1 2 are all parameters developed by Chiaverini whose values are given in the

table below, specifically for a mixture of 96% HTPB and 4% Ultra-Fine Activated Aluminum (UFAL)

Correlation Parameter Table

Parameter 96% HTPB / 4% UFAL Units C 0.0535 n/a 1

C 14.197 n/a 2

n 0.63 n/a m 0.122 n/a k 57.11 (m*MPa)-1

Table 5 - Correlation Parameter Table - Regression Rates This equation was placed into a spreadsheet and analyzed, see Appendix 8. However,

there was some difficulty and error in using this particular regression rate equation.

Further research was done; and a paper by the same individual, Chiaverini, showed

another, simpler regression rate equation:

⎟⎟⎠

⎞⎜⎜⎝

⎛⋅

−⋅=

s

a

TRE

Ar exp

is activation energy and is given as 20.557 kJ/mol Ea

A is the Arrhenius pre-exponential constant, given as 11.04 mm/s


T is the surface temperature of the fuel grain [K] s

R is the universal gas constant which is equal to 8.3143 J/(mol-K) r is the solid propellant regression rate [mm/s]

This equation was also put into a spreadsheet, Appendix 8, and resulted in a regression

rate of 0.9315 mm/s. This number was a closer match to the regression rates that had

been found in research papers. The exact regression rate can not be calculated until

testing of the solid propellants has been completed. Due to the non-exact nature of this

calculation, the regression rate we decided to use is 1 mm/s. This adjustment should put

us in a range of having slightly extra fuel should our theoretical regression rate be slightly

off due to human calculation error.

6.1.2.3 Mass Flow Rate / Fuel Grain Sizing

After the regression rate was found, the next step was finding the mass flow rate.

First the assumption of L/D = 10 and a 0.03 m inner diameter was made, which is

accurate with other theoretic calculations we ran across. An outer diameter was

calculated using the regression rate, a burn time of one second, and the inner diameter.

The outer diameter after one second was found to be 0.032 m, which was used in

conjunction with the inner diameter and length to find the volume of HTPB. The volume

of HTPB was then multiplied by the density of the HTPB to give the mass flow rate after

one second for our solid propellant of 0.0272 kg/s.

Using an oxidizer-to-fuel ratio of 8:1, the mass for HTPB was calculated. By

taking 1/9 (due to the ratio) of the fuel mass that was calculated earlier, the HTPB mass

came out to be 0.3913 kg. By placing the mass of HTPB over the mass flow rate after

one second of HTPB, the burn time was calculated to be 14.4 seconds.

Finally, a back calculation was done to find an outer diameter for 14.4 seconds.

This was done by adding to the inner diameter the burn time multiplied by the regression

rate multiplied by two for each side of the solid propellant. The resulting outer diameter

came out to be 0.0635 m. We increased this outer diameter so that we could have extra

fuel to insulate the combustion chamber walls due to the heat that the combustion

chamber will produce. Also extra fuel was used to allow standard pipe sizes for the

combustion chamber to be used in our design.


An oxidizer flow rate was calculated as shown in the Appendix 4 using the

remaining 8/9 ratio of fuel over the burn time, resulting in 0.2174 kg/s of oxidizer.

Results of solid propellant calculations: Inner Diameter = 0.03 m Outer Diameter = 0.0635 m

Length = 0.3 m Weight = 3.522 kg

6.1.3 Fuel Chamber Intended for Ground Testing 6.1.3.1 Basic Configuration

Figure 6 shows a cross-sectional view of the test chamber and identifies all the

major components. After the oxidizer enters the chamber through the injector it enters a

pre-combustion chamber. The purpose of the pre-combustion chamber is to ensure the

disassociation of the nitrogen and oxygen and to further atomize the particle size of the

nitrous oxide before it reaches the fuel grain.

As seen in Figure 6 we have decided to go with a simple cylindrical port fuel

grain. The amount of thrust a hybrid rocket is able to produce is dependent on the surface

Test Chamber Wall Exit Nozzle

Injector Fuel Grain

Garolite Laminated Ceramic

Snap Ring

Fig. 6 – Cross Sectional View of Fuel Chamber


area that the fuel grain and oxidizer are interacting. With this in mind making a star

pattern or a multi-port fuel grain would help to improve thrust. But because of the

simplicity of the mold and potential danger of fuel grain separation prior to burning we

have chosen a cylindrical pattern for the fuel grain.

The post combustion chamber is intended to help make the flow more uniform as

it enters the nozzle. This will in turn generate a more predictable and steady exit plume

which translates into a more predictable and steady thrust produced by our engine.

Because of the extreme heat we are expecting from the combustion process a Garolite

Laminated Ceramic will be inserted to line the walls of the pre and post-combustion

chambers. Extra solid fuel will be used to insulate the walls over the length of the fuel

grain.

The last component to be inserted into the fuel chamber is our graphite exit

nozzle. A Snap Ring will be inserted to hold all of these components into place at the

back end of the exit nozzle. The injector plate is held on to the front portion of our

chamber by eight, SAE Grad 2, ¼-20 bolts.

6.1.3.2 Obtaining Required Data

The fuel chamber must not only safely house all of the components that we need

inside to fire a hybrid engine, but it must also allow us to collect the necessary data we

need to predict what is occurring inside the chamber during firing. The team needs to

collect temperature and pressure readings from inside the chamber in order to properly

predict what materials and dimensions they need for a fuel chamber that would be used in

the actual METEOR rocket. Because of redundant weight issues this chamber would

need to have a low factor of safety and be able to withstand the volatile environment that

can be expected from the combustion process. All of this means that the team needs to

collect accurate, time dependant data throughout the firing of the rocket.

Pressure Measurements

The team initially went on the hunt for thermocouples and pressure transducers

that could possible withstand such high temperatures. This search concluded after

quickly realizing that within our budget and time constraints that this was not possible.


After talking to a few different people in industry we were suggested the idea of a

listening tube in order to collect pressure data. The idea of the tube is to effectively allow

the gases from the chamber cool before coming in contact with the pressure transducer.

The idea is simple, attach extended tubes to the chamber through a port hole in the

side of the chamber, and collect the static pressure safely. However there are many

complications associated with this process which the team needs to reconcile prior to

using this method to collect pressure data. The listening tube that would be directly

tapped into the flow would need to be able to withstand the worst case pressure scenarios.

In addition at the interface of the tube and the chamber it would need to be able to

withstand the extreme temperatures associated with the combustion process.

Furthermore in order for the tube to be directly tapped into the flow we would need to

drill a hole through the ceramic liner we are using to protect our chamber. All of these

problems need to be cured prior to the team being able to collect pressure data.

One other complication with using listening tubes, but is easily taken care of is

the time delay seen by the transducer from the real time pressure in the chamber. A

simple test can be conducted to back out this delay and use that to correlate our collected

data. By simply assembling the tube, not connected to the chamber, with the pressure

transducer, one would need to fill the chamber with some set pressure. After reaching a

steady pressure in the tube start your data acquisition and at some recorded time release

the pressure from the tube. The difference from the time you release the pressure to the

time that that is reflected in your data is you time delay associated with your listening

tube.

Temperature Measurements

In order to take temperature measurements in the chamber the team has decided to

embed the thermocouples into the exterior of the chamber and use heat transfer equations

to back out the information. Before this can be successfully done tests will need to be

conducted on the chamber, ceramic liners, and solid fuel in order to obtain their heat

transfer coefficients. This method will allow us to not only back out what is going on

inside the chamber, but also estimate the temperatures that would be seen outside the

ceramic liners and fuel grain where the actual fuel chamber interface is located.


6.1.4 Injector Design 6.1.4.1 Basic Concept

The injector serves multiple purposes, it main objective is to reliably deliver

oxidizer into the solid fuel chamber in order to stimulate the combustion process, but a

secondary purpose of the injector is to atomize the oxidizer flow as much as possible.

Making the particle size of the Nitrous Oxide as it enters the fuel chamber as small as

possible increases the rate of disassociation of the oxygen and nitrogen which ensures no

wasted oxidizer and more predictable thrust [24]. Based on designs that have been

experimentally proven to improve this asset of the nozzle we decided to look at three

different designs: a swirling nozzle [25], a shower head nozzle, and using a technique

called gasification [24].

All three of these methods serve the same purpose is decreasing the particle size

of the oxidizer. The swirling nozzle sends the flow into a vortex motion prior to reaching

the combustion point of the chamber

increasing the mixing and decreasing the

particle size, see Figure 7 for a schematic of a

basic swirling nozzle design. It can be seen

from these diagrams that actually machining

this part is no easy process and would require

us to outsource this part to a specialized

machine shop. This would greatly increase

the cost for our injector.

The method of gasification is a

relatively new science that is just beginning t

be examined in the scientific realm. Figure 8

illustrate the effects of gasification and its

benefits. The basic idea of gasification is to

introduce an inert gas into the flow of

oxidizer as it enters the injector which will greatly decrease the particle size as can be

seen in Figure 8. The down side of this method is that it introduces more redundant

Fig. 7 – Swirling Nozzle

o


structures to the system which would increase the paper design weight of the actual

rocket.

Fig. 8 – Spray Pattern for Various Gasification Injection Gas Ratios

The third method that we considered was a showerhead type of design. We

figured this was a simple, cheaper method of trying to accomplish the same goals of the

previous two types of injectors. We completed a pugh analysis on these designs, Table 6,

to weigh the different factors of each design. From these results the team decided to go

forward with the showerhead design.


Pugh Analysis: Injector Design Showerhead Swirling Nozzle Gasification

Manufacturability 3 1 2 System Complexity 3 2 1

Cost 3 1 2 Effectiveness 3 4 5

Weight 3 2.5 1 Total points: 15 10.5 11

Table 6 - Pugh Analysis of Oxidizer Injection Methods 6.1.4.2 Machining / Assembly

In order to machine our injector we will use the same method as the University of

Colorado used to machine their similar injector, Electron Discharge Machining (EDM).

This method of machining allows us to make holes through the injector on the order of

.025” or less. The smaller the size of the holes and the greater the number, allows us to

increase the atomization of the oxidizer which is our design intention.

Fig. 9 – Initial Injector Concept Drawing


Initially we designed the injector to be a two piece part that needed to be bolted

together and then from there attached to the thrust plate. To cut down on parts and

potential sealing issues we redesigned the injector to be one single piece of material that

acts as an injector and thrust plate combined; see Appendix 1 for the final part drawing.

Figure 9 and Figure 10 highlight the differences between the hand sketch of our original

design compared to our final part drawing.

Fig 10 – Final Injector Concept Cross-Section

6.1.5 Ignition System There were four basic types of ignition systems that we considered for use

in our Hybrid Rocket; a pure oxygen, a propane gas, a glow plug / N20, and a

pyrotechnic system. In this section, we will discuss the pros and cons of the different

types of ignitions. A Pugh analysis was completed on the four types of ignition systems

and those results can be seen in Table 7.

Our first ignition system concept used pure oxygen and an electronic spark to

provide a 300°C environment to disassociate the Nitrous Oxide. The parts list consisted

of an oxygen tank, tubing, electronic on/off valve, a nozzle, and an electronic spark plug.

The oxygen tank would be contained on board and have enough oxygen to burn for

approximately 3 seconds. Oxygen would be run from the tank through the piping,

electronic on/off valve, through the injector, and into the pre-combustion chamber. An

electronic spark would then ignite the oxygen which would heat the chamber up to a

minimum temperature of 300°C.


Ignition Systems Pugh Analysis

Glow Plug / Nox Pyrotechnic O2 Propane

System Complexity 3 1 2 3 Cost 3 3 1 2 Safety 3 4 1 2 Effectiveness 3 1 5 0 Repeatability 3 4 5 0 Weight 3 2 2 2 Manufacturability 3 1 4 4 Possible Damage to Other Components 3 5 2 5 Total: 24 21 22 18

Table 7 - Pugh Analysis of Ignition Systems Various ideas were tossed around as far as creating the spark is concerned. The

spark could simply be created from two wires a short distance apart in the pre-

combustion chamber. A voltage applied across these two wires would create a spark and

consequently ignite the oxygen. Most likely the voltage would be applied on and off for

1-3 seconds to ensure a good oxygen ignition. A schematic of the oxygen system is

shown in Figure 11.

Fig. 11 – Pure Oxygen Ignition System Schematic

+ -

Pre-Combust

ion Chamber

O2

NOx

On Off

O2 Nozzle

HTPBHTPB


One of the main objectives of this hybrid rocket was to make all of its fuels as

safe as possible for the user and environment. Our team decided against the pure oxygen

ignition system for several reasons. The first of those is that oxygen is extremely

flammable and there would be a high risk to any students and faculty in the vicinity of the

test. Concerns with rocket approval by our mentors and sponsors were also taken into

consideration.

The second ignition system considered by the team examined the use of propane

gas instead of oxygen, to provide a 300°C environment with the same schematic as

Figure 11, just a propane tank in place of the oxygen. Propane is much easier to control

and is less flammable than oxygen. The risks to students and faculty would be greatly

decreased with the use of propane. Upon further investigation, it became apparent that

propane would not work, given the low amount of oxygen at altitudes above 100,000 feet

and propane sill requires oxygen to burn. The conclusion to this concept was that the

propane ignition system would not have enough oxygen at 100,000 feet to burn and ignite

the engine.

Due to the fact that the oxygen and the propane ignition systems would not work

it was back to the drawing board to come up with another system. Upon a literature

review of other rockets ignition systems, it was found that two more concepts remained

for investigation; a N2O / glow plug ignition and a pyrotechnic ignition.

The glow plug ignition system uses an offshoot of N2O to ignite the main feed of

N2O. A small amount of N2O would be tapped off the main feed system, this would run

through an on/off solenoid and then into a small chamber. In this chamber there will be a

glow plug, such as the one pictured in Figure 12, which would disassociate the N2O.

The glow plug upon disassociation of the N2O would then ignite the oxygen that

is separated from the nitrogen which would then heat the main feed of N2O. A schematic

of this concept is shown in Figure 13.

Fig. 12 – Glow Plug


Main N2O Feed Separation Plate Combustion Chamber

Fig. 13 – Glow Plug Ignition System Schematic

N2O Solenoid Glow Plug

There are several points of concern that will require some experimentation to

ensure that this ignition system would work. An optimal flow rate of the N2O passing

over the glow plug would have to be established, if the flow rate is too high, the glow

plug may not be hot enough to ensure disassociation of the N2O. Experimentation may

also be done with trading out the glow plug for a segment of Ni-Chrome wire across the

flow. Another concern is if the ignited main feed flame can reach the HTPB in order for

the HTPB to ignite. This system is relatively inexpensive but must be investigated prior

to being put to use in the rocket engine. The team decided that this system would be

more complicated than a pyrotechnic system which is the next subject of our

investigation.

The fourth ignition system concept analyzed by our team uses a small amount of

controlled pyrotechnics to provide a flame and environment hot enough to ignite the N2O

and HTPB. A schematic of a typical pyrotechnic igniter is shown in Figure 14.


Ammonium Perchlorate Ring * Pyrogen: BP/ NC Laquer NiChrome Wire Epoxy Aluminum Wires

* The Ammonium Perchlorate ring was added to increase the burn time.

Fig. 14 – Pyrotechnic / Ni-Chrome Igniter

The basic concept behind the pyrotechnic igniter or electric match is a current

runs through the Ni-Chrome wire heating up the wires to a temperature hot enough to

ignite the BP / NC lacquer. Due to the BP / NC lacquer being a very quick “explosion”

our team decide to add a ring of slower burning Ammonium Perchlorate to the igniter.

The main feed of Nitrous Oxide should be turned on while the ignition of the Ammonium

Perchlorate ring takes place. Upon exposure to the burning Ammonium Perchlorate, the

nitrogen will disassociate from the oxygen and consequently ignite the HTPB. Now that

a brief explanation of the pyrotechnic ignition system has been completed, information

on the individual components is helpful; Ni-Chrome wire, BP / NC Lacquer, and

Ammonium Perchlorate.

Nickel - Chromium Alloy (Ni-Chrome) wire is a very thin high resistance wire.

Ni-Chrome provides an intense amount of heat when a low voltage is applied to it. This

is due to the high resistance of the material. The resistance also depends on the length of

the wire. Ni-Chrome wire will be used to ignite the BP/NC lacquer mix which will then

ignite the Ammonium Perchlorate, and consequently ignite the N2O / HTPB. A table of

approximate amperes for corresponding temperatures and wire diameters for the Ni-

Chrome wire is shown in Appendix 11.


A Black Powder (BP) / Nitro Cellulose Lacquer (C H N O6 7 3 11) combination is

typically applied over a short segment of the Ni-Chrome wire. NC lacquer can easily be

made from dissolving ping pong balls in acetone [30]. The correct combination of ping

pong balls / acetone would require some experimentation. The purpose of combining the

two is to soften the ping pong balls into pliable Nitro Cellulose plastic, black powder is

then mixed into the softened plastic. Again, the correct combination will require some

experimentation. The BP / NC lacquer mixture can then be applied to the end of the Ni-

Chrome wire by simply dipping the Ni-Chrome into the BP / NC lacquer mixture and

then allowing the acetone to evaporate. NC lacquer can also be purchased from most

local chemical suppliers.

Ammonium Perchlorate (NH ClO4 4) is a solid oxidizer that is commonly used in

rocketry. The space shuttle’s solid rocket boosters are comprised of approximately 70%

Ammonium Perchlorate. The advantages to using this oxidizer are that it burns at a lower

temperature than N2O / HTPB, and it is part oxygen, so it will burn in a vacuum

environment. Ammonium Perchlorate is relatively inexpensive and comes in the form of

a white powder. Ammonium Perchlorate is made up of 11.91% Nitrogen, 3.4%

Hydrogen, 30.22% Chlorine, and 54.47% Oxygen (Cary Academy, 2/12/2006).

Ammonium Perchlorate is also a safe material to handle; it is even occasionally used as a

food additive.

The final design for this rocket consists of Ni-Chrome wires coated with BP / NC

lacquer embedded in an Ammonium Perchlorate ring. The ring will be inserted in the

pre-combustion chamber (Figure 15). Aluminum wires will run out of the exit nozzle

and attach to a battery. These aluminum wires should burn up upon ignition. Some

experimentation will have to be completed to achieve the desired burn time based on the

amount of Ammonium Perchlorate inserted in the pre-combustion chamber. It will also

have to be determined whether or not the BP/NC lacquer can be left out of this ring. It

may take a small combination of both to achieve a good consistent ignition. Precautions

will have to be taken to ensure that there no shrapnel from the aluminum wires will

damage the exit nozzle.

A small amount of Ammonium Perchlorate may also be mixed in with the HTPB.

This will help accelerate the ignition of the HTPB. Experimentation will be done with


different amounts of Ammonium Perchlorate as well as at different locations. In one test

it will be engrained in just the front section of the HTPB; alternatively, it could be

engrained along the entire chamber.

Rocket Body

Nichrome Wire

Ammonium Perchlorate

HTPB HTPB

Ni-Chrome Wire / Ammonium Perchlorate Ring Top View Side View

~3 cm ~ 6 cm

Fig. 15 – Pyrotechnic Ignition System Schematic

Aluminum Wires to Battery

6.1.6 Feed System 6.1.6.1 Basics

The purpose of our feed system is not only to provide a controllable flow of

Nitrous Oxide into the fuel chamber but also to provide a means of being able to shut off

the combustion process in case of an emergency or at the at completion of a test. The

following describes the basic filling procedures for the tanks and the purposes of each

component that is identified in the feed system schematic in Appendix 10.

6.1.6.2 Oxidizer tank filling

The first step in the testing of the hybrid rocket engine will be to load the oxidizer

tanks with approximately 7 pounds of Nitrous Oxide. To accomplish this, we will use the

setup as shown below in Figure 16.


Fig. 16 – Oxidizer / Nitrogen Tank Fill Schematic

A pressurized supply of Nitrous will be connected to the fill connection of the oxidizer

tank assembly (With the manual shut off valve in the OPEN position). The tanks will

then be filled to the appropriate level and the manual shut off valve will then be shifted

CLOSED. After shutting off the fill line, the connection to the pressurized Nitrous can

then be removed. We will use three separate tank assemblies to allow for multiple test

firing before having to travel to a Nitrous supplier for refilling.

6.1.6.3 Nitrogen Tank Loading:

The Nitrogen is a critical part of ground testing as it provides a means for the

immediate shut down of the hybrid rocket for either a test condition or in the event of an

emergency. To fill the Nitrogen tank, first check that the tank to system valve is

CLOSED and the fill valve is OPEN. Once these are checked, connect the fill connector

to the Nitrogen supply and while monitoring the pressure transducer of the tank, fill the

tank with 900 +75, -0 psig. After the desired pressure has been reached, switch the fill

valve to CLOSED and then remove the Nitrogen supply connection.


6.1.6.4 Oxidizer Properties

Nitrous Oxide is the oxidizer of choice when considering different fluids due to

its unique property of being self pressurizing at room temperature. To get optimum

combustion, the desired combustion chamber pressure is 550 psig. It turns out that

Nitrous Oxide will pressurize itself to approximately 800 psig at room temperature. The

pressure and density varies considerably with temperature as is shown in Table 8.

Temperature

(°C)

Vapor Pressure

(Bar Abs.)

Liquid Density

(kg/m

Vapor Density

(kg/m3 3) )

0 31.27 907.4 84.86

5 35.47 881.6 98.41

10 40.07 853.5 114.5

15 45.10 822.2 133.9

20 50.60 786.6 158.1

25 56.60 743.9 190.0

Table 8 - Pressure of Nitrous Oxide Based on Temperature [29]

6.1.6.5 Valves

Many considerations must be considered for valve selection in order to achieve

optimum performance from the hybrid engine. Major valves that are needed in any feed

system are solenoid valves, relief valves, check valves and needle valves.

Solenoid Valves

A solenoid is a device in that when electrical current is applied, a magnetic field is

created that provides an axial force. In the case of a valve, this magnetic force when used

with a magnetic metal such as 430 stainless steel will allow for opening and closing of a

flow orifice. Sizing of this orifice is critical to the overall system performance. Other

design considerations are operating pressure, material compatibility and electrical power

consumption.

All liquid flow is a function of pressure drop across the orifice, orifice coefficient

and orifice diameter as shown in the below formula.


PAoCm dΔ=

•

**2** ρ

Where:

C = Orifice Coefficient d

Ao = Orifice Area

ρ = Fluid Density

ΔP = Pressure Drop Across Orifice

Orifice coefficients can be looked up in any fluids book and it would be shown

that for a sharp edged orifice, this coefficient value (Cd) can be approximated as 0.65.

Now knowing this value, you can adjust either pressure drop or orifice diameter based on

a know flow rate to size the valve orifice. A major design choice that must be closely

looked at is the design pressure drop through the valve. To provide for the optimum

system performance, the designer must limit this pressure drop to within reason. The

entire system must be considered for this value, and in our case we want to use this feed

system for many engine sizes / flow rates. If we were to use our entire allowable pressure

drop through the valve, we would not be able to achieve any higher flow rates that the

current configuration. As described later, by setting an injector inlet pressure, we want to

be able to adjust the pressure drop through the system to thereby change the flow rate for

different tests.

Materials used in the solenoid valve and with any other fluid component as seen

later must be resistant to corrosion from exposure to various atmospheric conditions. The

solenoid valve we have selected is made completely from stainless steel and nickel plated

steel, therefore meeting the corrosion resistance requirement.

The amount of electrical power consumed, while not an issue for ground testing

has great importance when considering space flight. The size and thereby the weight of a

spacecrafts batteries depends mainly of the power requirements of the craft. One then

needs to limit the power needed to operate the valve.


Relief Valves

The main purpose of any relief valve is to prevent pressure build up in a system

that could lead to failure. This therefore makes relief valve a key part of any safety

considerations. There are three considerations when sizing a relief valve.

The first consideration that must be determined is what pressure to set the valve to

open at. To do this one must consider both the operating pressure and the weakest

component in the system. For example, this rocket is designed for a 1000 psig maximum

operating pressure and the weakest components can handle up to 1200 psig. Therefore,

the relief valves must be set to less than 1200 psig and more or equal to 1000 psig where

the 1000 psig pressure will be seen.

The second consideration is what flow rate the valve can handle. If a certain

system operates at 50 scfm, the relief valve must be able to vent this flow rate or the

system will continue to over pressurize and fail. As shown in Tables 9 and 10, all relief

valves in the rocket system are capable of meeting the flow rate in that part of the system.

A third consideration is material selection. Due to the criticality of the operation

of these valves and their exposure to outdoor conditions, no corrosive material should be

used in any part of the valve. Corrosion of a part could lead to binding of moving parts

and thus no opening.

Check Valves

The primary purpose of a check valve is to permit flow in only one direction.

Check valves are used in this rocket system to isolate the Nitrogen and Nitrous Oxide and

prevent them from mixing. The design considerations for check valves are materials and

flow / pressure.

Material selection as mentioned above is critical for any fluid system component.

The use of any corrosive material such as carbon steel is prohibited from use in any of the

check valves.

The only other consideration when sizing a check valve is the maximum operating

pressure and flow able to pass through the valve. As shown in the solenoid valve flow

explanation, you want to minimize pressure drop through any component except where

necessary.


Needle Valves

The purpose of the needle valve in our system is to allow for manual manipulation

of the flow rate by a change in pressure drop across the system. As mentioned in the

injector section later, we are designing the feed system to provide a fixed pressure at the

inlet to the injector. Based on a desired flow rate, we would adjust the needle valve,

widening or narrowing the orifice, to achieve this flow. As mentioned above, flow is

dependant on pressure drop and orifice size so by changing both of these, one can reduce

or increase the system flow rate. In an actual spacecraft, in place of this needle valve is

what is called a flow orifice. This orifice is the smallest in the entire system thereby

having the greatest influence of the total flow rate. We will not go this route due to our

desire to test actual pressure drop lost in the system and test a variety of flow rates.

Temperature: 15 C

NOX Density: 822.2 kg/m3

Orifice Coefficient: 0.65 Pressure Drop: 137895.14 Pascals Flow Rate: 0.21 kg/sec * Assumes 20 psid pressure drop across valve

Equivalent Area: 2.1455E-05 m2

ESEOD (Metric): 0.005227 M ESEOD (English): 0.2058 In

Table 9 - Control Valve Sizing Calculation

Injector

The injector is the final component in the rocket feed system. The purpose of the

injector is to atomize the Nitrous Oxide from a liquid form to a gaseous form. To

complete this gasification, there must be a large pressure drop through the injector flow

holes. The size and number of these holes is directly dependant on the flow rates desired

for testing. For our initial design, a 0.21 kg/sec flow rate was considered. Using this

flow rate, we designed for a 95 psig pressure drop across the injector and a 5-hole pattern

to help distribute the oxidizer evenly throughout the chamber.


Number of Holes: 5 Temperature: 15 C

NOX Density: 822.2 kg/m3

Orifice Coefficient: 0.65 Pressure Drop: 655001.915 Pa Flow Rate: 0.21 kg/sec * Assumes 95 psid pressure drop across injector

9.84421E-06 Equivalent Area: m2

ESEOD (Metric): 0.000708 M ESEOD (English): 0.0279 In

Table 10 - Injector Orifice Sizing Calculation 6.1.7 Exit Nozzle 6.1.7.1 Nozzle Shaping

The nozzle is unquestionably one of the most important parts of any rocket motor.

It is responsible for converting the energy created by combustion into kinetic energy

which can propel the rocket. This is done by adjoining two separate nozzles to create

very high speed flows from pressurized gasses which are relatively motionless. The first

nozzle is a subsonic, converging nozzle. In this nozzle, the area gradually decreases to

speed up the flow until it reaches sonic speed. Once the flow has reached Mach 1, the

dynamics of the gas change greatly. The supersonic nozzle must be of a diverging

profile, which will expand the gasses and increase their velocity even further.

The profile of the converging nozzle is not particularly important. The main goal

when designing the subsonic portion is to minimize frictional losses as the gas passes

through it. This is done using a bell shaped nozzle [31].

The gas then enters the nozzle throat. This is the portion of the converging-

diverging (C-D) nozzle in which the flow must necessarily be at Mach 1. This is also the

portion of the nozzle with the smallest cross-sectional area. Because these flow

conditions must be reached at this location, the throat is the part of the nozzle that

determines its mass flow rate. We have decided to slightly elongate the throat to

minimize erosion of the material. The only penalty of modification is a slight increase in

friction. The benefits of maintaining optimal nozzle geometry far outweigh the frictional

losses through this minute straight section.


The design of the supersonic portion of the nozzle is very important for

maintaining an efficient flow. The profile of the nozzle must be designed for isentropic,

shock-free operation. To do this we employed a technique known as the Method of

Characteristics [26]. For the initial design, we utilized a flow software program designed

at MIT known as TODOR. This program uses the Method of Characteristics to optimize

the profile of the nozzle wall so that it follows the natural flow of the expanding gasses.

The pressure of the gasses can be determined from the cross-sectional area of the nozzle

at any point. The nozzle should then be terminated when the pressure inside of it is

approximately equal to the pressure outside of the rocket motor.

One of the main benefits of the METEOR rocket project is that powered flight

always occurs in a virtual vacuum. Because of this we are able to design a nozzle for

isentropic operation throughout the duration of the rocket’s flight, without the need to

change nozzle geometry, as is done on many ground-launched rockets.

6.1.7.2 Nozzle Materials

The combination of high exhaust temperatures and high velocity flow through the

nozzle make material selection very important. The gasses will be entering the nozzle at

around 3300 K, cooling to 3000 K as they pass through the throat. There are no light

metals that will stand up to this heat so we were forced to consider other nozzle materials,

based more on weight and melting point than strength properties. After much research

into material properties, we decided to use graphite as a nozzle material. It possesses

relatively low weight and very high melting temperatures, which can exceed 4000 K.

Even with such a robust nozzle material, we still expect to have a need to replace nozzles

after only a few test firings.

6.1.7.3 Nozzle Attachment

Ideally, the nozzle for the test motor would be attached to the combustion

chamber using shear pins. This would allow it to blow out if it was somehow blocked

and chamber pressure was to build too high. Unfortunately, due to the properties of the

graphite nozzle material, a pin type attachment would not be sufficiently strong. We


were therefore forced to utilize another lightweight attachment scheme which

incorporates a stainless steel snap ring aft of the nozzle.

The combustion chamber spacers, fuel grain, and nozzle will all be inserted

through the rear of the combustion chamber. The addition of a thin stainless steel washer

between the nozzle and snap ring may prove necessary. Weight considerations would

suggest that a washer not be used, but it would add considerable strength to our nozzle

end if it were in place. For testing, very high strength is not a desirable trait because of

the safety aspect of a nozzle blowout, as compared to a chamber rupture if failure were to

occur. Because of our design which “sandwiches” the chamber spacers, fuel grain, and

nozzle inside the chamber, addition of the washer would merely require shortening of the

ceramic spacers, and therefore would not require any redesign.

6.2 Test Stand 6.2.1 Introduction

To complete a main objective of the senior design project the team must statically

test a rocket motor that will be similar to the final stage of the production rocket. There

is no location on campus that is set up to do this so a test stand must be designed and

validated for the static tests. During the design process team members must keep in mind

the safety of observers and the environment. An acceptable design must be produced,

approved and built by the early stages of Senior Design II to be able to complete the

required testing objective.

6.2.2 Basic Configuration

The design team began with a design that would be able to be housed in a test cell

in the Mechanical Engineering machine shop (Figure 17). This concept calls for

something similar to a sled for the motor to be strapped to. When the motor fires the sled

will move forward and put pressure on the load cell. This will then give the observers in

the room next to the test cell a reading of thrust. The exhaust created by the motor will

be removed from the room by a large exhaust fan.


Figure 17 – Concept for testing within test cell.

Attention did not need to be given to the possibility of the combustion chamber

exploding because of the design of the room. Through speaking with Dr. Wellin it was

learned that load cells work better in tension than in compression, this adjustment will be

made on future design concepts.

Also, after speaking with Dave Hathaway it was learned that the desired test cell

was not available due to the cost of turning it from its current state back into a useable

test cell. Mr. Hathaway suggested that the test be set up right outside of the machine

shop. The motor would be oriented vertically, pointing towards the ground. Two pieces

of lexan, currently owned by RIT, and the brick wall of the machine shop would be used

to create a triangle for the test. There was a concern about having the exhaust of the

rocket fire into the air. Any debris that may be ejected from the motor would be

uncontrolled and could fall on a team member or an innocent bystander, possible causing


injury. It was from this discussion that the team is now pursuing a horizontally mounted

test stand.

From here the team turned its focus to using the driveway on the West side of the

machine shop. The motor would face North, would be contained by 2 lexan panels and a

debris blanket on the top to slow down any shrapnel from a possible explosion. This

setup might prove slightly harder to pass by campus safety, however with enough safety

equipment and a large area blocked off from pedestrians, there is a good possibility for

approval.

With this information the entire senior design team took part in a drawing

exercise. Each member drew their own concept of the test stand. The drawings were

then passed around the table so that each person could make comments on the drawings.

No talking was allowed during this exercise. A Pugh analysis was then done on the most

promising designs, see Table 11. What evolved from the Pugh analysis was something in

a completely different direction than the team had started with. Instead of using a load

cell, the team settled on strain gauges mounted on a cantilevered beam to measure thrust

(Figure 19).

Pugh Analysis of Test Stand Concepts

Issues Cantilever Beam Adjustable Rollers Tracked System Manufacturability 3 1 2 Vibration Issue 3 3.5 1 Binding 3 2 2 Transportability 3 2 2 Durability 3 2 2 Different Diameters 3 2.5 3 Data Acquisition 3 3 4 Complexity 3 1 1.5 Redundancy 3 2 2 Measurements 3 3 3 Safety 3 3.5 3 Securing Rocket 3 2 2 Storability 3 2 2.5 Total 39 29.5 30

Table 11 - Pugh Analysis of Rocket Test Stands

At this point the desired location was changed once again to a field on the RIT

campus. Other locations were discussed such as the Geneseo Airport or a Civil Defense


location near campus. However, the main focus of the team is to procure a location on

campus due to the ease of transporting large amounts of heavy equipment and pressurized

tanks.

Figure 18 – Horizontal test stand concept


Figure 19 – Preliminary drawing of chosen test stand design.

6.2.3 Safety Considerations

The motor will be strapped to a length of angle stock to ensure it is straight.

When it fires the vertical beam will bend and the strain gauges will report a voltage that

can be turned into a thrust value. Three strain gauges will be placed on the side of the

beam that will be in tension. Each gauge should read the same, if they do not we know

that there was some twisting involved and can account for it. The vertical beam will be

clamped to a metal structure that will be attached via concrete anchor bolts to a slab of

concrete.

A completely new structure made of angle stock and lexan will be used to encase

the motor. The lexan will be used for three of the vertical sides and the top, the back will

be left open to allow the exhaust gases to exit. An exhaust deflector will be added to the


rear of the motor, just outside the lexan, to deflect the exhaust and anything that may

come from the combustion chamber upwards.

The test stand size was optimized after doing stress calculations for all bolts and

beams. In most cases dimensions of structural members were decreased and dimensions

of bolts were increased. The needed torque in all bolts is easily attainable with hand

tools.

All compressed gases will be housed in their own cage a few feet from the motor.

They will be controlled remotely by servos. This is an added safety feature because if the

motor gets loose it will lose its supply of Nitrous Oxide and quit burning. In regards to

safety, a complete safety analysis of the test stand and test procedure was completed to be

reviewed by the department head and the campus and local authorities.

Appendix 12 – Safety Report outlines the safety concerns of the team and actions

the team took to counteract them. This report has been submitted to facilities

management and we are awaiting approval to secure a test location on campus.

6.2.3.1 Built in Redundancies

To ensure the safety of the team and any observers during testing several

redundant safety measures were instituted into the procedure and set-up of our test stand.

By properly following the safety check list outlined in the safety report one can be

assured that the rocket will not fire until the team is prepared and safely away from the

rocket. The following is a list of the redundant measure taken by the team to ensure their

safety:

• Thorough stress calculations were performed on every critical component of the

test stand, outlined in section 7 of this document

• Throughout the duration of the test visual contact will be maintained in case an

emergency stop has to be instituted

• Both electronic and manual shut off valves are attached to all the tanks hooked up

the test chamber

• Pressure relief valves are attached to the test chamber and all tanks to ensure they

do not exceed the specifications of their respective cylinders


• The nitrogen purge tank can immediately smother any flame in the combustion

chamber if it is deemed necessary

• If the rocket happens to break free of the engine cradle the oxidizer hose will

disconnect effectively stopping the combustion process

• A deflection plate will be located behind the rocket during firing in case any

materials or components should happen to break free

• All possibly flammable material will be removed from the immediate area prior to

testing

• Multiple bolts and straps hold the beam and test chamber in place, all of which are

capable of independently serving their intended purpose

• Lexan (bullet proof polycarbonate material) and steel angle stock box will contain

all the components of the test stand, and keep them separate from the tanks

• The hoses and ignition system will be the last things attached to the rocket prior to

firing

• If the relief valves fail and the pressure inside the test chamber rises it will not be

able to reach the burst pressure of the test chamber cylinder because the high

pressure would inhibit the flow of oxidizer into the chamber

• Straps holding beam to cradle have the ability to compensate for temperature

fluctuations

6.3 Paper Design 6.3.1 Introduction

The last main deliverable laid out by the team at the start of Senior Design I was

the development of a preliminary paper design for the actual rocket. This design will be

very conceptual seeing how much of this design relies on what data we will obtain by

testing our rocket. Material selection and exact sizing will not be possible until the

conclusion of Senior Design II, and we can draw conclusions from our data. This section

of the paper is intended to spark ideas for future teams to develop upon.


6.3.2 Aluminum Truss System

- Connects the Thrust Plate to the Satellite Base Plate.

- Purpose: The purpose of the Aluminum Truss System is to focus all the loads

that can be expected in supporting the nose cone and satellite with a 30 g

acceleration to the thrust plate. The thrust plate is going to be the focal point

of our structural system and most likely made out of inconel or some other

materials with high corrosion resistance and strength properties. Figure 20

illustrates how the aluminum truss system attaches and supports the satellite

base plate to the thrust plate

Fig. 20 – Front and Isometric View of Aluminum Truss System, Satellite Base Plate, Thrust Plate and Satellite Release Mechanism Assembly

6.3.3 Satellite Base Plate

- Supported by aluminum truss system and acts as the base support of the pico-

satellite.

- Satellite release mechanism and aluminum trusses attached to the underside of

this plate.

- On top of this plate contains a cradle for the satellite to sit in. In addition, four

vertically placed compression springs are around the cover plate to assist in

the separation of the nose cone from the rocket body, see Figure 22 for the

basic concept behind our satellite containment and release.


6.3.4 Thrust / Injector Plate

- Is going to be made out of Inconel. Has high strength and thermal properties.

- Going to be of a circular shape.

- Top surface will have holes drilled for aluminum truss and oxidizer tanks

bracket connections. In addition, there will be miniature holes drilled in the

center for the oxidizer to flow through. The tiny holes will assist in the

atomization and spraying of the oxidizer into the combustion chamber. There

will be two stands on each side of the injector holes. These holes will be

tapped for the N2O connection.

- Bottom surface will contain two indented circular rings. Inside ring will

contain the combustion chamber assembly. The second ring will be the for

the insulation tubing.

6.3.5 Satellite Release Mechanism

- Components: Slider Pin, Internal housing, Compression spring, cap insert.

- Purpose: Attaches nose cone to rocket base.

- Putting it together: Insert pin into the compression spring, place pin/spring

assembly into the internal housing, screw-on cap insert into internal housing,

attach to cover base, attach hoses, and pressurize tanks

- Operation: The pressure from the oxidizer tank pushes the pin into the nose

cone and holds it in place (this compresses the spring). When the oxidizer is

used up, the pressure in the tank will decrease and the force of the spring will

overcome the pressure from the oxidizer tank and pushes the pin away from

the nose cone. The vertically placed compression springs that are placed on

top of the cover plate now uncompress. This forces an equal and opposite

reaction which should force the rocket to decelerate and the nose cone to

accelerate away from the satellite; while the satellite maintains its current

velocity.

- A problem brought up doing the peer design review was the exact timing of

the pin release. The thought was that one tank loses pressure faster than

another, thus causing a poor or early release of the nose cone. The idea to fix


this is to attach all pins to a throttling valve so each mechanism is receiving

the same amount of pressure.

- Components specifics:

a) Slider Pin: Will be covered in a Teflon coating to provide ease of

sliding and sealing so the N2O does not leak through.

b) Internal housing: Going to be made of aluminum square stock. The

sides will be milled downed and hole will be drilled in the milled flats

and then the internal housing can be attached to the underside of the

cover plate. There will be a total of four release mechanism. They will

be placed in between the aluminum trusses to conserve on the limited

space we have inside the rocket. There is a hole tapped on the backside

of housing for the oxidizer attachment and the front is drilled and

partially tapped for the cap insert.

Fig. 21 – Pressurized Satellite Release Mechanism

c) Cap insert: A circular cap that will contain external threads to be

screwed into the internal housing. A hole will be drilled in the middle

where pin will slide in and out. The main purpose for the cap insert is

for assembly purposes. Easy removal and installation of the spring/pin

assembly inside the internal housing.


d) Compression Spring: The spring will be place around the pin. All

springs used must be tested to make sure they have the same strength

and spring constants so the releasing of the pins occurs at the same time.

6.3.6 Satellite Containment

- Nose cone will be a uniform piece (no explosive bolts or splitting into two).

- Design of nose cone will be semi-conical shape (top of nose cone more flat

than pointy). Reasoning for this is due to our 90 degree angle brackets that

will act as the top support for the satellite as it sits in the cradle. If a full

conical shape, we would have to make additional supports to ensure the top

brackets where parallel with the satellite. The flat top makes sure that are

brackets are parallel with the satellite.

- Nose cone will contain four ledges towards the bottom. When attached to the

rocket body, they will cause the compression of the springs that are attached

to the cover plate. This supports in the separation of the nose cone from the

rocket body.

- Four slots will be drilled into the side of the nose cone where the pins will

lock into place.

- In the end, the nose cone will be pushed onto the rocket body; this causes the

satellite to be support by the cradle and brackets that are on top of the nose

cone. The springs will get compressed by the ledges, and when the pins are

pressurized, the nose cone will be fully attached to the rocket body and the

satellite will be held securely in place.


- Satellite will be placed inside a cradle that sits on top of the cover plat.

Fig. 22 – Pico-Satellite Containment and Release Schematic


6.3.7 Hybrid Engine Configuration

The overall configuration of the

hybrid engine for our rocket is highly

dependent on the environment that we

expect the rocket to be operating in.

Because the atmosphere is so thin at

100,000 feet and above drag effects on

our rocket will be minimal. Keeping

this consideration in mind, and the

length to diameter ratio of 10:1 we

selected for our fuel chamber the

configuration we chose greatly limits

the length of our rocket. See figure 23

for the overall configuration of our

rocket.

An additional benefit for

aligning the engine in this manner is t

future team will be able to utilize the

heat generated in the fuel chamber to

pre-heat the oxidizer in the tanks, whi

when properly designed will increase

the efficiency of our rocket.

hat

ch

Fig. 23 – Hybrid Rocket Configuration


7. Engineering Analysis / Design Validation

7.1 FEA Analysis 7.1.1 Introduction

The main purpose of the finite element analysis done on the test stand assembly

and combustion chamber is to evaluate the factor of safety in our design, obtain stresses,

strains, and deformation plots, and to compare all of these results with our hand

calculations.

Note from COSMOS Works:

Do not base your design decisions solely on the data presented in this report. Use

this information in conjunction with experimental data and practical experience. Field

testing is mandatory to validate your final design. COSMOS Works helps you reduce

your time-to-market by reducing, but not eliminating field tests.

7.1.2 Test Stand Analysis 7.1.2.1 Materials Table 12 lists of all the materials used in the test stand analysis and their properties are outlined in tables 13 and 14.

No. Part Name Material Mass Volume Weld beads to attached test base brackets to test base

1 AISI 1020 .107 lb .19 in^3

2 Base brackets AISI 1020 1.18 lb 4.14 in^3 Steel

3 Test Base AISI 1020 45.86 lb 160.67 in^3 SteelAISI 4140 4 Test Stand Post 4.70 lb 16.47 in^3 Steel

Table 12 - Test Stand Assembly Parts


AISI 1020

Property Name Value Units Value Type Elastic modulus 2e+011 N/m^2 Constant Poisson's ratio 0.29 NA Constant Shear modulus 7.7e+010 N/m^2 Constant Mass density 7900 kg/m^3 Constant Tensile strength 4.205e+008 N/m^2 Constant Yield strength 3.516e+008 N/m^2 Constant

AISI 4140

Table 13 – Material Properties of AISI 1020 Steel

Property Name Value Units Value Type Elastic modulus 2.05e+011 N/m^2 Constant Poisson's ratio 0.29 NA Constant Shear modulus 8e+010 N/m^2 Constant Mass density 7850 kg/m^3 Constant Tensile strength 6.55e+008 N/m^2 Constant Yield strength 4.15e+008 N/m^2 Constant

Table 14 – Material Properties of 4140 Annealed Steel 7.1.2.2 Loading and Restraints

For our analysis, we placed a fixed restraint on the test base because it will be securely

fixed to our concrete bed. For loading, we placed a 200 lb force at the top of our test post

because that is the force that the test beam at maximum should receive and 2800 lb force

at each base bracket hole to represent the bolt loading on the test beam.


7.1.2.3 Study Properties

Here are charts about our meshing information and solver techniques COSMOS Works

used in its analysis of our test stand:

Mesh Information Mesh Type: Solid mesh Mesher Used: Standard Automatic Transition: Off Smooth Surface: On Jacobian Check: 4 Points Element Size: 0.28435 in Tolerance: 0.014217 in Quality: High Number of elements: 42232 Number of nodes: 77961

Table 15 – FEA Analysis Mesh Information

Solver Information Quality: High Solver Type: FFE Option: Include Thermal Effects Thermal Option: Input Temperature Thermal Option: Reference Temperature at zero strain: 298 Kelvin

Table 16 – FEA Analysis Solver Information


7.1.2.4 Stress Results Below contains a plot of the stress seen in the test stand as well as where the maximum

and minimum stresses are located:

Type Min Location Max Location

(-8.11398 in, (1.19852 in, 0 psi 17847.5 psi VON: von Mises stress Node:

11563

4.66527 in, 6.80813 in, Node: 77705 -3.48033 in) 4.14467 in)

Table 17 – Location of Maximum and Minimum Stress on Test Beam

Figure 24 – Stress Distribution on Test Beam


7.1.2.5 Strain Results Below contains a plot of the strain seen in the test stand as well as where the maximum

and minimum strains are located:


(-0.0790332 in, (0.325503 in, 0 ESTRN: 0.000387696 Equivalent strain

Element: 6587

4.62979 in, 6.18573 in, Element: 2160 7.4947 in) 5.37275 in)

Table 18 – Location of Maximum and Minimum Strain on Test Beam

Figure 25 – Strain Distribution on Test Beam


7.1.2.6 Displacement Results Below contains a plot of the displacement seen in the test stand as well as where the

maximum and minimum displacements are located:


(1.69852 in, (0.648524 in, 0 in 0.0811551 inURES: Resultant

displacement Node: 1 4.66527 in, 16.6653 in,

Node: 69271 7.64467 in) 5.51967 in) Table 19 – Location of Maximum and Minimum Displacement on Test Beam

Figure 26 – Displacement of Test Beam


7.1.2.7 Design Check Results Below contains multiple plots of the factor of safeties seen in the test stand assembly:

Figure 27 - Yield Factor of Safety Distribution on Test Beam


Figure 28 - Ultimate Factor of Safety Distribution on Test Beam

7.1.2.8 Conclusion After reviewing the design check plots, we feel our design is sufficient enough for actual

testing. The reasoning behind this due to the test post FOS distribution. The results

given to us by COSMOS Works closely match our hand calculations results; which

supports our design criteria. The lowest yield factor of safety in our test post occurs at

the test beam with a value of 3.4. The lowest ultimate factor of safety of our test post is

5.3.


7.1.3 Combustion Chamber Analysis

7.1.3.1 Materials

Here is a list of all the materials used in the combustion chamber analysis and their

properties:

No. Part Name Material Mass Volume 1 Combustion Chamber AISI 304 9.98261 kg 0.00124783 m^3

Table 20 – Materials Used in Test Chamber AISI 304

Property Name Value Units Value Type Elastic modulus 1.9e+011 N/m^2 Constant Poisson's ratio 0.29 NA Constant Shear modulus 7.5e+010 N/m^2 Constant Mass density 8000 kg/m^3 Constant Tensile strength 5.1702e+008 N/m^2 Constant Yield strength 2.0681e+008 N/m^2 Constant Thermal expansion coefficient 1.8e-005 /Kelvin Constant Thermal conductivity 16 W/(m.K) Constant Specific heat 500 J/(kg.K) Constant

Table 21 – Material Properties of AISI 304 Stainless Steel

7.1.3.2 Loading and Restraints

For our analysis, we placed a fixed restraint at the two end faces of the combustion

chamber because those faces will contain additional components and will be fixed in

place. For loading, we placed a 1000 psi pressure inside the chamber because that should

be the maximum pressure the chamber should see. Note: We have placed pressure relief

values that are activated when the chamber rises to a pressure of 750 psi.


7.1.3.3 Study Properties

Here are charts about our meshing information and solver techniques COSMOS Works

used in its analysis of our test stand:

Mesh Information Mesh Type: Solid mesh Mesher Used: Standard Automatic Transition: Off Smooth Surface: On Jacobian Check: 4 Points Element Size: 0.20958 in Tolerance: 0.010479 in Quality: High Number of elements: 56540 Number of nodes: 90790

Table 22 – FEA Analysis Mesh Information of Test Chamber

Solver Information Quality: High Solver Type: FFE Option: Include Thermal Effects Thermal Option: Input Temperature Thermal Option: Reference Temperature at zero strain: 298 Kelvin

Table 23 – FEA Analysis Solver Information of Test Chamber


7.1.3.4 Stress Results

Below contains a plot of the stress seen in the combustion chamber as well as where the

maximum and minimum stresses are located:


(2.14313e-016 in, (1.30275 in, 204.698 psi 10099.1 psi VON: von Mises

stress Node: 398 7.6358 in, -5.516 in,

Node: 61389 1.75 in) 3.15715e-007 in)

Table 24 – Location of Maximum and Minimum Stress on Test Chamber

Figure 29 – Stress Distribution on Test Chamber


7.1.3.5 Strain Results

Below contains a plot of the strain seen in the combustion chamber as well as where the

maximum and minimum strains are located:


(1.62249 in, (1.30039 in, 0.000253603 9.45831e-006 ESTRN: Equivalent strain Element: 40137

-7.7927 in, -5.5279 in, Element: 37699 0.0255057 in) 0.0388766 in)

Table 25 – Location of Maximum and Minimum Strain on Test Chamber

Figure 30 – Strain Distribution on Test Chamber


7.1.3.6 Displacement Results

Below contains a plot of the displacement seen in the test stand as well as where the

maximum and minimum displacements are located:


(0 in, (0.284063 in, 0 in 0.000163155 inURES: Resultant displacement Node:

86

7.845 in, -5.74537 in, Node: 52465 1.375 in) 1.2173 in)

Table 26 – Location of Maximum and Minimum Displacements on Test Chamber

Figure 31 – Dislocations of Test Chamber


7.1.3.7 Design Check Results Below contains multiple plots of the factor of safeties seen in the test stand assembly:

Figure 32 – Cross Sectional View of Yield Factors of Safety in Test Chamber


Figure 33 - Cross Sectional View of Ultimate Factors of Safety in Test Chamber

7.1.3.8 Conclusion

After reviewing the design check plots, we feel our design is sufficient enough for actual

testing. As predicted the failure does occur at hole for our listening tube; however, the

lowest yield factor of safety of our design is 3, which closely relates to our hand

calculations. The ultimate factor of safety is 7.4; which also occurs at the hole for our

listening tube.


7.2 Rocket Calculations

7.2.1 Rocket Sizing and Thrust Calculations

The following section walks you through the preliminary calculations performed

by the team in order to predict the necessary size, mass, and mass flow rate to generate

the needed values outlined in the specification portion of this paper (Section 5).

7.2.1.1 Givens and Assumptions: Isp = 235 sec d = 0.03 m i

10=id

LVtotal = 9200 m/s for complete combustion

mL = 1 kg ρ = 930 kg/m3HTPB

m = 0.75 kg as max = 30g 2melec = 0.3 kg g = 9.81 m/s

E = 20.557 kJ/mol A = 11.04 mm/s a

T R = 1000 K = 8.3143 J/(mol-K) s

M =0.1 kg/mol 7.2.1.2 Mass Estimation:

smve 2300=( ) ⎟

⎠⎞

⎜⎝⎛= 281.9sec235

smve gIv spe =

4

9200sm

v =ΔN

vv total=Δ

smv 2300=Δ

kgkgkgm 3.075.00.11 ++=elecsL mmmm ++=1 kgm 05.21 =

( )⎟⎟⎟⎟

⎠

⎞

⎜⎜⎜⎜

⎝

⎛

⋅=

smsm

kgm2300

2300exp05.20⎟⎟

⎠

⎞⎜⎜⎝

⎛ Δ⋅=

evvmm exp10⎟⎟

⎠

⎞⎜⎜⎝

⎛=Δ

1

0lnmm

vv e

kgm 572.50 =

10 mmm p −= kgm p 522.3= kgkgm p 05.2572.5 −=


7.2.1.3 Regression Rate:

⎟⎟⎟⎟

⎠

⎞

⎜⎜⎜⎜

⎝

⎛

⋅⋅

−⋅=

KKmol

JmolkJ

smmr

10003143.8

557.20exp04.11⎟⎟

⎠

⎞⎜⎜⎝

⎛⋅

−⋅=

s

a

TRE

Ar exp

smmr 9315.0= However, calculation of regression is very inexact. The formula

used above is very general and based off experiments done by Chiaverini. Therefore, a regression rate of 1 mm/s will be assumed for sizing calculations of the initial rocket

engine. More exact calculations can be completed following testing.

7.2.1.4 Sizing: ( )( )mL 03.010=idL ⋅= 10 mL 3.0=

After 1 second of burning:

( sec11203.00 ⎟⎠⎞

⎜⎝⎛+=

smmmd( )( )trdd i 20 += ) md 032.00 =

( )LddV iHTPB22

04−=

π ( ) ( )( )( mmmVHTPB 3.003.0032.04

22 −= )π

3510922.2 mVHTPB−×=

( )353 10922.2930 m

mkgmHTPB

−×⎟⎠⎞

⎜⎝⎛=&VmHTPB ρ=&After one second,

skgmHTPB 0272.0=&

( )kgmHTPB 522.391⎟⎠⎞

⎜⎝⎛=pHTPB mm ⎟

⎠⎞

⎜⎝⎛=

91Oxidizer-to-fuel ratio = 8:1

kgmHTPB 3913.0=

HTPB

HTPBb m

mt&

=

skgkgtb

0272.0

3913.0= sec4.14=bt

Back-Calculate to obtain an outer diameter for the grain size:

( )sec4.141203.00 ⎟⎠⎞

⎜⎝⎛+=

smmmd( )( )bi trdd 20 += md 0588.00 =

Increase outer diameter to allow extra fuel to insulate the combustion chamber walls and to utilize standard pipe sizes: md 0635.00 =


7.2.1.5 Oxidizer Mass Flow Rate:

( )

sec4.14

522.398 kg

mox

⎟⎠⎞

⎜⎝⎛

=&b

p

ox t

mm

⎟⎠⎞

⎜⎝⎛

= 98

& s

kgmox 2174.0=&

7.2.2 Exit Nozzle Shaping 7.2.2.1 Given and Assumptions: Total Fuel Burned, m = 3.5225 kg fTarget Burn Time, t = 15 seconds bMass Flowrate, m = 0.235 kg/s & Assuming the following conditions: T = 3300 K 0P = 3.8 MPa 0γ = 1.20 7.2.2.2 Area Sizing: Where subscript “t” denotes nozzle throat conditions:

1

0

1.201.20 16

112

1.20 13.8 10 12

2.145

t

t

t

P P

P Pa

P MPa

γγγ−−

−−

−⎛ ⎞= −⎜ ⎟⎝ ⎠

−⎛ ⎞= × +⎜ ⎟⎝ ⎠

=

01

112

13300 1.20 112

3000

t

t

t

T T

T K

T K

γ=−

+

=−

+

=


12

12

6

5 2

0.235 8314.4 3000

2.145 10 30 1.20

9.1194 10

tt

t gas

t

t

R TmAP M

kg J Ks kmolA kgPa

kmolA m

γ

−

⎛ ⎞⎛ ⎞ ∗= ⎜ ⎟⎜ ⎟ ⎜ ⎟∗⎝ ⎠ ⎝ ⎠

⎛ ⎞ ⎛ ∗⎜ ⎟ ⎜= ⎜ ⎟ ⎜×⎜ ⎟ ⎜ ∗⎝ ⎠ ⎝

= ×

&

⎞⎟⎟⎟⎠

Where Mgas is the molecular weight of the exiting gasses, which for these calculations was assumed to be 30, which represents a mixture of equal parts CO2, H O, and N . 2 2 The throat diameter is then:

= 0.010776 m Dt = 1.078 cm Dt

For the ground test nozzle, the ambient pressure, Pa was assumed to be 101.3 kPa. The mach number of the exiting gasses for the given conditions is then:

1

2 0

1.20 16 1.20

23

2

2 11

2 3.8 10 11.20 1 101.3 10

8.2962.88

ea

e

e

e

PMP

M

MM

γγ

γ

−

−

⎡ ⎤⎛ ⎞⎛ ⎞ ⎢ ⎥= −⎜ ⎟⎜ ⎟ ⎢ ⎥−⎝ ⎠ ⎝ ⎠⎢ ⎥⎣ ⎦⎡ ⎤⎛ ⎞×⎛ ⎞ ⎢ ⎥= −⎜ ⎟⎜ ⎟ ⎢ ⎥− ×⎝ ⎠ ⎝ ⎠⎢ ⎥⎣ ⎦

=

=

The cross-sectional area at the exit can be found in the following manner:

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Giving us an exit diameter of: De = 0.025721 De = 2.572 cm 7.2.2.3 Exit Velocity: Finally, the velocity of the exiting gasses can be found using the following equation:

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⎡ ⎤⎛ ⎞⎛ ⎞⋅⎢ ⎥⎜ ⎟ ⎛ ⎞×⎛ ⎞ ⎜ ⎟= −⎢ ⎥⎜ ⎟ ⎜ ⎟⎜ ⎟ ⎜ ⎟− ×⎝ ⎠ ⎝ ⎠⎜ ⎟⎢ ⎥⎜ ⎟⎝ ⎠⎢ ⎥⎝ ⎠⎣ ⎦

=


8. Senior Design II 8.1 Deliverables Below are the primary and secondary objectives for our design team to complete

based on discussions with our mentors and evaluation of our situation at the end of Senior

Design I. The primary focus of our team will be on collecting as much data and

information as possible on our hybrid engine and all of its components. We will need to

order several of our parts prior to the completion of Senior Design I and this document.

8.1.1 Primary Deliverables at End of SD II

• Operating Cantilevered Beam Test Stand

• Design for Back-up Cart & Track / Load Cell Test Stand

• Nozzle Design for Vacuum and Atmospheric Pressure

• Injector / Thrust Plate for Test Chamber

• Experimentally Determined Internal Temperatures and Pressures

• Experimentally Measured Thrust

• Reliable and Safe Feed System for Testing

• Reliable and Safe Ignition System

• Propellant Experience

o Molding

o Sizing

o Regression rates

o Oxidizer flow

o Burn temperature

o Average volumetric flow rate

8.1.2 Secondary Objectives

• Continue Paper Design

• Design Fuel Chamber Made of Composites

• Test Different Solid Fuel Patterns

o Maximum impulse


o Constant Thrust

• Stage Separation Devices

• Sizing of All Stages of Actual Rocket

• Locate Electronically Controlled Regulatory Valve

• Locate Valves and Components for Actual Rocket Feed System

8.2 Future Plans 8.2.1 Time Table

Appendix 6 outlines a tentative schedule for Senior Design II in the form of a

Gantt Chart. Much work needs to be done with the materials we are ordering, in

particular the propellants and ignition materials, before we can confidently assure a safe

rocket firing. The following sections outline a few of these preliminary tests that need to

be completed.

8.2.2 Propellant Testing

8.2.2.1 Molding

The solid propellant, HTPB, and curative mixture that we have selected has only a

two hour cure time, depending on how much curative is added. This will be especially

advantageous as we conduct preliminary testing on different molding techniques. At the

moment we are considering PVC or a cardboard mold lined with saran wrap to ensure the

HTPB does not stick to the walls of the mold. In addition to developing the mold we will

have to develop a safe method of inserting the fuel grain into the chamber.

8.2.2.2 Combustion

In addition to settling on a method of molding and curing the HTPB the team will

need to gain experience with the burning tendencies of the propellant. Tests will be

performed to determine the temperature and rate of combustion under ambient

conditions. When these have been determined we can better assess what to expect from

the combustion process inside the chamber based on expected temperatures.


8.2.3 Ignition Testing

8.2.3.1 Glow Plug Ignition

To test the validity of the glow plug method of ignition many different tests will

need to be conducted. First we must figure out what kind of flow rate we can run across

the plug and still get disassociation of the Nitrous Oxide. Furthermore we will need to

determine if we are able to ignite the flow with the plug, or if an additional spark will be

required. Then we will have to see if this idea is capable of initiation the combustion

process and interact with HTPB.

8.2.3.2 Pyrotechnic Ignition

Testing with the pyrotechnic method of ignition will include trying several

different mixtures of pyrotechnics, many different schematics of Ni-Chrome wire and a

thorough investigation of the burn properties of NC Lacquer and Ammonium Perchlorate.

The goals of these tests are to determine the most reliable method, but also to ensure that

the schematic we choose will not potentially damage any other components in the test

chamber.

8.2.4 Data Acquisition Component Testing

The set up and collecting of useful data from our data acquisition system was one

component of our design that the team assumed could be easily controlled an performed.

After discussing the scope and ambitions of our desired test results with Dr. Wellin and

different people in industry we soon came to realize that we mistakingly overlooked

many of the details in organizing our system. A focus during the first few weeks will be

placed on gaining experience with the thermocouples, pressure transducers, signal

converters, and all the other components included in a sound DAQ schematic. In

addition the students will need to gain experience with the LabView interface, and

inputting useful data.

8.2.5 Feed System / Oxidizer Flow

Understanding the flow, and verifying our expected pressure drops across

different components is an essential part of predicting the behavior of our rocket.


Particular attention will be given to the pressure drop and atomization of the flow at

different rates through our injector nozzle. We must ensure that we are able to produce

the desired flow rate through the nozzle to keep an optimum oxidizer to fuel ratio inside

the test chamber.

8.2.6 Combustion Chamber Tests

Tests will have to be completed on the Test chamber to ensure that it will be able

to contain the worst case pressures that we can expect based on the preliminary tests

listed in the previous sections of 8.2. Attaching the injector / thrust plate and raising the

pressure inside the chamber will be sufficient enough to validate our design.

8.2.7 Test Stand Materials

The test stand must be assembly and loads given for worst case scenarios will be

applied near the location that the rocket would be attached. The team will look for any

yielding of the material, or any bolts of welds that look questionable. Testing these

components in a safe environment to verify our calculations is a good practice because

we do not want them to fail while the rocket test chamber is attached.

8.3 Testing the Rocket After all of the preliminary tests have been completed in section 8.2 and the team

and our advisors are confident that the system we have designed and tested is completely

safe for all observers and the environment we will go forward with testing our assembled

test chamber. There is no room for questionable performance when dealing with our

system, and a thorough investigation of all components must be completed before firing

our rocket. Safety is the biggest concern for our team, mentors and RIT. In order for the

METEOR project to be successful and continue on campus we must put forth a good

example of a sound analysis and regard for safety.

8.3.1 Test Results

After successfully completing tests on our rocket, the team is then responsible for

interpreting the data collected and drawing conclusions. Ideally the data we collect will


be sufficient to complete the deliverables outlined in section 8.1.1 of this report. In

addition to completing these deliverables the team will be responsible for outlining future

tests to be completed, and suggestions for the direction of future METEOR rocket teams.

8.4 Budget Throughout the process of developing our design and weighing its feasibility

against estimated specifications the teams main focus has been on safety. Another

consideration that had influence on our decision making process was the budget allotted

to our team. It is very easy to design a hybrid engine with multiple high tech devices

which may improve or optimize our design, however dealing with a constrained budget

the team had to come up with some innovative methods of avoiding some of these costly

materials. Our team was given a maximum budget of $10,000 for all of our testing

materials and data acquisition devices. A great emphasis is placed on the word maximum

when referring to our budget, because the funding we are receiving comes out of a

collective fund which is supporting the entire METEOR project here on campus. If our

team wants to lay the ground work for future teams to succeed leaving them any

additional money that we can afford to pass up will only further ensure the success of the

METEOR here on campus.

As seen in our Bill of Materials in Appendix 2 the minimal amount our team

expects to spend is around $4500. This is less than half the amount of money we are

allowed to spend. This budget will increase because our team will be ordering multiple

parts, especially in the feed system and data acquisition components in case of any

damaged parts during testing. Taking this in consideration we still don’t expect our

budget to rise above $6000. One thing that may be suggested at the end of our

examination during Senior Design II is looking at some more technical, or temperature

resistant data acquisition sensors which could improve the experimental data we collect.


References

1. Patru, D., J. Kozak, and R. Bowman, The METEOR Lab. 2005, Rochester Institute of Technology: Rochester, NY.

2. Pegasus User's Guide. 2000, Orbital Sciences Corporation. p. 103. 3. CubeSat Community Website. [cited 2006 18 Feb]; Available from:

http://littonlab.atl.calpoly.edu/. 4. Engineered Films - High Altitude Balloons. [cited 2005 12 Dec]; Available

from: www.ravenind.com/RavenCorporate/eng_films/high_alt_balloons_index.htm.

5. Patru, D., J. Kozak, and R. Bowman, Pico-Satellite Launch System and Pico-Satellites: Design Concept. 2005, Rochester Institute of Technology: Rochester, NY.

6. Boltz, F.W., Low-Cost Small-Satellite Delivery System. Journal of Spacecraft. 39(5): p. 818-820.

7. Alexander, B., et al., Converting the Minuteman Missile into a Small Satellite Launch System, University of Texas at Austin: Austin, TX.

8. Baker, A.M., M. Heywood, and R. Newlands, The Hybrid Engine as a Green Propulsion Unit for Amateur Rockets. ESA SP., 2001(484): p. 327-334.

9. Krauss, O., Design and test of a Lab-Scale N2O/HTPB Hybrid Rocket, University of Colorado at Boulder: Boulder, CO. p. 11.

10. Boltz, F.W., Optimal Ascent Trajectory for Efficient Air Launch into Orbit. Journal of Spacecraft. 41(1): p. 153-157.

11. AGI: Analysis for Land, Sea, Air, and Space. [cited 2006 Feb 23]; Available from: www.agi.com.

12. Sutton, G.P. and O. Biblarz, Rocket Propulsion Elements. 7th ed. 2000: Wiley-Interscience. 751.

13. Cappola, J.A., The Study and Selection of Rocket Propellants for Launching Pico-Satellites. 2005, Rochester Institute of Technology: Rochester, NY.

14. Hill, P. and C. Peterson, Mechanics and Thermodynamics of Propulsion. 2nd ed. 1992, New York: Addison-Wesley Publishing Company. 760.

15. Chiaverini, M.J., et al., Regression Rate Behavior of Hybrid Rocket Solid Fuels. Journal of Propulsion and Power, 2000. 16(1).

16. Chiaverini, M.J., et al., Regression-Rate and Heat-Transfer Correlations for Hybrid Rocket Combustion. Journal of Propulsion and Power, 2001. 17(1).

17. Werthman, W.L. and C.A. Schroeder. A Preliminary Design Code for Hybrid Rockets. in 32nd Aerospace Sciences Meeting & Exhibit. 1994. Reno, NV: American Institute of Aeronautics and Astronautics.

18. Greiner, B. and J. R. A. Frederick. Results of Labscale Hybrid Rocket Motor Investigation. in AIAA/SAE/ASME/ASEE 28th Joint Propulsion Conference and Exhibit. 1992. Nashville, TN: American Institute of Aeronautics and Astronautics.

19. Hybrids, Part 1. [cited 2006 14 Feb]; Available from: http://www.hawkfeather.com/rockets/hybrids1.html.


http://littonlab.atl.calpoly.edu/

http://www.ravenind.com/RavenCorporate/eng_films/high_alt_balloons_index.htm

http://www.agi.com/

http://www.hawkfeather.com/rockets/hybrids1.html

20. Shanks, R. and M.K. Hudson, A Labscale Hybrid Rocket Motor for Instrumentation Studies. Journal of Pyrotechnics, 2000(11).

21. McCormick, et al., Design, Optimization, and Launch of a 3" Diameter N20/Aluminized Parafin Rocket. 2005, Stanford University: Stanford, CA.

22. Mungas, D. and Kulkarni, Design, Construction and Testing of a Low-Cost Hybrid Rocket, in Aircraft Engineering and Aerospace Technology. 2003. p. 262-71.

23. Jacobson, J. Ammonium Perchlorate - Fuel Oxidizer. 2006 [cited 2006 13 Feb.]; Available from: http://web1.caryacademy.org/chemistry/rushin/StudentProjects/CompoundWebSites/2000/AmmoniumPerchlorate/Oxidizer.html.

24. Waidmann, W., Thrust Modulation in Hybrid Rocket Engines. Journal of Propulsion and Power, 1988. 4(5): p. 421-7.

25. Tamura, T., S. Yuasa, and K. Yamamoto. Effects of Swirling Oxidizer Flow on Fuel Regression Rate of Hybrid Rockets. in 35th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. 1999. Los Angeles, CA: American Institute of Aeronautics and Astronautics.

26. Anderson, J.D., Fundamentals of Aerodynamics. Third ed. 2001: McGraw-Hill. 912.

27. Federal Aviation Regulations Part 101. 1958, Federal Aviation Administration. 28. US Military Specifications (MIL). [cited 2006 15 Jan]; Available from:

www.combatindex.com/mil_docs/pdf/. 29. Newlands, R., The physics of Nitrous Oxide. 2004, AspireSpace. 30. Priming - Pyro Universe. [cited 2006 23 Feb]; Available from:

http://www.pyrouniverse.com/fusemaking/priming.htm. 31. Fox, R.W., A.T. McDonald, and P.J. Pritchard, Introduction to Fluid Mechanics.

Sixth ed. 2003: Wiley.


http://web1.caryacademy.org/chemistry/rushin/StudentProjects/CompoundWebSites/2000/AmmoniumPerchlorate/Oxidizer.html

http://web1.caryacademy.org/chemistry/rushin/StudentProjects/CompoundWebSites/2000/AmmoniumPerchlorate/Oxidizer.html

http://www.combatindex.com/mil_docs/pdf/

http://www.pyrouniverse.com/fusemaking/priming.htm

Appendix 1 – Drawing Package

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Appendix 2 – Bill of Materials

McMaster Carr Parts ListBolts - Lexan Frame 1/4 x 1.5" 65 92865A546 McMaster $9.03 100 $9.03Bolts - Injector Plate 1/4 x .375" 8 92185A535 McMaster $4.10 10 $4.10Nuts 1/4" - 20 73 95462A029 McMaster $2.14 100 $2.14Locking Washers 1/4" spring lock 73 91104A029 McMaster $2.09 100 $2.09Test Beam 12 x 3.0 x .375" 1 6554K231 McMaster $17.73 1 $17.73Test Beam 12 x 2.5 x .5" 1 6554K87 McMaster $21.34 1 $21.34Anchor Bolts .5" diameter 42 92914A640 McMaster $1.97 50 $98.50Angle Stock - cradle 3" Leg x < 3' 1 9017K73 McMaster $31.93 7 $223.51 - lexan frame 3" Leg x 3' 11Deflector Plate - Frame 1.25" Sq tube 3 6527K14 McMaster $18.69 3 $56.07 - Shield 2' x 2' x .063" 1 6544K16 McMaster $25.30 1 $25.30 - Support 2' x 2' mesh 1 9302T124 McMaster $12.18 1 $12.18Base Brackets 10"x10"x.5" 1 6544K36 McMaster $50.16 1 $50.16Engine Cradle Brackets 10"x10"x.25" 1 6544K24 McMaster $31.30 1 $31.30Steel Tube OD=3.5; ID=2.5; t=.5" 2 89495K453 McMaster $365.15 2 $730.30Snap Ring Assy BD=2.5"; t=.078" 2 91580A253 McMaster $18.55 3 $55.65Garolite Laminated Ceramic Ring OD=2.5; ID=2.0; L=48" 1 87285K57 McMaster $115.08 1 $115.08Injector plate (304 ss) D = 3.5; t = 2" 1 9208K62 McMaster $35.62 2 $71.24O-rings OD = 2.75; t = .06" 2 9396K118 McMaster $6.44 10 $6.44Brass Ball Valve 3/8" FNPT 1 46495K59 McMaster $22.03 2 $44.06Brass Relief Valve (750 psig) 1/4" MNPT 1 5825T21 McMaster $80.14 2 $160.28Brass Pipe Nipple 3/8" MNPT x 2" Lg. 5 50785K209 McMaster $2.60 7 $18.20Teflon Coated Braided Hose (2' Lg.) 1/4 MNPT 1 4468K402 McMaster $19.03 1 $19.03Teflon Coated Braided Hose (10' Lg.) 3/8" MNPT 1 4468K403 McMaster $50.26 1 $50.26Check Valve (Buna-N Seat) 3/8" FNPT 2 7775K13 McMaster $17.51 3 $52.53Check Valve (Buna-N Seat) 1/4" FNPT 1 7775K12 McMaster $11.34 2 $22.68Brass Tee Fitting 3/8" FNPT x 3/8" FNPT x 3/8" FNPT 3 50785K73 McMaster $3.29 4 $13.16Brass Needle Valve 3/8" FNPT 1 46425K13 McMaster $30.33 2 $60.66Brass Tee Fitting 1/4" FNPT x 1/4" FNPT x 1/4" FNPT 1 50785K72 McMaster $2.17 1 $2.17Brass Elbow 3/8" FNPT 1 50785K37 McMaster $2.24 3 $6.72Brass Hex Fitting 3/8" MNPT 3 5485K231 McMaster $1.64 5 $8.20Stainless Steel Pipe 1/4" MNPT x 6" Lg. 1 46755K72 McMaster $10.35 1 $10.35Stainless Steel Pipe 1/4" MNPT x 12" Lg. 1 46755K112 McMaster $19.68 1 $19.68Brass Reducer Coupling 3/8" FNPT to 1/4" FNPT 2 50785K183 McMaster $2.05 4 $8.20Brass Rod (For Thread Adapters) Raw Material (1' Lg.) 1 8970K781 McMaster $35.45 1 $35.45Brass Union 1/4" FNPT 1 50785K243 McMaster $11.46 1 $11.46Brass Reducing Hex Fitting 3/8" MNPT to 1/4" MNPT 1 5485K321 McMaster $1.71 2 $3.42Pipe Thread Sealant Tape Roll Teflon Tape (1/4" Wide) 1 6802K11 McMaster $1.95 3 $5.85Pressure Gauge, Dual Scale (0-1000 psi) 1/4" MNPT 2 4000K713 McMaster $10.04 2 $20.08

$2,104.60

AeroCon Parts ListHTPB and Papi 94 1 Gal. / 1 pint 6 HTPB Kit Aerocon $56.00 6 $336.00Graphite Nozzle D = 3.0" x 12" 2 Graphite Aerocon $54.00 3 $162.00Nichrome Wire 40 gauge x 30ft 1 40 Gauge NiCr 60 Aerocon $9.00 1 $9.00Nichrome Wire 34 gauge x 30ft 1 34 Gauge NiCr 60 Aerocon $6.00 1 $6.00HotHead Electric Match Heads Chip / 50 gauge nichrome 10 HotHead Elec Match Aerocon $9.00 50 $9.00*** Hold off on ordering highlighted bottles until we talk to N2O vendor to see how tanks come $522.00

Skylighter Parts ListAmmonium Perchlorate 1 lb. 2 CH5000 Skylighter $9.15 2 $18.30Nitrocellulose Lacquer 1 qt. 1 CH8198 Skylighter $29.65 1 $29.65

$47.95

Home DepotBolts - Test Beam / Cradle 1/4 x 2" 8 Home DepotCasting Tubes Various PVC pipe Home Depot

CoAx Parts ListCoaxial Solenoid Valve (Oxidizer) 3/8" FNPT 1 KB152C880VTN3/8TZA CoAx $331.00 2 $662.00

Omega Parts ListThermocouples E Type 2 CO1-E Omega $31.00 2 $62.00IR Temperature Transmitter 0-400 oF 4 OS136-1-V2 Omega $175.00 5 $875.00IR Temperature Transmitter 300-100 oF 1 OS136-2-V2 Omega $175.00 1 $175.00Pressure Snubber Liquid / H2O 2 PS-4E Omega $10.00 2 $20.00Pressure Snubber Air / Gas 1 PS-4G Omega $10.00 1 $10.00Heat Flux Sensor K-type 1 HFS-3 Omega $130.00 1 $130.00Pressure Transducer 1/4" MNPT 2 PX302-1KGV Omega $185 3 $555.00

$1,827.00

CotronicsCastable Ceramics 10 lb 1 740 Castable Ceramic Cotronics $79.95 1 $79.95Adhesive Backed Ceramic Tape 1/32" X 1" X 50' 1 397-21PS Cotronics 22.65$ 1 $22.65Adhesive Backed Ceramic Tape 1/32" X 2" X 50' 1 397-22PS Cotronics 39.95$ 1 $39.95Adhesive Backed Ceramic Tape 1/16" X 1" X 50' 1 397-41PS Cotronics 26.95$ 1 $26.95

Appendix 3 – Gantt Chart SD I

Appendix 4 – Risk Assessment

Risk Assesment

PROJECT FEASABILITY ASSESSMENT - METEOR ROCKET DESIGN (P06006) - Week 4

1 (LOW)

2 3 (MED)

4 5 (HIGH)

RATIONALE REMEDY/REQUIRED ACTION

RESOURCE FEASIBILITY

1) SUFFICIENT TEAM MEMBERS SKILLS X R1: COMPUTER SKILLS AND EDUCATIONAL BACKGROUND IN PROJECT AREA MAY BE LACKING

R1: SUPPLY TEAM WITH REQUIRED READING MATERIAL AT SOON AS POSSIBLE FOR BACKGROUND

2) SUFFICENT SHOP AVAILIBILTY XR2: MACHINING TECHNIQUES FOR ROCKET BODY MAY REQUIRE TOOLS OR TECHNOLOGY NOT AVAILABLE ON RIT CAMPUS

R2: LOOK INTO HIRING COMPANIES TO CONSTRUCT PARTS WITH LOW TOLERANCES OR DETAILED MATERIALS

3) SUFFICENT NUMBER OF PEOPLE X R3: TEAM SHOULD BE MECHANICALLY SOUND, MAY NEED TO RESEARCH ELECTRICAL CONTROL OF ENGINE IGNITION

R3: DETERMINE COMPLETE PROJECT REQUIREMENTS AND DISCUSS POSSIBLE NEEDS WITH RIT FACULTY

4) SUFFICENT GUIDANCE FROM MENTOR X R4: ALL MENTORS AND SPONSORS ARE LOCATED ON CAMPUS R4: NONE REQUIRED

5) SUFFICENT ASSEMBLY AREA X R5: ROCKET SIZE IS NOT SUFFICIENT ENOUGH TO RUN INTO ASSEMBLY SPACE PROBLEMS

R5: ACQUIRE PRIVILIDGES TO A ROOM OR LAB WHERE THERE IS SPACE TO STORE OUR PARTS AND SUPPLIES IMMEDIATELY

6) SUFFICENT COMPUTER TOOLS X R6: SOME SOFTWARE THAT MAY BE REQUIRED AT DIFFERENT POINTS OF DESIGN PROCESS

R6: SPEAK WITH FACULTY ADVISORS AND SEE IF THERE ARE ANY SOFTWARE PACKAGES WE NEED TO ACQUIRE

7) SUFFICENT MECHANICAL TOOLS X R7: MAY REQUIRE SUBASSEMBLIES TO BE CONSTRUCTED OUTSIDE RIT

R7: DETERMINE PARTS NECESSARY ON MORE TECHNICALLY ADVANCED FUNCTIONS AS SOON AS POSSIBLE

8) SUFFICIENT INPUT FROM SPONSOR X R8: ALL MENTORS AND SPONSORS ARE LOCATED ON CAMPUS R8: NONE REQUIRED

9) SUFFICIENT TESTING DEVICES XR9: GROUND TESTING ENGINE REQUIRES SAFE ENVIRONMENT AND HIGH TECH SENSORS TO PROPER TESTING

R9: DETERMINE WHAT TESTS NEED TO BE RUN ON HYBRID ENGINE WHILE IN USE TO PREDICT FUNCTION QUICKLY AND ORDER MATERIALS

10) SUFFICIENT TESTING LOCATIONS XR10: GROUND TESTING ENGINE REQUIRES REMOTE AREA THAT CAN BE MADE SAFE, ENVIRONMENTALLY AND PHYSICALLY

R10: A LOCATION ON CAMPUS WILL BE SOUGHT AFTER BY LOOKING AT EVERY POSSIBLE MISHAP AND PROVIDING REDUNDANT SAFETY MEASURES

ECONOMIC FEASIBILITY

1) AVAILABLITY OF FUNDS X E1: FUNDS HAVE ALREADY PROMISED FROM EE OFFICE E1: NONE REQUIRED

2) SUFFICENT BUDGET XE2: MATERIALS NEEDED FOR ROCKET BODY WILL BE RARE METALS AND VERY EXPENSIVE, BUDGET MAY BE TIGHT, MACHINING WILL HAVE TO BE DONE OUTSIDE RIT

E2: DETERMINE ALTERNATIVES FOR BODY MATERIALS AND LOOK UP SAMPLE PRICES AND MACHINING COSTS IMMEDIATELY

3) ADDITIONAL BUDGET X E3: IN ORDER TO EXPAND ON THIS PROJECT IN THE FUTURE ADDITIONAL FUNDS WILL BE REQUIRED

E3: DEMONSTRATE PROJECT FEASABILITY BY COMPLETING DESIGN OBJECTIVES

(c) E. Hensel and P. Stiebitz Page 1 Confidential - Not for Distribution

Risk Assesment

1 (LOW)

2 3 (MED)

4 5 (HIGH)

RATIONALE REMEDY/REQUIRED ACTION

SCHEDULE FEASIBILITY

1) PDR MILESTONE X S1: PROJECT IS CURRENTLY ON SCHEDULE S1: NONE REQUIRED

2) CDR MILESTONE X S2: PROJECT IS CURRENTLY ON SCHEDULE S2: NONE REQUIRED

3) FACULTY/MENTOR SCHEDULE X S3: MENTORS AND SPONSORS SECURED S3: NONE REQUIRED

4) SPONSOR AVAILABILITY X S4: SPONSORS ON CAMPUS DURING SEMESTER S4: NONE REQUIRED

6) TEAM MEMBER SCHEDULE X S5: LARGE TEAM WITH MANY DIFFERENT CONCENTRATIONS, CLASSES OFTEN OVER LAP

S5: ORGANIZE SCHEDULE AND AGREE TO MEETING TIMES FOR EVERY WEEK

7) TESTING OF HYBRID ENGINE BY MAY 2006 X S6: SCOPE OF PROJECT IS VERY AGGRESSIVE AND INNOVATIVE

S6: STICK TO PROJECT TIMELINES AND WORK HARD THROUGHOUT SEMESTER

TECHNOLOGY FEASIBILITY

1) NEW INVENTIONS REQUIRED X I1: STAGE SEPARATION AND HYBRID ENGINE SIZES MAY REQUIRE SOME INNOVATION

I1: THOROUGH SEARCH AND EXAMINATION OF AVAILABLE PARTS THAT PERFORM SIMILAR PURPOSES

2) TECHNOLOGY COMMERCIALLY AVAILABLE X I2: TO CUT BACK ON COSTS AND BURDENS TO TEAM WANT TO LOCATE AS MANY COMMERCIALLY AVAILABLE PARTS

I2: THOROUGH SEARCH AND EXAMINATION OF AVAILABLE PARTS

3) ARE TEAM MEMBERS EDUCATED IN TECHNOLOGY X I3: SOME TECHNOLOGY REQURED IS NOT COVERED IN TYPICAL ENGINEERING DISCIPLINES

I3: PROVIDE PAPERS AND MATERIALS FOR TEAM TO READ FOR BACKGROUND MATERIAL

4) IS TECHNOLOGY SIMPLE AND EASILY ASSEMBLED X I4: SOME PARTS ARE TECHNOLOGICALLY INNOVATIVE AND COMPLEX

I4: ANY SUBASSEMBLIES THAT REQUIRE TECHNOLOGICALLY ADVANCED TECHNIQUES WILL BE OUTSOURCED

5) INFORMATION AND DATA READILY AVAILABLE X I5: BECAUSE TECHNOLOGIES ARE NEW, INFORMATION ON ROCKET FUELS AND RATES ARE SCARCE

I5: DIFFERENT COMPANIES, TECHNICAL PAPERS AND UNIVERSITIES WILL BE SEARCHED AND CONTACTED FOR INFORMATION

ADDITIONAL

1) ADDITIONAL 1 (FUTURE USE)

2) ADDITIONAL 2 (FUTURE USE)

(c) E. Hensel and P. Stiebitz Page 2 Confidential - Not for Distribution

Appendix 5 – Objective Trees

Hybrid Rocket

Feed System Propellants Fuel Chamber

HTBP PMMA Activated Aluminum

See Appropriate Objective Chart

Properties

Solid Fuel Liquid Oxidizer

Regression Rate

Specific Impulse

Sizing

Insulation

Valve Selection

Residual Fuel

Flow Regulation

Gasification

Shower head oxidizer nozzle

Delivery

Hoses

Thermal Density

Flow Rate

Total Mass Flow Rate

Air Tight Attachment FEA & Temperature

Analysis

Fuel Pattern

Surface Area

Method to Cut Pattern

Nozzle Injector Plate

Oxidizer Tanks

Truss System

See Appropriate Objective Chart

L-Brackets and Straps

Ceramic Putty

Bolted Connection

Sold Fuel

Bolted Connection

Ceramic O-ring

Chamber Pressure

Survivability

Material

Throat and Exit Area

Sea Level (Testing)

Vacuum

Atmospheric Pressure

Materials

Test Stand

Location DAQ Construction

Software

Sensors

What data is needed

Operating Conditions (T&P)

Welding Accessibility

Order Parts

Budget Structural Analysis

Design

Motor Specs (Size & Thrust)

Local Authority Approval

Design Approval

Safety

Added Redundancies

Structural Integrity

Government Laws/ Mandates

Environmental Protection

Design

Motor Specs (Size & Thrust)

Combustion Reaction

Chemicals Used

Materials Used

Appendix 6 – Timeline SD II

Weeks 1-4 – Preliminary Component Tests Propellant Testing:

• Test different methods of molding • Test burning characteristics

Ignition Testing • Glow plug concept testing • Pyrotechnic ignition testing

o Ammonium Perchlorate mixtures o NC lacquer experience o Ni-Chrome voltage requirements

Data Acquisition Testing • Thermocouples • Pressure Transducers • Strain Gauges • LabView

Feed System / Oxidizer Flow • Pressure drops across injector • Pressure drops through feed system

Test Stand / Test Chamber • Verify Stress capabilities • Pressure and Force Tests

Weeks 5-7 – Testing Rocket Verify Safety and validity of Design

• After completing preliminary tests get approval for firing rocket from advisors

Test Rocket – Collect Data

Weeks 7-10 – Analyze Data Perform as many tests as possible

• Collect data for to complete all SD II objectives • Predict Environment of chamber • Accurate thrust readings

Analyze data and design tests for future teams • Based on collected data make recommendations

for future teams projects

Appendix 7 – Rocket Calculations

Givens and Assumptions: Isp = 235 sec di = 0.03 m

Vtotal = 9200 m/s 10=id

L for complete combustion

mL = 1 kg ρHTPB = 930 kg/m3

ms = 0.75 kg amax = 30g melec = 0.3 kg g = 9.81 m/s2

Ea = 20.557 kJ/mol A = 11.04 mm/s Ts = 1000 K R = 8.3143 J/(mol-K) M =0.1 kg/mol Mass:

gIv spe = ( ) ⎟⎠⎞

⎜⎝⎛= 281.9sec235

smve

smve 2300=

Nv

v total=Δ 4

9200sm

v =Δ smv 2300=Δ

elecsL mmmm ++=1 kgkgkgm 3.075.00.11 ++= kgm 05.21 =

⎟⎟⎠

⎞⎜⎜⎝

⎛=Δ

1

0lnmm

vv e ⎟⎟⎠

⎞⎜⎜⎝

⎛ Δ⋅=

evvmm exp10 ( )

⎟⎟⎟⎟

⎠

⎞

⎜⎜⎜⎜

⎝

⎛

⋅=

smsm

kgm2300

2300exp05.20

kgm 572.50 =

10 mmmp −= kgkgmp 05.2572.5 −= kgmp 522.3=

Regression Rate:

⎟⎟⎠

⎞⎜⎜⎝

⎛⋅

−⋅=

s

a

TRE

Ar exp ⎟⎟⎟⎟

⎠

⎞

⎜⎜⎜⎜

⎝

⎛

⋅⋅

−⋅=

KKmol

JmolkJ

smmr

10003143.8

557.20exp04.11

smmr 9315.0= However, calculation of regression is very inexact. The formula

used above is very general and based off experiments done by Chiaverini. Therefore, a

regression rate of 1 mm/s will be assumed for sizing calculations of the initial rocket engine. More exact calculations can be completed following testing.

Sizing:

idL ⋅= 10 ( )( )mL 03.010= mL 3.0=

After 1 second of burning:

( )( )trdd i 20 += ( sec11203.00 ⎟⎠⎞

⎜⎝⎛+=

smmmd ) md 032.00 =

( )LddV iHTPB22

04−=

π ( ) ( )( )( mmmVHTPB 3.003.0032.04

22 −= )π

3510922.2 mVHTPB−×=

After one second, VmHTPB ρ=& ( )353 10922.2930 m

mkgmHTPB

−×⎟⎠⎞

⎜⎝⎛=&

skgmHTPB 0272.0=&

Oxidizer-to-fuel ratio = 8:1 pHTPB mm ⎟⎠⎞

⎜⎝⎛=

91 ( )kgmHTPB 522.3

91⎟⎠⎞

⎜⎝⎛=

kgmHTPB 3913.0=

HTPB

HTPBb m

mt&

=

skgkgtb

0272.0

3913.0= sec4.14=bt

Back-Calculate to obtain an outer diameter for the grain size:

( )( )bi trdd 20 += ( )sec4.141203.00 ⎟⎠⎞

⎜⎝⎛+=

smmmd md 0588.00 =

Increase outer diameter to allow extra fuel to insulate the combustion chamber walls and to utilize standard pipe sizes: md 0635.00 =

Oxidizer Mass Flow Rate:

b

p

ox t

mm

⎟⎠⎞

⎜⎝⎛

= 98

& ( )

sec4.14

522.398 kg

mox

⎟⎠⎞

⎜⎝⎛

=& s

kgmox 2174.0=&

Appendix 8 – Regression Rates

t (seconds) Diameter (m) iterated A (m2) Go (kg/m2s) Axial Location, x (m)Instantaneous Regression

Rate, r (mm/s)First Regression, r (mm)

using long formula

Second regressionrate (mm/s) - using

second formula Regression (mm)1 0.03 0.028274334 7.381645558 0.02 0.155695447 0.155695447 0.929797194 0.9297971942 0.030155695 0.028421073 7.343533733 0.035 0.184034477 0.368068954 0.607279466 1.214558933 di = h = 0.03 m3 0.030368069 0.028621231 7.292178079 0.05 0.207088128 0.621264385 0.472186704 1.416560111 L / di = 104 0.030621264 0.028859862 7.23188187 0.065 0.227166813 0.908667253 0.394186276 1.576745102 L = 0.3 m5 0.030908667 0.029130733 7.164636538 0.08 0.245257967 1.226289834 0.342067272 1.710336359 m dot = 0.2348 kg/s6 0.03122629 0.029430085 7.091760434 0.095 0.261895937 1.571375619 0.304160279 1.824961676 t = 15 s7 0.031571376 0.029755321 7.0142451 0.11 0.277409906 1.941869342 0.275013083 1.925091579 p = 2.897895 MPa8 0.031941869 0.030104503 6.932886876 0.125 0.292020195 2.336161556 0.251702486 2.0136198849 0.032336162 0.030476114 6.848350456 0.14 0.305883014 2.752947129 0.232506073 2.092554661

10 0.032752947 0.030868925 6.761204293 0.155 0.319113925 3.191139254 0.216336332 2.163363323 Parameter11 0.033191139 0.031281912 6.671942323 0.17 0.331801241 3.649813649 0.202469281 2.227162096 C1

12 0.033649814 0.031714202 6.58099831 0.185 0.344014211 4.128170528 0.19040231 2.284827715 C2

13 0.034128171 0.032165043 6.488755867 0.2 0.355808283 4.625507684 0.179774387 2.337067033 n14 0.034625508 0.032633772 6.395555807 0.215 0.367228634 5.141200875 0.170318743 2.384462404 m15 0.035141201 0.033119802 6.301701741 0.23 0.378312612 5.674689187 0.161833491 2.427502371 k (m*MPa)-1

16 0.035674689 0.033622602 6.207464502 0.245 0.389091493 6.225463894 0.154162692 2.46660307917 0.036225464 0.034141695 6.11308574 0.26 0.399591755 6.793059843 0.147183743 2.50212362318 0.03679306 0.034676642 6.018780925 0.275 0.409836039 7.377048707 0.140798739 2.53437730419 0.037377049 0.035227038 5.924741905 0.29 0.419843875 7.977033628 0.134928424 2.56364004720 0.037977034 0.035792511 5.831139128 0.305 0.429632246 8.592644924 0.129507842 2.590156833

*using alternative formula to compare with first Regression *The two equations used include the first equation that was the long regression rate equation, and the do = 0.03 msecond one was the short regression rate equation that had activation energy included in it ρHTPB = 930 kg/m3

14.197

Givens

Correlation Parameter Table96% HTPB / 4% UFAL

0.0535

0.630.12257.11

Copied and pasted the values calculated in column F into column G until the values converged. In other words, Go was iterated until it converged.

Appendix 9 – Rocket Nozzle Calculations

Total Fuel Burned, mf = 3.5225 kg Target Burn Time, tb = 15 seconds Mass Flowrate, m = 0.235 kg/s & Assuming the following conditions: T0 = 3300 K P0 = 3.8 MPa γ = 1.20 Where subscript “t” denotes nozzle throat conditions:

1

0

1.201.20 16

112

1.20 13.8 10 12

2.145

t

t

t

P P

P Pa

P MPa

γγγ−−

−−

−⎛ ⎞= −⎜ ⎟⎝ ⎠

−⎛ ⎞= × +⎜ ⎟⎝ ⎠

=

01

112

13300 1.20 112

3000

t

t

t

T T

T K

T K

γ=−

+

=−

+

=

12

12

6

5 2

0.235 8314.4 3000

2.145 10 30 1.20

9.1194 10

tt

t gas

t

t

R TmAP M

kg J Ks kmolA kgPa

kmolA m

γ

−

⎛ ⎞⎛ ⎞ ∗= ⎜ ⎟⎜ ⎟ ⎜ ⎟∗⎝ ⎠ ⎝ ⎠

⎛ ⎞ ⎛ ∗⎜ ⎟ ⎜= ⎜ ⎟ ⎜×⎜ ⎟ ⎜ ∗⎝ ⎠ ⎝

= ×

&

⎞⎟⎟⎟⎠

Where Mgas is the molecular weight of the exiting gasses, which for these calculations was assumed to be 30, which represents a mixture of equal parts CO2, H2O, and N2.

The throat diameter is then: Dt = 0.010776 m Dt = 1.078 cm For the ground test nozzle, the ambient pressure, Pa was assumed to be 101.3 kPa. The mach number of the exiting gasses for the given conditions is then:

1

2 0

1.20 16 1.20

23

2

2 11

2 3.8 10 11.20 1 101.3 10

8.2962.88

ea

e

e

e

PMP

M

MM

γγ

γ

−

−

⎡ ⎤⎛ ⎞⎛ ⎞ ⎢ ⎥= −⎜ ⎟⎜ ⎟ ⎢ ⎥−⎝ ⎠ ⎝ ⎠⎢ ⎥⎣ ⎦⎡ ⎤⎛ ⎞×⎛ ⎞ ⎢ ⎥= −⎜ ⎟⎜ ⎟ ⎢ ⎥− ×⎝ ⎠ ⎝ ⎠⎢ ⎥⎣ ⎦

==

The cross-sectional area at the exit can be found in the following manner:

12( 1)2

1.20 12(1.20 1)2

5 2

4 2

112* 12

1.20 11 2.889.1194 10 2* 1.20 12.882

5.1959 10

et

ee

e

e

MAAM

mA

A m

γγγ

γ

+−

+−

−

−

−⎛ ⎞+⎜ ⎟⎛ ⎞= ⎜ ⎟⎜ ⎟ +⎝ ⎠ ⎜ ⎟

⎝ ⎠

−⎛ ⎞+⎜ ⎟⎛ ⎞×= ⎜ ⎟⎜ ⎟ +⎝ ⎠ ⎜ ⎟

⎝ ⎠= ×

Giving us an exit diameter of: De = 0.025721 De = 2.572 cm

Finally, the velocity of the exiting gasses can be found using the following equation:

11 2

0

0

121.20 1

3 1.20

6

2 11

8314.4 33002*1.20 101.3 1011.20 1 3.8 1030

2230.81

ee

gas

e

e

R T PVM P

J K PakmolV kg Pakmol

mV s

γγγ

γ

−

−

⎡ ⎤⎛ ⎞⎛ ⎞ ⎛ ⎞⎛ ⎞ ⋅⎢ ⎥⎜ ⎟= −⎜ ⎟ ⎜ ⎟⎜ ⎟⎢ ⎥⎜ ⎟⎜ ⎟−⎝ ⎠ ⎝ ⎠⎜ ⎟⎝ ⎠⎢ ⎥⎝ ⎠⎣ ⎦

⎡ ⎤⎛ ⎞⎛ ⎞⋅⎢ ⎥⎜ ⎟ ⎛ ⎞×⎛ ⎞ ⎜ ⎟= −⎢ ⎥⎜ ⎟ ⎜ ⎟⎜ ⎟ ⎜ ⎟− ×⎝ ⎠ ⎝ ⎠⎜ ⎟⎢ ⎥⎜ ⎟⎝ ⎠⎢ ⎥⎝ ⎠⎣ ⎦

=

Appendix 10 – Feed System Schematic

Balloon

# Qty Item Description 1 4 7 lb non-DOT bottle 2 2 Pressure Transducer 3 6 Brass Ball Valve 4 4 Brass Relief Valve (1000 psig) 5 1 Brass Relief Valve (750 psig) 6 13 Brass Pipe Nipple 7 1 Teflon Coated Braided Hose (1' Lg.) 8 1 Teflon Coated Braided Hose (1' Lg.) 9 1 Teflon Coated Braided Hose (6' Lg.)

10 1 Coaxial Solenoid Valve (Oxidizer) 11 1 Coaxial Solenoid Valve (Nitrogen) 12 1 Check Valve (Buna-N Seat) 13 1 Check Valve (Buna-N Seat) 14 8 Brass Tee Fitting 15 1 Brass Needle Valve 16 1 Brass Tee Fitting 17 3 Brass Pipe Nipple 18 2 Brass Hex Fitting 19 2 Brass Ball Valve 20 1 Brass Cross Fitting 21 1 Brass Elbow 22 10 Brass Hex Fitting 23 1 Pressure Gauge, Dual Scale (0-2000 psi) 24 1 Stainless Steel Pipe 25 7 Brass Reducer Coupling 26 1 Brass Rod (For Thread Adapters) 27 1 Brass Union 28 1 Brass Reducing Hex Fitting - 1 Pipe Thread Sealant Tape Roll - 2 Pressure Gauge, Dual Scale (0-1000 psi) - 1 Aluminum Rectangular Bar

19

181126 26

9

14 28

2

2 22 25

14 6

5

24

Combustion Chamber 16

Injector

25

6 3

6

12

8 15

6

10

15

6

7

2

6

22 3

6

22

14

21

22

25 4

N2O

2619

20

17

N2

Hybrid Rocket Feed System

Lexan Frame 1

23

16

17 17 18

4

26

1

Appendix 11 – Ni-Chrome Wire Temperature Properties

Current/Temperature Table - Ni Cr A & Ni Cr C Approximate Amperes to heat a Straight Oxidized wire to given temperature

Degrees F 400 600 800 1000 1200 1400 1600 1800 2000 Degrees C 205 315 427 538 649 760 871 982 1093

AWG INCH. /DIA. Amperes 8 0.128 22.40 32.00 41.00 52.00 65.00 79.00 95.00 111.00 128.00 9 0.114 18.80 26.80 34.50 44.00 55.00 67.00 80.00 94.00 108.00

10 0.102 16.20 23.30 29.70 37.50 46.00 56.00 68.00 80.00 92.00 11 0.091 13.80 19.20 24.80 31.50 39.00 48.00 57.00 67.00 78.00 12 0.081 11.60 16.10 20.80 26.50 33.50 40.80 48.00 56.00 65.00 13 0.072 9.80 13.60 17.60 22.50 28.20 34.20 41.00 48.00 55.00 14 0.064 8.40 11.60 15.00 18.80 23.50 29.00 34.60 40.50 46.00 15 0.057 7.20 10.00 12.80 16.10 20.00 24.50 29.40 34.30 39.20 16 0.051 6.40 8.70 10.90 13.70 17.00 20.90 25.10 29.40 33.60 17 0.045 5.50 7.50 9.50 11.70 14.50 17.60 21.10 24.60 28.10 18 0.04 4.80 6.50 8.20 10.10 12.20 14.80 17.70 20.70 23.70 19 0.036 4.30 5.80 7.20 8.70 10.60 12.70 15.20 17.80 20.50 20 0.032 3.80 5.10 6.30 7.60 9.10 11.00 13.00 15.20 17.50 21 0.0285 3.30 4.30 5.30 6.50 7.80 9.40 11.00 12.90 14.80 22 0.0253 2.90 3.70 4.50 5.60 6.80 8.20 9.60 11.00 12.50 23 0.0226 2.58 3.30 4.00 4.90 5.90 7.00 8.30 9.60 11.00 24 0.0201 2.21 2.90 3.40 4.20 5.10 6.00 7.10 8.20 9.40 25 0.0179 1.92 2.52 3.00 3.60 4.30 5.20 6.10 7.10 8.00 26 0.0159 1.67 2.14 2.60 3.20 3.80 4.50 5.30 61.00 6.90 27 0.0142 1.44 1.84 2.25 2.73 3.30 3.90 4.60 5.30 6.00 28 0.0126 1.24 1.61 1.95 2.38 2.85 3.40 3.90 4.50 5.10 29 0.0113 1.08 1.41 1.73 2.10 2.51 2.95 3.40 3.90 4.40 30 0.01 0.92 1.19 1.47 1.78 2.14 2.52 2.90 3.30 3.70 31 0.0089 0.77 1.03 1.28 1.54 1.84 2.17 2.52 2.85 3.20 32 0.008 0.68 0.90 1.13 1.36 1.62 1.89 2.18 2.46 2.76 33 0.0071 0.59 0.79 0.97 1.17 1.40 1.62 1.86 2.12 2.35 34 0.0063 0.50 0.68 0.83 1.00 1.20 1.41 1.60 1.80 1.99 35 0.0056 0.43 0.57 0.72 0.87 1.03 1.21 1.38 1.54 1.71 36 0.005 0.38 0.52 0.63 0.77 0.89 1.04 1.19 1.33 1.48 37 0.0045 0.35 0.46 0.57 0.68 0.78 0.90 1.03 1.16 1.29 38 0.004 0.30 0.41 0.50 0.59 0.68 0.78 0.88 0.98 1.09 39 0.0035 0.27 0.36 0.42 0.49 0.58 0.66 0.75 0.84 0.92 40 0.0031 0.24 0.31 0.36 0.43 0.50 0.57 0.64 0.72 0.79

Appendix 12 – Safety Report

Safety Assessment Report For

Project METEOR Senior Design Team #06006

Team Members:

David Dale – ME (Project Manager) John Chambers – ME Brad Addona – ME

Jessica LaFond – ME Chris Hibbard – ME

Anthony Fanitzi – ME Jeff Nielsen – ME Dan Craig – ME

Faculty Mentors

Dr. Dorin Patru

Dr. Jeffrey Kozak

Faculty Advisor

Dr. Alan Nye

1.) Introduction: The purpose of this safety report is to outline and document all potential safety concerns and how they will be solved to statically test a hybrid rocket engine. This test will complete one main objective of the design team’s senior project. 2.) System Description 2.1 Purpose and Intended Use

The purpose of this system is to have a location and test setup that may be used multiple times by different senior design groups to test multiple rocket engines. It is expected that the test stand will be used less than 10 times per year.

2.2 System Development Through research, brainstorming, Pugh analysis and discussions with

professors the members of this senior design team (hereafter referred to as Team) believe that the current test setup is the best for our purposes based on ease of safety, construction, cost and data acquisition.

2.3 System components Refer to Appendix A: “System Components” 2.4 Functional Diagrams / Sketches / Schematics Refer to Appendix B: “System Diagrams” 3.) System Operations 3.1 Operating, Testing and Maintaining Procedures

Refer to Appendix C: “Operating, Testing and Maintaining Procedures” and Attachment 1: “Check-off sheet”

3.2 Special Safety Procedures

All safety issues will be controlled with the measures outlined in section 3.1. In addition there will be two fire-extinguishers (1 H20, 1 C02) on hand. Also, one of our team members is a volunteer Fireman and his expertise will be used if necessary. Future teams will have a volunteer fireman onsite during every testing.

3.3 Operating Environments The expected operating environment is a location with a large open area

(50’ x 100’) so it can be assured no bystander will be harmed in anyway. Ideal weather conditions are low wind, during daylight hours and no precipitation.

3.4 Facility Requirements or Support Equipment It is desired to have a concrete pad (minimum 6’x 6’ x 4”) to mount the test stand to. A van or pickup truck may be needed to transport the test

stand, rocket engine, fuels, instrumentation, safety measures (shields, deflectors, fire extinguisher), and needed computers to the test site.

4.) Systems Safety Engineering 4.1 Ranking Hazardous Conditions Description Category Environmental, Safety, Health Results

Catastrophic I Could result in death, permanent total disability, loss exceeding $1M, or irreversible severe environmental damage that violates law or regulation.

Critical II

Could result in permanent partial disability, injuries or occupational illness that may result in hospitalization of at least three personnel, loss exceeding $200K but less than $1M, or reversible environmental damage causing a violation of law or regulation

Marginal III

Could result in injury or occupational illness resulting in one or more lost work days, loss exceeding $10K but less than $200K, or mitigatible environmental damage without violation of law or regulation where restoration activities can be accomplished.

Negligible IV Could result in injury or illness not resulting in a lost work day, loss exceeding $2K but less than $10K, or minimal environmental damage not violating law or regulation

4.2 Ranking Hazard probabilities Description Level Occurrence

Frequent A Continuously experienced Probable B Occurs frequently

Occasional C Will occur several times Remote D Unlikely, but can be expected to occur

Improbable E Unlikely to occur, but possible 4.3 Identifying Hazardous Conditions 4.3.1 List of all hazards Refer to Appendix D: “Hazards”

4.4 Hazardous Materials Refer to Appendix E: “Hazardous Material”

5.) Conclusions and Recommendations 5.1 Results

The conclusion of this safety assessment is that there are risks involved with the testing of a hybrid rocket, but the risks can be mitigated through the initial and redundant safety measures taken by the project design team. The risks can be reduced enough that we can confidently test a hybrid rocket without putting the members of our team, the environment or a third party in danger.

5.2 Stress Analysis Refer to Appendix F: “Stress Analysis of Critical Components”

APPENDIX A

System Components

1. Concrete Bed

2. Anchor Bolts

3. Base

4. Base Bolts

5. Test Beam

6. Engine Cradle

7. Engine Cradle Bolts

8. Rocket Engine Assembly

9. Holding Straps

10. Oxidizer delivery system

11. Sensors / Strain Gages

12. DAQ Computer

13. Lexan shields

14. Angle Stock / Lexan Frame

15. Exhaust deflector

16. 10’ Hose

17. Ignition System

18. Nitrous Oxide Tank

19. Cage for Nitrous Oxide and N2 Tanks

20. 2 Fire extinguishers (1 CO2, 1 H20

21. Video Camera

22. Generator

23. Power Supply

24. Ear Plugs

25. Safety glasses

26. First Aid Kit

27. Tools (socket set, wrenches)

APPENDIX B

System Diagrams

APPENDIX C

Operating, Testing and Maintaining Procedures

Setup Procedure

Install Test Beam

1. Put base plate in place and tighten down with concrete anchor bolts to a minimum

of Treq = 6 lb*in

2. Install test beam

3. Install bolts in vertical part of base plate and tighten to a minimum of

Treq = 139.5 lb*in

4. Affix engine cradle to top of beam, install and tighten bolts to a minimum of

Treq = 30 lb*in.

Secure Test Chamber

5. Put engine assembly in engine cradle

6. Put 4 straps around rocket and cradle, tighten down to minimum Treq = 5.96 lb*in.

i. Visually inspect to ensure engine assembly is in contact when flat edge

of engine cradle

Connect Data Acquisition / Test Electronics

7. Connect thermocouples, pressure sensors and strain gauges to rocket, run wires to

DAQ computer

8. Test / Calibrate all gauges

9. Turn on and off inline valves for each tank to ensure successful operation

10. Connect 10’ Hose to the Injector on the Rocket Engine Assembly

Secure Lexan Panels

11. Put pre-assembled lexan box into place and bolt to the concrete pad

12. Set up thrust deflector approximately 18” from edge of lexan

Clear Area

13. Walk perimeter and make sure there is no one in the area that is not with the team

14. Remove any flammable debris that may be around test location

15. Connect 10’ hose to cage containing the tanks

16. Connect ignition system to power supply

17. Team members will then be positioned behind barriers to ensure their safety and the test will begin. Setup Checklist

Setup # Action Completed Initial

Concrete pad inspected for cracks or crumbling 1 Base plate welds inspected for cracks or fatigue 1 All Anchor Bolts installed 1 All Anchor Bolts tightened to at or above required torque 3 All base plate beam bolts installed 3 All base plate beam bolts tightened at or above required torque 3 Test beam pulled with 100 lbs of force and it did not come loose 4 Enging Cradle welds inspected for cracks or fatigue 4 Engine Cradle bolts installed 4 Engine Cradle bolts tightened at or above required torque 4 Test beam with 3 times expected force and it did not come loose Rocket engine body checked for cracks or fatigue 6 4 straps put around the rocket body 6 4 straps tightened around the rocket body 6 Ensure Engine is in contact with flat on Engine Cradle Visually inspect lexan panels for cracks or holes 7 Install sensors 8 Test sensors to make sure they work and are calibrated 9 Test electronic control valves on all tanks 10 Connect interior hose 11 Lexan Box bolted into place 11 Physically make sure panels are secured 12 Set up thrust deflector 18" away from edge of lexan

13/14 Clear perimeter of any bystanders / debris

Test Procedure

Countdown

t = 0 Apply current to igniter, start DAQ

t = 0.5 turn on NOx

t = 1 Turn off current to igniter

1 < t < test time Pure NOx

t = test time Electronically shut off NOx, turn on N2

t = test time + 3 Electronically shut off N2, stop DAQ

1. Visually inspect rocket before approaching to ensure flame has extinguished

2. Manually shut off redundant valves on tanks

3. Remove outer hoses

4. Safely remove lexan box and Engine Assembly

APPENDIX D Hazards

Hazard / Mishap Category Level Control Residual

RisksConcrete could crack and break

III E Will inspect concrete before use and use an adequate pad.

None

Anchor bolts may not hold in concrete

III D There will be multiple, redundant, bolts in the concrete, more than one would have to fail for this to be a risk

None

Not tighten down bolts enough and post releases

III D This will be checked in our pre-test inspection

None

Welds break at top or bottom

III D Welds will be checked before each test, if welds break the rocket will still be held down by the tethers

None

Nox tanks explode

I D Tanks will be separated from observers by bulletproof material

Shrapnel

Straps break or are not tightened properly

III C This will be checked in our pre-test inspection, also there will be multiple straps

None

Post breaks III E Post will be inspected before test, if a failure does occur the rocket will still be held down by the tethers

Damage to Lexan or Electronics

Rocket explodes due to excess pressure

I E There will be a pressure relief valve on the rocket engine

Shrapnel

Exhaust ignites surroundings

II D There will be an exhaust deflector and we will be cognizant of what is behind the test stand, making sure there is nothing flammable within a reasonable distance

Small fire

Injector plate separates from rocket body

II D The straps will hold the tank, the plate on the front of the engine cradle will stop the injector plate

None

Hazard / Mishap Category Level Control Residual

RisksChunks of fuel are spit out from rocket

III C If anything solid is spit out from the nozzle, it will first hit the deflector, breaking it up and then fall harmlessly to the open area behind the rocket

Debris contacts observer or small fire

Nozzle disengages from rocket during firing

II D The nozzle will be moving away from any observer or bystander and will fall harmlessly on the ground

Debris fly forward and contact observer

Valve for any tank is stuck open

IV D No harm in this, except we cannot stop the test, we have to wait for the rocket to burn itself out

None

bolts break in top or bottom clamp

III E There are multiple bolts in top and bottom, so multiple failures would have to occur if this was to become a problem. However the tethers would still hold the rocket

None

Heat causes straps to loosen

III D Straps self tighten with temperature variation

None

Hoses burst or become unattached

III D If the hoses become unattached the rocket will simply stop, the burst will be contained within the bulletproof panels and shrapnel blanket

None

Foreign object enters test stand

IV C The test stand will be enclosed on 5 of 6 sides, if an object enters the test stand it could harm the object (if living) but will not be detrimental to the test or any observers

None

Shields could break or shatter

I E The shields have been tested with bullets, if they did break they would hopefully slow down any projectile enough that it would fall harmlessly to the ground a few feet from the test cell

If the shields fail there is nothing between the rocket and pressurized tanks but there is still an additional barrier between observers

APPENDIX E

Hazardous Materials 1. Nitrous Oxide, N2O 2. Nitrogen, N2 3. Hydroxyl Terminated Poly-Butadiene (HTPB – tire rubber) 4. Pyrogen / NC Laquer (part of ignition system) 5. Ammonium Perchlorate (part of ignition system) 6. Lexan (Polycarbonate)

APPENDIX F Stress Analysis of Critical Components

Beam Calculations The Test Beam has been modeled as a solid cantilevered beam, though it has holes in it. This is a valid analysis due to the fact that the parts of the beam that have holes in it are pre-loaded with a force that will ensure the clamp will not spread apart under expected loading conditions.

The beam has been sized for a factor of safety of 2. Bolt Analysis To find the required pre-load of the bolts in the bottom of the test beam the geometric center of the bolt pattern was found. From this the maximum expected force in each bolt was determined. This was done in two steps because the bolts will take up the shear force and the couple created by the applied force. To size the bolts the expected force that the bolts will see was set equal to the pre-load, again to prevent separation of the plates used for clamping. A factor of safety of 2 was used for the calculations. From the pre-load the proof load can be found which leads to the proof strength. Bolts are then selected from Table 8-9 in the Mechanical Engineering Design text book (Shigley, 7th ed.). The simple calculation for bolt torque is done to find out how much torque should be applied to the bolt during the setup process. Concrete Anchor Bolts The anchor bolts hold the whole test stand to the ground. We must be sure they hold. The concrete will fail long before the actual bolts will fail (~120,000psi for bolts, ~2900psi for concrete), so calculations must be done to determine how much force the concrete can handle. As long as the anchor bolts are embedded at least 2.375” into the concrete the bolts will have a factor of safety over 40. The torque equation was again used to find out how much torque should be applied to the bolts during the setup process.

Strap Analysis Ideally the thrust of the rocket will be perfectly horizontal. Because of the front plate, a moment will be created and the back end of the rocket will tend to move in the upward direction. This is counteracted by the straps that will hold the rocket in the test stand. To plan for the worst case scenario we have done calculations that include only one strap, located at the farthest point from the moment (longest possible moment arm, greatest force) and a thrust that is angled 10° away from horizontal. Even in this worst possible scenario the strap only needs to withstand 39lb. The torque on the bolt was calculated with a factor of safety of 2. The bands themselves have a factor of safety of over 8. Pressurized Cylinder (Combustion Chamber)

The tangential stress was found to be the maximum stress experienced by the cylinder. A concentration factor for a pressurized cylinder with one hole in it was used for our calculations. Since there are two holes in our cylinder, we combined them into one larger hole, this is an acceptable assumption. Using Peterson’s Stress Concentration Factors a factor of safety of 4 was found for the cylinder, using 304 Stainless Steel. It is also important to note that if the pressure inside our cylinder did for some reason increase to above 1000 psi, that all flow of Oxidizer would halt and this would in turn prevent any possibility of our tank bursting. Injector Plate Bolts These bolts were sized in the exact same manor as the bolts in the previous paragraph “Bolt Analysis”. Using a factor of safety of 2 the pre-load was set to the expected force, a proof load and proof strength was found from this value and bolts were chosen from Table 8-9 in the Mechanical Engineering Design textbook (Shigley, 7th ed.). The required torque on the bolts was also calculated. Lexan Analysis Attached at the end of Appendix F is an experiment that was completed by a group of students supervised by Dr. Wellin determining the impact strength of Lexan. As witness from their results the Lexan is able to withstand significant impacts from bullets sized pieces of shrapnel.

Appendix 13 – MSDS Sheets

Ammonium Perchlorate - MSDS.pdf

BBKWQ-NClacquer.pdf

MSDS N2O.pdf

MSDS_GP.pdf

N2O hazmat.pdf

Nitrogen MSDS.pdf

Nitrous Oxide MSDS.pdf

http://meteor.rit.edu/docs/Ammonium%20Perchlorate%20-%20MSDS.pdf

http://meteor.rit.edu/docs/BBKWQ-NClacquer.pdf

http://meteor.rit.edu/docs/MSDS%20N2O.pdf

http://meteor.rit.edu/docs/MSDS_GP.pdf

http://meteor.rit.edu/docs/N2O%20hazmat.pdf

http://meteor.rit.edu/docs/Nitrogen%20MSDS.pdf

http://meteor.rit.edu/docs/Nitrous%20Oxide%20MSDS.pdf

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