The Development of a Paraffin Based Experimental Hybrid Sounding Rocket UCLA

The UCLA Rocket Project

1

The Development of a Paraffin Based Experimental Hybrid

Sounding Rocket

Kurt U. Zimmerman1, Brian C. Kentosh

2, Peter C. Chang

3, Mackenzie J. Booth

4, Richard E. Abrantes

5, Phuoc Hai N. Tran

6,

Craig P. McGrath6, Brandon A. Dizon

7, Adel M. Shalabi

7, William A. Silva

7, Matthew J. Marshall

7

University of California, Los Angeles, CA, 90095

The UCLA Rocket Project entered the third year of design and testing of its custom hybrid rocket engine

which will be used as the propulsion system for UCLA’s entry into the 6th

Intercollegiate Rocket Engineering

Competition put on by the Experimental Sounding Rocket Association of Green River, Utah. The hybrid

engine combines paraffin wax with 5 micron aluminum powder as its fuel and liquid nitrous oxide as its

oxidizer. It was designed by student members of the UCLA Rocket Project and has a specific impulse of 202

seconds, a thrust of 1004 pound force, and a burn time of 9.9 seconds. In conjunction with the 15 foot long

carbon fiber and fiberglass air frame designed and fabricated by the Rocket Project, the hybrid engine is

capable of propelling a 10 pound payload to 25,000 feet above ground level. An on board avionics package

comprised of off the shelf accelerometers and microcontrollers is capable of accurately predicting the rocket’s

apogee and initiating an engine shut down in order to reach exactly 25,000 feet. Students in the Rocket

Project also designed, developed and constructed all of the electrical interfaces for the various sensors and

electronic valves used in both testing and launch of the experimental rocket. For recovery and reusability, the

rocket employs a dual deployment drogue and main parachute system, which allows the rocket to land safely

within 6 minutes of reaching apogee.

Nomenclature

a = regression rate coefficient l = port length t = time

ak = acceleration L = nose cone length T0 = stagnation temperature

Acs = cross section area m = mass Tn = temperature bulk air

Ainj = area injector orifice ṁ = mass flow v = velocity

CD = coefficient of drag M = mach number x = distance from tip

CP = specific heat constant pressure n = regression rate exponent Λ = shock angle

c* = characteristic velocity P = pressure ρ = density

FD

= drag force r = radius

g = gravitational constant ṙ = change in radius

k = flow loss coefficient R = base radius

I. Introduction

HE UCLA Rocket Project has continued its three year project to design and fabricate its custom experimental hybrid

rocket engine, known as the Hybrid Propulsion Experiment (HyPE). The goal of this development has been to compete

in the Experimental Sounding Rocket Association’s (ESRA) Intercollegiate Rocket Engineering Competition (IREC). The

objective of the advanced competition in the 6th

IREC includes carrying a 10 pound payload to 25,000 feet above ground

level (AGL) and returning the rocket and payload in reusable condition sans expendables such as propellant. The UCLA

Rocket Project started working on the HyPE during the 2008-2009 year with six active members participating in the project.

1 President, UCLA Rocket Project

2 Propulsion Lead, UCLA Rocket Project

3 Electronics Lead, UCLA Rocket Project

4 Aerodynamics and Structures Lead, UCLA Rocket Project

5 Recovery Manager, UCLA Rocket Project

6 Propulsion Team Member, UCLA Rocket Project

7 Electronics Team Member, UCLA Rocket Project

T


2

Figure 2. Carbon fiber

tube lay-up process

Figure 1. Quarter Cutaway View of

UCLA’s Entrant into the 6th

IREC

Figure 3. Coefficient of Drag vs Mach Number Plot Generated

using RASAero5

Today there are over 30 active members with diverse backgrounds from aerospace and mechanical engineers, to electrical

engineers, material scientists, and computer scientists.

In addition to completion of the design, fabrication and testing of the HyPE engine, the UCLA Rocket Project also

designed and fabricated its own carbon fiber and fiberglass airframe, designed the dual deployment recovery system, all of

the electronic controls and interfaces, the test equipment and custom avionics package capable of predicting the apogee of the

rocket in flight for accurate and dynamic engine shut down capability. All of these projects were undertaken with the intent of

competing in and winning the advanced category of the 6th

annual IREC competition.

II. Structural and Aerodynamic Systems

A. Airframe

The primary purpose of the airframe is to provide a rigid structure

through which the thrust can be transferred, to house the various

subsystems while inducing minimal drag, and to allow easy access to

components within the airframe. Accessibility of internal components is

important for integration in the field, so the airframe diameter was

increased to 8 in. The increased diameter is also beneficial in shortening

the overall rocket length to 15ft, giving an aspect ratio of 22.5, which

eliminates the extreme bending moment seen in UCLA’s entrant to the 5th

IREC, the RATTworks rocket.

There were three key requirements that were considered in choosing a

material for the airframe: strength, weight, and cost. In order to achieve

both high strength and low weight, composite materials were deemed the

best option. After conducting extensive

research on available composites, a

unidirectional pre-impregnated (pre-preg) carbon fiber was chosen for the coupling tubes,

and a woven pre-preg was chosen for the body tubes. Not only do these materials have a

high strength to weight ratio, but they are relatively easy to work with. In addition, their

widespread use in the aerospace and marine industries made it possible to eliminate material

costs through in-kind material donations.

It is necessary to cure pre-preg composite parts in either an oven or autoclave. The

UCLA Rocket Project’s oven size limited the lengths of tubes to 23 in. This necessitated the

use of coupling tubes, made of 8 layers, to link together outer body tubes, made of 3 layers.

The process of fabricating the carbon fiber tubes for the airframe was relatively simple,

but tedious. Prior to each lay-up, the carbon fiber was cut to size, and an aluminum mandrel

was wrapped in a release film to prevent the cured carbon fiber tube from adhering to the

mandrel. Layers of carbon fiber were then wrapped around the mandrel individually. A

layer of pressure tape was applied after every other layer and then subsequently removed

after ten minutes. The pressure tape provided compression on the fibers and prevented the

formation of creases in the final part by reducing the bulk factor of the material1. After the

final layer of carbon fiber was applied, a layer of

porous peel ply was added. This facilitated the

bleeding of excess resin out of the tube, which

decreased weight and increased strength. A final

layer of pressure tape was added, and then the

tube was allowed to cure within the oven for 2.25

hours at 275 degrees ºF.

B. Drag Analysis

Extensive research was conducted on

various types of drag to optimize the performance

of the rocket. Three primary forms of drag had to

be taken into account while designing the rocket:

skin friction drag, pressure drag and base drag2,3,4

.

Skin friction drag is affected by the wetted area

and surface smoothness2, with the wetted area


3

Figure 4. Drag comparison for various nose

cone shapes at a fineness ratio of 3:12

being a design parameter that is determined by the nose cone shape and the fineness ratio of the nose cone. Pressure drag is

based upon the cross sectional area of the body and is relatively small at subsonic speeds. One of the specific types of

pressure drag is wave drag, which begins to affect objects moving at velocities greater than Mach 0.8. Base drag is caused by

the low pressure region created at the aft end of the rocket, and can have a substantial contribution to overall drag once the

engine has stopped thrusting2. In addition to these main three forms of drag, the rocket also experiences interference drag

around the camera fairing and where the fins are attached.

To determine the overall drag of the rocket, shown in Figure 3, an excel calculator was created, which took in

various geometric inputs and then plotted the different types of drag. This made it possible to quantify which types of drag

had the greatest effect. This calculator was used in conjunction with the free software RASAero5.

C. Nose cone

After studying the different nose cone shapes available, an LD Haack

shape was selected, as research showed that it outperformed most other

shapes in the transonic region, shown in Figure 42,4,6

. The LD Haack isn’t

a simple geometric form; rather it is a mathematically derived shape that

has been optimized to reduce wave drag. The equation for an LD Haack

is:

sin 2

2

Rr

(1)

Where

12

cos 1x

L

(2)

Where R is the base radius, L is the nose cone length and x is the distance

from the tip of the nose cone.

The optimum length ultimately came to be a compromise between skin friction drag and wave drag, as increasing the

length increased the skin friction drag, yet reduced the wave drag by reducing the curvature. Data from the both the in house

excel calculator and RASAero led to a fineness ratio of 4.5:1.

Aerodynamic heating was taken into consideration due to the rocket attaining supersonic velocities. The maximum

temperature that a fluid can be heated to near a moving body is the stagnation temperature2,3,7

, where Tn is the temperature of

the bulk air, v is the rocket’s velocity (1340 ft/s) and CP is the specific heat at constant pressure:

2

02

n

p

vT T

C

(3)

Equation (3) gave a maximum stagnation temperature of 144 ºF (335K). To alleviate problems with aerodynamic heating the

nose cone tip was slightly rounded to increase the surface area through which heat is absorbed.

Due to the presence of radio frequency transmitters in the nose cone, it was necessary to choose a material that did not

block radio waves. Furthermore, this material should be capable of withstanding temperatures up to 144 ºF. Fiberglass with a

room temperature curing resin fulfilled these requirements. Many of these resins have relatively low glass transition

temperatures, Tg. However, after contacting several companies in the resin industry, a suitable mixture was found in 820

green laminating system resin from Adtech. In order to manufacture the nose cone, a male plug followed by a female

fiberglass mold had to be made before the actual nose cone could be fabricated. The plug, made of wood, polystyrene foam

and Bondo, was used to provide a shape around which the mold could be made. The mold was made into two halves for ease

of lay-up. The majority of the nose cone was fiberglass, however some carbon fiber was added in the tip and around the

shoulder to alleviate aerodynamic heating and provide additional strength.

D. Boattail

Boattails reduce base drag by reducing the cross sectional area at the base of the rocket, and thereby minimizing the low

pressure region. Boattails may also incur extra wave drag if curvature isn’t kept to a minimum2,8

. In theory, the absolute

minimum wave drag for a boattail is provided by a shape that has zero change in cross sectional area at the base; however for

ease of fabrication, a simple LD Haack was used9. This shape has better wave drag characteristics than a conical boattail,

however, and enabled it to be fabricated from the same mold that was used to make the nose cone.

E. Fins

Fins are an important element for ensuring that the rocket remains stable during flight by moving the center of pressure

behind the center of gravity. A clipped delta shape was chosen to maximize stability. The clipped delta is a shape that is easy


4

Figure 5. Dimensioned Fin

Figure 7. Stress Simulation Showing a

Factor of Safety Plot

Figure 8. Assembled HyPE 1B Engine.

Figure 6. Typical Supersonic

Airfoil Configurations2

to manufacture, provides a good amount of rigidity to counter fin flutter, and has relatively low drag compared to several

other designs2.

The primary design factor for the fins was to provide a stability margin of 1.5 to 2.0 calipers, or reference diameters, for

both the wet and dry configurations of the rocket. The Barrowman Equations and the open source software OpenRocket10

were used to determine the fin dimensions and initial estimate of the rocket’s center of pressure location relative to its center

of gravity. The RASAero software was utilized to determine a more comprehensive center of pressure location as a function

of Mach number11

. With the aid of the above software, the optimum fin dimensions, shown in Figure 5, were determined to

be 12 inches for the root chord, 4 inches for the tip chord and 8 inches for the height. In order to ensure that the fins stay

within the mach cone, equation (4) was used, where M is the Mach number and Λ is the shock angle2:

1 1

90 sinM

(4)

With a maximum mach number of 1.25, a sweep angle of 36.9º is determined. For the

dry configuration, a stability margin of 2.31 calipers was achieved, while for the wet

configuration, the weight increases to 180.16 lbs with a stability margin of 1.86.

The second primary design criterion was to minimize drag and fin flutter. Fin flutter

is a major concern as it can be severe enough that enough stress can be put on the fins

that the material fails causing the fins to be violently ripped off of the rocket body.

Additionally fin flutter also creates a turbulent flow and increase drag. Several of the

factors that affect fin flutter are the dynamic pressure, Mach number, material stiffness

and fin mass12,13

. To prevent fin flutter, the fins must be prevented from bending or

twisting. Smooth flow around the fin will act to

dampen out the oscillations12,13

.

Four aluminum brackets were made for each fin, in order to keep them rigidly

attached. The fins were then bolted to the engine mount tube and bonded to both the

engine mount and the boattail. Fiberglass was added to the bonds to provide strength.

The fins were made of 20 layers of unidirectional pre-preg, alternating the layer’s

fiber direction by 90º for increased strength. To reduce drag, the fins used a modified

double wedge cross section (Figure 6), which is optimum for supersonic flights2. The

tip of the fin was left flat to prevent flow around the tip.

F. Launch Rail and Test Stand

A test stand was built in order to conduct static hot fire tests of the

experimental engine. Based on the projected weight of the rocket, it became

obvious that a reinforced launch rail would be necessary. A modular setup

was designed and built to accommodate both requirements and reduce

material cost.

After several design studies, a triangular cross section was chosen, where

two side by side 6ft long sections would make up the test stand setup and

four 6ft sections stacked end to end would become the launch rail. The

modularity of the system not only decreased material costs, but it also

improved transportability. The test stand version has a factor of safety

(FOS) of 3.88 based on a 1000 lb load (Figure 7) and in the launch rail configuration the system deflects 3.31 in (84mm)

based on a 100lb load from off axis thrust applied perpendicularly to the top of the rail, the worst case scenario. The system

was designed in SolidWorks and has shown an ability to handle the expected loads. The final launch rail/test stand was

constructed from ASTM A36 steel and assembled using a combination of welds and bolts.

III. Propulsion Systems

One of the primary focuses

of this project was the Hybrid

Propulsion Experiment 1B

(HyPE 1B) propulsion system.

The overall layout is shown in

Figure 8.

The HyPE 1B is a student-

designed hybrid rocket engine, using aluminized paraffin as the solid fuel and liquid nitrous oxide as the oxidizer. Regulated


5

Figure 9. Fuel grain port detail. The

minor rippling is believed to result from

slight misalignment during spin casting

operations.

nitrogen pressurizes the oxidizer tank, flowing liquid nitrous oxide into the combustion chamber where it combusts with the

paraffin. D-class hobby rocket motors serve as the ignition source. A graphite bell nozzle accelerates the hot gases and

produces thrust. Loads are carried from the combustion chamber to the oxidizer tank through a thrust structure and

subsequently transferred to the rocket body. All tanking is performed remotely, and in-flight engine cutoff is made possible

by an on-board pneumatic system. Table 1 provides a summary of the engine’s predicted performance characteristics.

Table 1. HyPE 1B Performance Summary

Thrust [lbf] Isp [s] C* [ft/s] O/F Ratio Burn Time [s] Pc [psi] Ae/At

1004 202 4,619 4.30 9.90 400 5.21

A. Previous Research

The previous year, the UCLA Rocket Project developed and test-fired the first HyPE engine, designated the HyPE 1A, to

represent that this was the Block I design of the HyPE engine. The aluminum combustion chamber of the HyPE 1A deformed

during its first and only hot fire test, which occurred after the 5th

IREC in Green River, Utah. To mitigate this, the combustion

chamber was converted from ¼ in thick aluminum to ½ in thick steel. This variant was designated the HyPE 1AT (HyPE 1,

Block I, Test Configuration). The HyPE 1AT allowed for research to continue while the upgraded HyPE 1B (Block II) was

being designed.

B. Propellant

The HyPE 1B fuel grain was based on research performed by Stanford

University, which showed that aluminized paraffin had a high regression

rate, allowing for high thrust via a single port design14-30

. This, coupled with

moderately high performance, made paraffin an attractive option.

The final composition of the HyPE 1B fuel was 74% paraffin wax, 18%

aluminum powder, 6% Vybar 103, and 2% carbon black. The addition of

aluminum increases the regression rate and Isp of the engine14

. Vybar and

carbon black were used to reduce sloughing, which is the expulsion of

unburned fuel out the nozzle. Vybar, commonly used as an additive in

candles, increases the melting temperature and helps slow the melting of the

fuel. Carbon black acts as an opacifier and absorbs thermal radiation,

preventing the heat from conducting past the surface layer.

To cast the fuel grain, the components were melted together in an electric

convection oven. The mixture was then poured into an ablative liner and

spun on a lathe until it cooled. This produces a grain with an even port down the central axis. Weights were taken before and

after casting to verify the mass of the fuel grain.

Characteristic velocity, c*, was determined through NASA Chemical Equilibrium with Applications (CEA). CEA did not

include data for paraffin, so this was input from data given by the Journal of Chemistry31

. From this analysis the optimal O/F

ratio was found to be 4.30, and this value was used to size the fuel grain for optimal combustion.

C. Pressurant System

The HyPE 1B employs a pressure regulated system to keep the oxidizer flow constant for the burn duration. This

performs better than the simpler blow down system where the pressurant gas is unregulated, causing oxidizer pressure to drop

as the gas expands.

The HyPE 1B utilizes a L45M SCBA tank manufactured by Luxfer Gas Cylinders. Specifications of this tank are given in

Table 232

.

Table 2: Specifications of Luxfer L45M32

Service Pressure Volume Diameter Length Weight

4500 psi 285 in3 5.4 in 18.4 in 6.6 lb

The L45M was sized for the HyPE 1A and as such is slightly undersized for the HyPE 1B, however the performance gain

of upgrading the pressurant was not substantial enough to justify the cost. The pressurant tank is filled with gaseous nitrogen

from a standard K-size gas cylinder. The nominal fill pressure is 2250 psia, which is the standard pressure of the nitrogen gas

cylinder.

The pressurant flow is controlled by a Mighty Mite high flow regulator manufactured and generously sponsored by

Dresser. A more stable oxidizer pressure is achieved due to its high flow rate.

D. Oxidizer Tank


6

Figure 11. Ox Valve Actuation. Valves

are mounted to thrust rods (not

pictured for clarity)

Figure 10. HyPE 1B Oxidizer Tank.

Left hand side connects to pressurant

tank, right hand side connects to

combustion chamber

Figure 13. HyPE 1B Combustion Chamber

Figure 12. Impinging injector

during a water flow test in a

demonstration of atomization.

The oxidizer tank used in the HyPE 1B was designed in-house and

holds 36 lbs of liquid nitrous oxide. The entire tank was constructed from

6061-T6 aluminum alloy for its affordability and availability. The layout of

the oxidizer tank is shown in Figure 10. The outside diameter, excluding

bulkhead flanges, is 7.5 inches, and the overall length including bulkheads

is 26.6 inches.

The nitrous oxide is filled as a soft-cryogen, between -40 and -60 ºF.

This increases the oxidizer density and lowers the pressure required to

maintain it as a liquid. To help maintain low temperatures, the tank is

insulated with low temperature foam insulation. In the event that the

oxidizer heats and over pressurizes, the system is equipped with a burst disk

which activates at 675 psig to vent pressure. This condition requires a

complete reset of the system, but prevents catastrophic failure.

The nominal oxidizer pressure is 500 psia. This puts the tank under 15 ksi of hoop stress. At -50 ºF operating temperature,

the wall has a yield stress of 38 ksi33

, and a FOS of 2.5. At temperatures of 100 ºF, the FOS drops to 2.3. This large FOS

primarily results from the expense needed to acquire a thinner walled tube of the required dimensions.

Additionally, the tank features a diffuser to even out the flow of gaseous nitrogen into the oxidizer tank and minimize the

introduction of gas into the liquid nitrous oxide. The tank bulkheads contain flanges to center the tank inside the rocket. Each

flange has four channels around the circumference to allow for wire harnesses

to pass alongside the tank and past the flanges.

E. Oxidizer Control Valve

To control the oxidizer flow valve actuation, a pneumatic piston system

was developed which is able to open or close the ball valve. This system uses

a dedicated on board compressed air reservoir and actuates in both directions,

achieving at least two full open close cycles. Since the HyPE 1B was not

designed to restart, the valve is only required to perform one cycle. The piston

must open the valve as part of the ignition sequence, and must close the valve

as part of the main engine cut off.

Due to the cost prohibitive nature of 3-way solenoid valves, this system

uses two simple solenoid valves arranged in two pairs. Each valve pair is

connected to one end of the piston.

F. Combustion Chamber

The HyPE 1B combustion chamber consists of a

chamber wall, ablative liner, injector, igniters, pre-

combustion chamber, fuel grain, post-combustion

chamber, and nozzle. The layout is shown in Figure 13.

All parts are made from 6061-T6 aluminum alloy unless

otherwise specified. The chamber itself is made from an

extruded aluminum tube. The outside diameter of the

chamber is 5 inches and the overall length from thrust

plate to nozzle exit is 32.1 inches.

1. Injector

The injector is the heart of the hybrid rocket engine. This is where the liquid

oxidizer is introduced to the solid fuel grain. The injector determines the flow rate

of oxidizer, which in turn controls many key parameters including: regression rate,

fuel burn rate, O/F ratio, and chamber pressure. Further, the effectiveness of the

injector determines how efficient combustion will be. Ideally the injector will

atomize the liquid into a gas but in practice the goal is to achieve the finest droplet

size possible.

The HyPE 1B uses an impinging showerhead injector as shown in Figure 12. In

this configuration, the oxidizer flow is split into streams which are all directed at a

single point, causing the streams to collide, creating very fine droplets. This

produces a highly atomized flow with enhanced combustion efficiency compared to


7

Figure 14. Ignition

Ring (Pyrotechnic

charge not shown)

a standard straight showerhead34

.

The injector orifice determines the pressure drop across the injector, or the ΔP. A large ratio of pressure drop to chamber

pressure, ΔP/P, stabilizes combustion and prevents backflow of hot combustion gases into the oxidizer tank. Meanwhile, a

lower ΔP/P saves weight by allowing for a less structural oxidizer tank as well as a smaller pressurant tank. The HyPE 1B

optimizes the ΔP/P at 25%. A check valve between the oxidizer tank and the combustion chamber prevents gases from

moving back up to the tank in the event of a backflow. Based on the oxidizer pressure of 500 psia and losses in the oxidizer

plumbing line, the HyPE 1B’s chamber pressure is designed at 400 psia.

The total injector orifice area, or the combined area of all the injector holes, is related to mass flow, oxidizer flow rate,

and pressure drop35

:

2

inj ox

ox

kA m

P

(5)

K is the flow loss coefficient of 1.735

which has been verified in-house by cold flow testing. Design oxidizer flow rate is

3.53 lb/s (1.60 kg/s) and nitrous density is 68 lb/ft3 (1089 kg/m

3)

36. This produces a design injector area of .0834 in

2. Four

concentric rings were used, with 6, 12, 18, and 24 holes each (60 in total). The impingement point was chosen to be 2.0 in

from the injector to allow space for the igniters.

2. Ignition System

Ignition energy requirements in a hybrid rocket depend on initial oxidizer flow rate and fuel

volatility, which can be met very simply through adequate heating of the fuel grain in the

presence of an oxidizer37

.

The HyPE utilizes a simple pyrogenic ignition system which provides hot gasses at high

pressure for initiation of the combustion process. An empirical formula for the heat required

for hybrid ignition, similar to the Bryan-Lawrence Equation for estimation of ignition energy

requirements for solid rocket motors38

, was not found. However, required heat can be estimated

using equilibrium combustion thermodynamics39

. If heat available from igniter combustion

exceeds the estimated required heat by an appreciable margin then ignition can be ensured40

.

The ignition system consists of an assembled “ignition ring” that lies in the pre-combustion

chamber just aft of and flush against the injector. Six half-slices of D-class Aerotech model

rocket motors act as the pyrogenic igniters seen in Figure 14. They are ignited pyrotechnically

by commercial string-light bulbs filled with FFFFG black powder triggered by a 12 volt power

source.

In practice, the “ignition ring” houses two alternating sets of three igniters wired in parallel, adding a layer of redundancy

to the system. Each igniter has its own pyrotechnic charge. The wires run through the fuel grain port and out of the nozzle

and are cleanly forced out upon engine start.

Design of the ignition system was motivated by simplicity, accuracy, reproducibility, and reliability, which have been

improved over three major development phases since its inception for the HyPE 1A engine. The concept has been shown

effective in its application in solid rocket motors and in the similar cylindrical pyrogenic preheater grain system used in the

RATTworks M900 that the UCLA Rocket Project launched at the 5th

ESRA IREC. Proof of concept and design optimizations

have been achieved through multiple firings of the Hype 1AT engine.

3. Pre/Post Combustion Chamber

The pre-combustion chamber holds the igniter system and also provides a space between the injector and fuel grain. This

allows for the oxidizer to impinge before reaching the grain, and also allows a recirculation area for enhanced mixing and

heating34,35

.The pre-combustion chamber length was characterized based on the ignition system.

The post-combustion chamber provides a space between the fuel grain and nozzle, allowing combustion to complete

before the products enter the nozzle. A longer post-combustion chamber leads to better combustion efficiency, but also to

increased weight of the engine. The size used was based on a rule of thumb presented in Humble which suggests the optimal

post-combustion length is twice the fuel port diameter35

. This gives the post-combustion length as 2.25 in.

4. Ablative Liner

The ablative pre-preg liner protects the combustion chamber wall from the hot combustion gases. By burning sacrificially,

the ablative produces a film of cooler gases which surround the chamber wall, protecting it. This is one of the simplest and

most cost effective methods of cooling an engine. The liner is expendable and must be replaced after each launch.

The best performing and most reasonable alternative was identified to be a silica-polyimide composite41

. However, for

convenience, the use of carbon fiber was investigated for an ablative. From preliminary in-house torch tests, carbon fiber was

found to work as an adequate ablative.


8

Figure 15. MOC: Bell Nozzle Design

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.160

0.01

0.02

0.03

0.04

0.05

0.06

length [m]

heig

ht

[m]

MOC Min Length Bell Nozzle

X: 0.1546

Y: 0.05123

Figure 16. Pneumatic Umbilical

Pusher System

G. Nozzle

In the HyPE 1A and HyPE 1AT, a divergent conical nozzle was

used. The flow at the exit of the conical nozzle is not all parallel to the

path of the rocket and so there are losses in usable thrust39

. A bell

nozzle has been selected to replace the conical nozzle as it channels all

the thrust parallel to the flight path of the rocket, increasing the

efficiency of the rocket nozzle.

The Method of Characteristics (MOC) was used for designing the

minimum length bell nozzle. The MOC works on the principal that fluid

properties are constant along characteristic lines. As these characteristic

lines intersect, a system of equations can be iteratively solved to update

the fluid properties at the new location, creating the boundaries of the

nozzle wall as shown in Figure 15. In order to minimize the length of

the divergent section of the nozzle, the angle after the throat is based on

the Prantl-Meyer expansion fan42

.

As a result of the new nozzle design, the same expansion ratio and

thrust has been maintained; however, efficiency has been increased by 2.2% (increasing usable thrust to 1004 lbs) and

decreasing the mass of the nozzle by 15%. The final bell nozzle design has been validated using simulations in SolidWorks

Flow Simulation and an in-house program using a finite difference MacCormack scheme with artificial viscosity42

.

H. Umbilical Fill System

The HyPE 1B utilizes umbilical fill lines which can be remotely actuated.

To achieve this end, the fill lines for both oxidizer (nitrous oxide), and

pressurant (nitrogen) connect to the rocket through quick disconnect couplings.

These couplings are connected to a pneumatic piston-pusher system, shown in

Figure 16, which mechanically disconnects the fill lines, pushing them away to

avoid damaging the rocket during liftoff.

I. Analysis

For propulsion analysis, development continued on the in-house New Optimization Program (NOP). The program accepts

a variety of engine performance inputs including oxidizer flow rate, fuel grain dimensions, and chamber pressure, and

calculates the static thrust. The program then accepts simulation inputs such as initial rocket mass and initial altitude. The

NOP simulates the rocket’s trajectory by locking the engine’s design parameters and allowing flow rates and pressures to

change as the burn progresses.

The NOP uses the 1976 US Standard Atmosphere43

for atmospheric conditions. It also accepts a Drag vs. Mach table as

an input, discussed in section II.B. Other minor considerations include the buoyant and the gravity forces which decrease

with altitude.

The program operates through a time step method. For each instant in time, the program calculates the current thrust,

atmospheric conditions, drag, weight, acceleration, velocity, position and propagates this information ahead in time, iterating

until apogee is reached. The program then plots the results of the simulation, including plots for altitude, velocity,

acceleration, thrust.

A rough overview of the calculations that are involved in the NOP follows. The increase in port radius for any Δt:

t t t tr r r t (6)

Where r is bore radius. This is used to calculate the burn characteristic of the fuel grain assuming constant oxidizer flow rate.

However, the NOP takes this a step further and allows oxidizer flow rate to change as a function of oxidizer pressure. Toward

the end of the burn, the pressurant tank drops below the regulator pressure and the oxidizer tank pressure begins to drop. This

converts the system from a pressure regulated system into a blow down system.

Rewriting the injector equation (5) gives:

2

2

oxox c c

inj ox

m kP P P P

A

(7)

The equation for chamber pressure is34

:

c

t

cP m

A

(8)


9

Figure 17. MECO System Flight Test

Platform

Where c* is characteristic velocity, and At is throat area. Substituting fuel and oxidizer mass flows for the total flow rate in

equation (8) and then substituting into equation (7) gives:

*2 1 1 2

22

2

n n n

ox ox ox f p ox

ox inj t t

k c cP m m ar l m

A A A

(9)

This equation is insightful in that everything inside the brackets is known at any time step. Pox is also known, as the NOP

keeps track of oxidizer pressure during the transition to a blow-down. This fully defines the engine’s performance for any

oxidizer pressure and port radius. Every time step, the NOP solves equation (9) numerically for oxidizer flow rate. This

allows the NOP to continue modeling the engine as it runs off-nominal These changes in oxidizer pressure cause changes in

both oxidizer and fuel flow rates, as well as changes in chamber pressure and Isp. This yields an improved performance

analysis.

To size the engine, oxidizer flow rate was varied as the main independent variable. As this changes, the fuel grain scales to

maintain the same O/F ratio. The required volume of nitrous was then found, and the CAD model and drag analysis were

changed based on the new geometries and parameters. The NOP was then run to determine the altitude given by this

configuration. The procedure was iterated until the projected altitude reached 28,000 ft to allow for margin in weight and

drag.

IV. Electronic Systems

A. Avionics

A main engine cutoff (MECO) system was designed in order to ensure that the rocket will not exceed the desired altitude

of 25,000 ft AGL. The system has two primary tasks:

1) Calculate the rocket’s current altitude, velocity and acceleration while in flight

2) Shut off the engine at the appropriate time to reach the desired altitude of 25,000 feet above ground level given its

current altitude, velocity and acceleration

The system will calculate the rocket’s current altitude using data from an Analog Devices ADXL345 three-axis

accelerometer. Acceleration readings are then integrated through the use of a trapezoidal Riemann’s sum to obtain velocity,

and then again to obtain altitude44

. The microcontroller also calculates the rocket’s remaining mass as a function of time,

using mass flow estimates from engine testing, while the rocket’s current mass is greater than its defined inert mass.

The MECO system will then use the calculated acceleration, velocity, position, and mass values and predict what altitude

the rocket will achieve if its engine were to instantaneously shut-off. Acceleration due to drag and gravity is modeled by45

:

2

, 1,

1, where

2

Dk sim D air D CS k sim

Fa g F C A v

m (10)

The simulation is iteratively solved until velocity reaches zero (simulated

apogee). If the resulting altitude prediction is greater than or equal to the

desired altitude of 25,000 feet AGL, the engine is shut-off by issuing a

signal to the oxidizer control valve. The density of air ( ) is calculated

as a function of altitude using a polynomial equation curve-fitted to U.S.

Standard Atmosphere43

values while the drag coefficient ( ) table is

obtained from RASAero as discussed in section II.B.

The ADXL345 can measure changes in acceleration as small as

0.125 ft/s2 and the onboard microcontroller performs these calculations

every 75 milliseconds. A small scale test of the system was carried out

on a model rocket using an A8-3 Estes motor, shown in Figure 17.

Figure 18 shows the result the tests. The accelerometer accurately calculated the instantaneous altitude to within 7% of the

actual value, which was measured using an angle finder. The error may be explained by the use of the angle finder method,

which carries large uncertainty. The apogee prediction code averaged around 6% error and generally underestimated the

apogee by a few feet.

There are a few non-idealities within the current MECO system that were not addressed. The system does not take into

consideration the attitude of the rocket. Attitude was determined using the remaining two axes of the accelerometer; however

after a preliminary engine noise test on the MECO system using vibrations during a static hot fire, it was found that

transverse vibrational noise on the accelerometer would make attitude correction nearly impossible. The current MECO

algorithm will only be a first-order approximation of the rocket’s true altitude. A time delay based on engine performance

modeling will be used to prevent premature MECO.


10

Figure 18. Integrated Position (Calculated Apogee = 131 ft.

Actual Apogee = 126 ft.)

0

20

40

60

80

100

120

140

0 1 2 3 4 5 6

Po

siti

on

(ft

)

Time (s)

Instantanious Altitude Predicted Apogee Actual Apogee

Figure 19. Schematic of the Control and DAQ Systems

Figure 20. Final PCB Assembly

The MECO system consists of redundant

3.3V Arduino Pro Minis based off an 8 MHz

Atmega 168 microcontroller, ADXL345 three-

axis accelerometers, an SD card data logger, and

a series of transistor operated relays connected to

a 14.8V Li-Po battery pack, which will be used to

actuate the oxidizer ball valve and vent valve, as

well as a means of switching between ground and

vehicle based control. The MECO command will

simultaneously close the oxidizer control valve

and re-open the oxidizer vent to assure no

oxidizer remains when the rocket is recovered,

“safing” the rocket. The redundant MECO system

will shut-off only when both microcontrollers

indicate an engine cutoff condition. Acceleration,

velocity, position data, as well as engine status

messages will be logged to the on-board memory.

B. Electronic Hardware Interface

The rocket test/launch control and data

acquisition infrastructure consists of two sets of

electronics bridged by long lengths of CAT5

Ethernet cables, which allows for safe rocket

operation as shown in Figure 19. The launch box is

the user interface for control and data acquisition.

The Switch Board, contains an array of

transistor-relay switches that toggle according to the

signals sent out at the Launch Box, which switch

high voltage control signals to actuate the control

hardware. The op-amp ICs, located on the Switch

Board, amplify the thermocouple and load cell

signals. This circuit uses a bipolar power supply,

which allows transmission of differential signals,

which are less susceptible to ground plane noise46

.

There have been numerous improvements to the

test/launch infrastructure since 201047

. The Switch

Board has rapidly evolved from circuit prototypes to a full-fledged custom PCB, shown in Figure 20, in an effort to improve

circuit organization and reliability. Microcontrollers are still used to manage time sequenced events such as the ignition

sequence, but due to the general unreliability of the microcontroller experienced

during testing, manual overrides have been added to the ignition sequence.

C. Software

The GUI for monitoring live launch and test data was developed in National

Instrument’s LabVIEW programming language. The program’s front panel tabs,

shown in Figure 21, displays information such as volumetric flow rate, pressure

transducer readings, force transducer readings, and thermocouple readings. Raw

sensor data is linearized and then simultaneously recorded to file and displayed on

a time domain graph. The software also has integrated abort case detection code

that monitors pressures and alerts the user if a pressure reading is out of spec,

allowing the operator to take steps to safe the rocket. Future work will be done to

automate the control of the rocket in these abort cases. The GUI also displays which switches and controls have been

activated.

The graphical programming language is organized into modular sections comprised of various calculation threads for each

sensor. The calibrations constants used to linearize the raw data were obtained by manually calibrating sensors in the lab, and

then manually inputting them into the LabVIEW interface on the front panel.


11

Figure 21. Labview Front Panel Showing All

Indicators for a Hot Fire Test

An automated launch/abort system was investigated in order

to simplify the launch control process. A simple Arduino

microcontroller was tested for controlling the launch task

signals between the launch box and the pad. After an initial

field test with minimal automation functions, it was determined

that the microcontroller was too unreliable for the factors of

safety and reliability desired in the mission. The National

Instruments software and hardware are potential candidates for

attempts at automating future launch sequences.

D. Scientific Payload

The 10 lb payload to be carried to 25,000 ft AGL is located

in two separate locations. The nosecone contains scientific

instruments to record the acceleration, rotation, ambient

pressure, and nosecone temperature. The engine bay contains an HD camera compartment for recording the flight. The

nosecone payload houses a small project box that holds a 3.3V Arduino Pro Mini microcontroller based off an 8MHz Atmega

168, which interfaces with a BMP065 barometric pressure sensor and an IMU3000 inertial measurement unit (IMU) via an

I2C digital communication protocol as well as a DS18B20 ambient temperature sensor device and k-type thermocouple. An

instrumentation amplifier is required to amplify the signal from the thermocouples. Data from the sensors is stored to an SD

card reader for post flight analysis. The barometric pressure sensor is used to monitor the change in pressure as the rocket

travels through the transonic region and the pressure at the rocket’s apogee. This data can then be used to design a more

accurate altimeter system which can be used to initiate MECO. The IMU unit is used to characterize the trajectory of the

rocket as well as record the transverse and vibrational forces acting on the rocket during flight. The data will then be used to

help improve future designs of the rocket’s nosecone, body, and fins. The thermocouples are used to record the temperatures

experienced by the nosecone as a result of air friction, which give insight into the flight environment. The Aiptek PenCam

HD Trio is located in the engine bay and will film the rocket launch, flight, and descent. The camera has a battery life of over

two hours and a recording time of approximately 1 hour 40 minutes, and records to on board memory for post-flight analysis.

The HD camera records video viewed down the length of the rocket by using an aerodynamic faring with a 3/4” square

mirror angled at a 45 degree angle within the faring. Modifications were made to the internal circuitry of the camera to

provide external wire leads to operate the power and record functions using the screw terminals48

.

V. Recovery Systems

For the “reusability” criterion to be met, the HyPE utilizes a dual deployment recovery system that includes a drogue and

a main parachute. The purpose of the dual deployment scheme is to prevent the accumulation of large drift distances during a

majority of its descent.

A. Parachute Descent Rate

A parachute’s physical parameters are used to determine a suitable descent rate that limits the stress placed upon the

rocket when it impacts the ground. A custom descent rate calculator was created for the drogue and the main parachute

systems. The descent rate output function helps determine the viability of parachutes utilized for recovery. This calculator is

based upon equation (11)46

. The function’s inputs involve the weight of the rocket, the air density at apogee, which is held

constant to simulate a worst-case scenario, the parachute’s canopy area, and the coefficient of drag.

2 Tt

ref D

Wv

A C

(11)

The desired main parachute descent rate is between 15 and 25 ft/s led to the choice of the Sky Angle XXL parachute. The

drogue parachute’s ability to minimize drift while reducing stress during the main parachute’s ejection led to the selection of

the TAC-1 parachute49

.

Table 3. Descent Rate Inputs and Outputs

Main Parachute Drogue Parachute

Cd 2.9250

1.551

Aref (ft2) 12950

28.352

vt (ft/s) 18.8 82.9

Approximated Total Descent Time (s) 339


12

Figure 22. Concept

Behind the Drift

Application56

Figure 23. Graphical Results of the Drift Application Program

B. Parachute Deployment

At 25,000 feet above ground level, the air density becomes significantly reduced, which

limits the ability of black powder to combust rapidly and to create enough of a pressure

gradient to deploy the drogue parachute53

. The Rouse-Tech CD3’s 25 gram carbon dioxide

deployment system was adopted to cope with these high altitude conditions and ensure

deployment of the drogue parachute at apogee. The CD3 charge consists of a sealed black

powder controlled plunger which punctures a carbon dioxide canister, releasing pressurized

gas53

. Since the main parachute will be deployed at an altitude of 1,300 feet above ground

level, the air density is high enough for black powder to be effective. As a result of multiple

tests it was shown that 2.5 grams of black powder was sufficient to jettison the main

parachute from its compartment54

.

A single level of redundancy is employed in the recovery system. The system has two

separate circuits for each set of ejection stages: the drogue deployment stage and main

deployment stage. Two PerfectFlite HiAlt45K altimeters, which are capable of measuring

altitudes of up to 45,000 feet, are used for apogee determination and triggering of ejection

charges. Each altimeter is connected to its own battery, switch, and deployment charges.

C. Rocket Drift during Descent

In an attempt to decrease the recovery time of the rocket, an iteration-based program has

been created that determines a projected landing site and bearing from the projected apogee

towards the ground. The program requires wind velocity, wind bearing, and several physical

rocket parameter inputs. Equating the drag force equation51

to Newton’s second law yields equation (12)55

2

( , )2

D ref rel

x y

C A va

m

(12)

By iterating this equation for a layer of air every 100 feet from apogee to the ground, we obtain outputs which can be used to

determine rocket’s theoretical drift distance and bearing for the 100 feet layer of air. In order to account for the possibility of

error and fluctuations in the launch day’s wind patterns, a factor of a sixth has been incorporated into the program to set

upper and lower bounds of error. This is indicated by the red, transparent circle on the graph titled Drift Projection onto the

Cardinal Directional Plane56

. The results were simulated based upon wind profile data on May 19th

of 2011 at the Universal

Coordinated Time of 08Z, or 3 a.m. in the state of Utah57,58

.

VI. Conclusion

The UCLA Rocket Project has spent the past three years researching and developing hybrid propulsion technologies

for a student designed and fabricated rocket to be flown at the 6th

IREC Competition in Green River, Utah. The completed

engine design burns aluminized paraffin with nitrous oxide to produce 1004 lbs of thrust with an Isp of 202 s and burn time

9.9 s. With this engine, the 15 ft long 8 in diameter rocket will carry a 10 lb payload to 25,000 ft above ground level. An

avionics package calculates the rocket’s trajectory, analyzing the optimal time to cut thrust to achieve a precise altitude.

Ground control hardware is used to both control the rocket and collect data during tests. The payload takes scientific

measurements and video to aid in the development and construction of future UCLA rockets. Numerous precautions have

been taken to assure the rocket can be operated safely.

Each individual subsystem design within the rocket has been scrutinized and evaluated for weight, reliability,

simplicity, and safety. The electronics, structures, propulsion, and recovery systems have all been optimized both in theory

and design to culminate in UCLA AIAA’s most ambitious project to date, the HyPE rocket.


13

Acknowledgments

The UCLA Rocket Project would like to thank its numerous sponsors, without whom this project could never have left

the drawing board. We would like to especially thank C-P Manufacturing Corporation for their generous donation of time

and skill in machining our precision engine components. Thanks to Northrop Grumman, Lockheed Martin, AFRL, Aerospace

Corporation, Engineering Alumni Association, and the UCLA Engineering School for their monetary donations. Thanks to

Dresser for the donation of several great regulators. Thanks to Airtech, SFU, 3M, and SpaceX for the donation of composite

layup related supplies. To ROC Carbon for the donation of the graphite and machining for our nozzles. To Mouser

Electronics for the donation of electronic components and parts. To Quality Precision Clearing Inc. for the oxygen cleaning

our oxidizer system, and to all our other sponsors. Last but not least, a special thanks to Dr. Richard Wirz for his continued

guidance and support of the Rocket Project.

References 1.Jackson, J. (2011, January 6). Working with Composite Materials. 2.Chin, S. S. (1961). Missile Configuration Design. McGraw-Hill Book Company, Inc. 3.Krasnov, N. F. (1970). Aerodynamics of Bodies of Revolution. New York: American Elsevier Publishing Company, Inc. 4.Stoney, J. W. (1958). Collection of Zero-Lift Drag Data on Bodies of Revolution from Free-Flight Investigations.Washington: NACA. 5.Rogers, C. E., & Cooper, D. (2011). Rogers Aeroscience RASAero Aerodynamic Analysis and Flight Simulation Program. Lancaster:

Rogers Aeroscience. 6.Stoney, J. W. (1954). Transonic Drag Measurements of Eight Body-Nose Shapes. Washington DC: NACA. 7.The Great Soviet Encyclopedia, 3rd Edition. (1970-1979). The Gale Group, Inc. 8.Jack, J. R. Theoretical Pressure Distribution and Wave Drags for Conical Boattails. 9.Harder, K. C., & Rennemann Jr., C. (1955). On Boattail Bodies of Revolution Having Minumum Wave Drag - NACA Technical Note

3478. Washington DC: NACA. 10. Niskanen, S. (2009). Development of Open Source model rocket simulation software. Helsinki: Helsinki University of Technology. 11. Barrowman, J. S., & Barrowman, J. A. (1966). The Theoretical Prediction of the Center of Pressure. 12. (1964). Flutter, Buzz and Divergence. In NASA Space Vehicle Design Criteria SP - 8003. Washington DC: NASA. 13. Martin, D. J. (1958). Summary of Flutter Experiences as a Guide to the Preliminary Design of Lifting Surfaces on Missiles.

Washington: NACA. 14. McCormick, A., Hultgren, E., Lichtman, M., Smith, J., Sneed, R., & Azimi, S. (2005). Design, Optimization, and Launch of a 3"

Diameter N2O/Aluminized Paraffin Rocket. Tucson, Arizona: AIAA. 15. Van Pelt, D., Hopkins, J., Skinner, M., Buchanan, A., Gulman, R., Chan, H., et al. (2004). Overview of a 4-in OD Paraffin-Based

Hybrid Sounding Rocket Program. Fort Lauderdale, Florida: AIAA. 16. Dyer, J., Zilliac, G., Doran, E., Marzoña, M. T., Lohner, K., Karlik, E., et al. (2008). Status Update Report for the Peregrine 100km

Sounding Rocket Project. Hartford, CT: AIAA. 17. Grosse, M., & Schlatzke, G. (2008). Development of a Hybrid Rocket Motor Using a Diaphragm for a Small Test Rocket. Hartford,

CT: AIAA. 18. Davidy, A. (2008). Development of an Analytical Solution to the Heat Transfer Equation in the Burning Paraffin Fuel. Hartford, CT:

AIAA. 19. Evans, B., Favorito, N. A., & Kuo, K. K. (2005). Study of Solid Fuel Burning-Rate Enhancement Behavior in an X-ray Translucent

Hybrid Rocket Motor. Tucson, Arizona: AIAA. 20. Carmicino, C., & Sorge, A. R. (2006). The Effects of Oxidizer Injector Design on Hybrid Rockets Combustion Stability. Sacramento,

CA: AIAA. 21. Lohner, K., Dyer, J., Doran, E., & Dunn, Z. (2006). Fuel Regression Rate Characterization Using a Laboratory Scale Nitrous Oxide

Hybrid Propulsion System. Sacramento, CA: AIAA. 22. Dunn, Z., Dyer, J., Lohner, K., Doran, E., Bayart, C., Sadhwani, A., et al. (2007). Test Facility Development for the 15,000 lb Thrust

Peregrine Hybrid Sounding Rocket. Cincinnati, OH: AIAA. 23. Karabeyoglu, A., Cantwell, B., & Zilliac, G. (2005). Development of Scalable Space-Time Averaged Regression Rate Expressions

for Hybrid Rockets. Tuscon, Arizona: AIAA. 24. Karabeyoglu, A., Zilliac, G., Cantwell, B., De Zilwa, S., & Castelluci, P. (2003). Scale-Up Tests of High Regression Rate Liquefying

Hybrid Rocket Fuels. AIAA. 25. Karabeyoglu, A., Zilliac, G., Castellucci, P., Urbanczyk, P., Stevens, J., Inalhan, G., et al. (2003). Development of High-Burning-

Rate Hybrid-Rocket-Fuel Flight Demonstrators. Huntsville, Alabama: AIAA. 26. Greatrix, D. R. (2007). Model for Predicting Fuel Regression Rate in Hybrid Rocket Engines. Cincinnati, OH: AIAA. 27. Doran, E., Dyer, J., Lohner, K., Dunn, Z., Cantwell, B., & Zilliac, G. (2007). Nitrous Oxide Hybrid Rocket Motor Fuel Regression

Rate Characterization. Cincinnati, OH: AIAA. 28. Zilliac, G., & Karabeyoglu, A. (2006). Hybrid Rocket Fuel Regression Rate Data and Modeling. Sacramento, CA: AIAA. 29. Zilliac, G., & Karabeyoglu, A. (2005). Modeling of Prepellant Tank Pressurization. Tuscon, Arizona: AIAA. 30. Kilic, S., Karabeyoglu, A., Stevens, J., & Cantwell, B. (2003). Modeling the Slump Characteristics of the Hydrocarbon-Based

Hybrid Rocket Fuels. Huntsville, Alabama: AIAA.


14

31. Abu-Awaad, F. M. (April 2004). The Gas-Phase Heats of Formation of n-Alkanes as a Function of the Electrostatic Potential

Extrema on their Molecular Surfaces. Journal of Chemistry , 81-86. 32. Luxfer Cylinders. Product: LCX composite cylinder specifications. 33. MIL-HDBK-5J: Metallic Materials and Elements for Aerospace Vehicle Structures. (2003). Department of Defense Handbook. 34. Sutton, G. P., & Biblarz, O. (2001). Rocket Propulsion Elements. John Wiley & Sons. 35. Humble, R. W. (1995). Space Propulsion Analysis and Design, 1st Ed. Learning Solutions. 36. Braker, W., & Mossman, A. L. (1980). Matheson Gas Data Book, 6th Ed. Lyndhurst, NJ: Matheson. 37. Schuler, A. L., & Wiley, D. R. (1989). Hybrid propulsion technology program: Phase 1. Volume 3: Thiokol Corporation Space

Operations. NASA CASI. 38. National Aeronautics and Space Administration. (1971). Solid rocket motor igniters. NASA space vehicle design criteria, chemical

propulsion . Cleveland: NASA CASI. 39. Hill, P. G., & Peterson, C. R. (1992). Mechanics and Thermodynamics of Propulsion, 2nd. Addison-Wesley Publishing Company,

Inc. 40. Story, G., Zoladz, T., Arves, J., Kearney, D., Abel, T., & Park, O. (2003). Hybrid Propulsion Demonstration Program 250K Hybrid

Motor . Huntsville: NASA CASI. 41. Richter, G. P., & Smith, T. D. (1995). Ablative Material Testing for Low-Pressure, Low-Cost Rocket Engines. Cleveland, Ohio:

Lewis Research Center. 42. Anderson, J. (2003). Modern Compressible Flow 3rd Ed. New York, NY: McGraw-Hill Companies, Inc. 43. National Aeronautics and Space Administration. (1976). US Standard Atmosphere 1976. 44. Young, H. D., & Freedman, R. A. (2010). University Physics (13 ed.). San Fransisco, CA: Pearson/Addison Wesley. 45. Anderson, J. D. (2011). Fundamentals of Aerodynamics (5 ed.). Boston: McGraw-Hill. 46. Lattice Semiconductor Corporation. (2001, May). Differential Signaling. Application Note. Hillsboro, OR. 47. Prendergast, L. M., Zimmerman, K. U., Silva, W. A., Chang, P. C., Sivaram, S., Kentosh, B. C., et al. (2010). Paraffin Wax Based

Experimental Hybrid Motor. 48. Chang, P. C. (2011, May 5). 5/5/11 Electronics Team Update. Retrieved May 23, 2011, from HyPE: Hybrid Propulsion Experiment:

http://aiaa.seas.ucla.edu/RP2011 49. Knacke, T. (1992). Parachute Recovery Systems: Design Manual. Santa Barbara, CA: Para Publishing. 50. The b2 Rocketry Company. (n.d.). Retrieved February 2011, from b2 Rocketry Web Site: http://www.b2rocketry.com 51. Munson, B. R. (2009). Fundamentals of Fluid Mechanics (Vol. 6). Wiley. 52. Giant Leap Rocketry. (n.d.). Retrieved February 2011, from Giant Leap Rocketry Web Site:

http://www.giantleaprocketry.com/products_recovery.asp 53. Rouse-Tech, Inc. (2003). Rouse-Tech CD3 Instruction Manual. Rouse-Tech, Inc. 54. Rouse Tech, T. (2011, April). Determination of Ejection Charges for Parachute Deployment. (R. Abrantes, Interviewer) 55. Kavehpour, H. P. (2011, February). Application of Drag Forces. (R. Abrantes, Interviewer) 56. Charlesworth, P. (2004). The Impact of Wind Profile on High Altitude Flights by Fin Stabilised Rockets. United Kingdom Rocket

Assocation. 57. Correlations of Air Pressure to Wind Velocity above 20,000 Feet. (n.d.). IPS Meteostar. 58. Wind Profile from Ground Level to 20,000 Feet. (n.d.). Salt Lake City: National Oceanic and Atmospheric Administration: National

Weather Service.

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