Design and Analysis of a CMC Radiation Cooled LOX Densified Propylene Engine

American Institute of Aeronautics and Astronautics

1

Design and Analysis of a CMC Radiation Cooled

LOX/Densified Propylene Engine

Anthony T. Hartman,

1 Jos R. Torres,

2 and Jessica L. Rasmussen

3

California State University, Long Beach, CA, 90840

Eric Besnard4

California State University, Long Beach, CA, 90840

The paper presents the design and performance analysis of an 800 lbf LOX/propylene

rocket engine, designed to power the upper stage of a Nanosat Launch Vehicle (NLV). A

precursor version of this engine with relaxed requirements was static fire tested twice and

has undergone the first recorded flight test of a LOX/propylene engine by a team at

California State University, Long Beach back in 2009. The revised engine has an expansion

ratio of 70 with a targeted specific impulse of 342 s, and combustion and nozzle efficiencies

of 95% or greater. The injector features a LOX-centered split triplet with a center-mounted

TEA-TEB igniter. A single piece Cf/SiCm ceramic matrix composite is used for radiation

cooling of the thrust chamber. Fuel film cooling (FFC) is added to lower the chamber wall

temperature to the 2200-2400C range near the throat, maintaining the C/SiC structural

integrity and minimizing throat erosion. Computational Fluid Dynamics analyses indicate

that 10% FFC is sufficient to achieve this goal while almost meeting the targeted specific

impulse. A proof-of-concept version of this engine with relaxed requirements designed for

near sea-level testing is presented.

Nomenclature

= nozzle expansion ratio = nozzle efficiency c* = combustion efficiency = ratio of specific heats CMC = ceramic matrix composite

Cf /SiCm = carbon-fiber-reinforced silicon carbide matrix

CNC = computer numerically controlled

GHe = gaseous helium

GN2 = gaseous nitrogen

Isp = specific impulse (s)

L* = characteristic chamber length (m)

LOX = liquid oxygen

MC = Mach number in the combustion chamber

O/F = oxidizer to fuel ratio

PC = chamber pressure

TEA/TEB = triethylaluminum and triethylborane

1,2,3

Student, Mechanical and Aerospace Engineering Department, AIAA Student Member 4Professor, Mechanical and Aerospace Engineering Department, AIAA Associate Fellow

47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit31 July - 03 August 2011, San Diego, California

AIAA 2011-5936

Copyright 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.


2

I. Introduction

everal launch vehicles currently used to launch large satellites into orbit typically cost up to hundreds of million

dollars. Since the size of scientific

instruments continues to shrink with

improvements in technology, there is a

rapidly increasing interest in nanosatellites,

small satellites weighing 10 kilograms or

less. Recent US military focus on

Operational Responsive Space shows a

demand for more responsive launch

vehicles to deliver smaller payloads to low

earth orbit for a cost much less than that of

traditional launch vehicles. Instead of

relying on secondary or tertiary payload

opportunities, a dedicated Nanosat Launch

Vehicle (NLV) can provide the needed

flexibility in operations and targeted orbit.

The NLV being developed by Garvey

Spacecraft Corporation (GSC) in

partnership with California State

University, Long Beach would provide a

method of flying nanosatellites at a fraction

of the time and cost of a typical launch

program. Figure 1 shows the layout of the

proposed GSC NLV.1,2,3

The initial Nanosat

Launch Vehicle concept was proposed to

use LOX/propane or LOX/methane as

propellants, however recent trade studies

show that the use of LOX and chilled, or

densified, propylene offers an increase in performance over ethanol and methane1.

The envisioned NLV will utilize key technologies such as composite propellant tanks, ceramic matrix

composites, or CMC, for thrust chambers and densified propylene as fuel2. One way to increase performance in a

liquid rocket engine is the choice of high energy propellant. While some recent research has shown testing and

development of engines using LOX/methane4,5,6

the NLV baseline design uses LOX/ densified propylene because

recent trade studies7 have shown that it offers an 11% increase in performance over methane, and a 4% increase over

propane. While the specific impulse of propylene is slightly lower than that of methane, by chilling the propylene

thus densifying it, smaller propellant tanks can be used to achieve the same required performance, ultimately

allowing the NLV to lift a greater payload mass.

Several static fire engine tests have been conducted and flight tests have been conducted of several full scale,

low to medium fidelity, first stages3;4;5;6

. Initial ambient-temperature static fire tests and flight tests have been

conducted using propylene8. A design for an enhanced upper stage engine of the NLV is presented in this paper. The

requirements and constraints of the upper stage engine are driven by the baseline NLV requirements which include

the use of LOX and densified propylene as propellants, to fit within the 60 cm diameter interstage of the NLV, and

the ability to mount a thrust vector control unit to the engine, while providing 3560 N (800 lbf) of thrust at a targeted

specific impulse of 342 s.

The next section gives an overview of the engine design and engine sub-systems.

S

Figure 1. NVL Baseline design (dimensions in cm.)


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II. Engine Description

A. Overview The engine is designed for the upper stage of the NLV and is comprised of three main sub-components; a thrust

chamber, an igniter and an injector (Figure 2). As noted above, the engine utilizes LOX and chilled propylene as

propellants with a design thrust of 3560 N (approximately 800 lbs). Trade studies yielded a chamber pressure, Pc, of

1.03MPa (150psi). The thrust chamber is a CMC chamber with an expansion ratio of 70. For ignition, the engine

uses Triethylaluminum and Triethylborane (TEA/TEB) that ignites in contact with oxygen, allowing for reliable

ignition as soon as the TEA/TEB comes in contact with the LOX. The engine injector is an assembly of welded

plates that will be brazed to the thrust chamber. The flat head injector utilizes a set of split triplet impinging jets to

inject both the propellants and hypergolic TEA/TEB into the combustion chamber. The optimum Oxidizer to Fuel

ratio, O/F, was determined to be 2.610% fuel film cooling is used to keep the CMC thrust chamber below 2,473-

2,673 K (2,200-2,400 C) during the burn.

Figure 2. Upper Stage Engine Assembly

A graphical representation of the top down engine design process for this engine is presented in Figure 3. The

inputs for the engine design, shown in the red boxes, are O/Feng (oxidizer to fuel ratio of the engine), Pc (chamber

pressure), (expansion ratio of nozzle), Pa (atmospheric pressure), F (thrust), fuel film cooling percentage, At (area of throat), c* (combustion efficiency), and (nozzle efficiency). Based upon the type of propellant used, there is an optimum O/F ratio that corresponds to a maximum theoretical Isp. Using NASAs CEA (Chemical Equilibrium with Applications) program

15, the optimum theoretical Isp of propylene was determined to be 373 s, with a nozzle

expansion ratio of 70 and a chamber pressure of 1.034 MPa7. This O/F ratio is then taken and used to determine the

thermochemical properties of the resultant combustion gas, (ratio of specific heats), (molecular weight), Tc (combustion temperature), and c* (characteristic exhaust velocity). This thermochemical analysis is based on the

assumption of equilibrium combustion gas properties in the thrust chamber. This analysis was performed with a tool

developed by CSULB. After the thermochemical analysis is performed, the theoretical vacuum Isp can then be

calculated and the assumed efficiencies can be applied to get the actual vacuum Isp. The resulting nominal engine

performance is seen in Table 1.


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The mass distribution of the engine components is summarized in Table 2. The estimated and calculated masses of

each of the subsystems/components and margin are applied based on AIAA standards.19

A growth factor of 15%

was applied to calculated category components which were calculated with various analysis tools and a 30% growth

factor was applied to components whos mass were estimated. The LOX/chilled propylene engine with mass margin weighs 4.32 kg and has a thrust to weight of 84.

Table 2. Mass Breakdown

Component Mass (kg) Growth Category Growth Factor Mass w/ Growth

Chamber 1.86 Calculated 15% 2.14

Injector 1.89 Calculated 15% 2.17

Igniter 0.01 Estimated 30% 0.01

Engine Thrust-to-Weight Ratio = 86 Total 4.32

B. Injector

To account for injector losses, a single discharge coefficient is applied to characterize the combination of various

losses and is assumed to be 0.8. The pressure drop across the injector is set to be 20% of the 1.03MPa chamber

pressure (0.206 MPa) which is between the upper and lower bound of the desired pressure drop for this type of

injector based on experimental engines.13

In order to maintain high performance and minimize weight, an impinging jet design which is relatively simple

to manufacture and is well characterized from previous experiments is selected. This design has been characterized

from previous tests and missions to provide good overall propellant mixing and atomization, and thus higher

efficiencies. A split triplet configuration was chosen for this impinging injector because the increase in number of

orifices allows for smaller diameter of orifices providing better mixing and atomization of the propellants while

INPUTS

Vacuum Thrust 3,560 N (800 lbf)

PC 1.03 MPa (150 psia)

Engine O/F 2.6

At 0.0017 m3

70

PERFORMANCE

Theoretical Vacuum Isp 373.1

Combustion Temperature 3470 K (6270 F)

Targeted c* 0.95

Targeted 0.98

Fuel Mass Flow Rate 0.291

LOX Mass Flow Rate 0.756

Real Vacuum Isp

(w/o FFC

losses) 346.3

Table 1. Nominal Engine Performance Parameters

Figure 3. Top-Down Design Process


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allowing film cooling orifices in the outer plenum The injection angles were calculated such that the resultant

momentum directs the impingement of the streams to converge just above the throat. This convergence is a critical

aspect in CMC engine design to minimize chamber burn through and hot spots.14

The flat head injector body is fabricated from Inconel

625. The injector has 36 sets of split triplet orifices that

create an impingement pattern that converges just upstream of the throat. The triplet geometry consists of two sets of

straight LOX jets being impinged by outer and inner fuel jets, with fuel film cooling orifices along the outer

diameter to cool the chamber wall. Figure 4 shows the flow of propellants and mixing patterns in the injector. The

plenums in the injector are sealed by welding a top plate which will also serve as the brazing surface to the chamber

wall. The injector also utilizes film cooling orifices which inject fuel along the thrust chamber wall, creating a

gaseous layer of unmixed fuel that serves as a cooling barrier. This further protects the CMC chamber from the

intense heat of combustion which will reach 3470 K. The distance between the outer fuel, outer LOX, inner LOX,

and inner fuel is 4.331, 3.657, 2.793, and 1.928mm. The distance between the film cooling orifices are 15.959 mm.14

Figure 5 shows the configuration of the split triplet, film cooling, and TEA/TEB igniter orifices on the injector face.

Figure 4. Injector Cross Section and Propellant Flow


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Figure 5. Injector Orifice Configuration

For the unlike impinging jets, a 60 degree impingement angle is commonly used as a baseline to achieve

satisfactory flow characteristics. When the impingement angle is increased, the risk for propellant backlash

dramatically increases, potentially causing damage to the injector face. Impingement angles are driven by the mass

flow rate and momentum of the propellant.14

The aforementioned baseline impingement angles were then modified for our engine configuration. As noted

above, the impingement angles were optimized by evaluating the angles of the injector orifices with respect to the chamber geometry. The optimized outer, inner, and resultant angles are 21.67, 6.21, and 8.86 as seen in Figure 6.

The length-to-diameter ratio of the orifices is between 7 and 10 to ensure adequate flow of the propellants upon

injection into the thrust chamber.

Figure 6. Optimized Impingement Angles into Thrust Chamber

Outer

Angle

Inner

Angle

Resultant

Angle

= 21.67 = 6.21 = 8.86


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C. Thrust Chamber

The engine thrust chamber features a bell-shaped nozzle produced from carbon-fiber-reinforced silicon carbide

matrix (Cf /SiCm). The C/SiC can withstand temperatures up to 2,473-2,673 K (2200-2400 C) without much

ablation/erosion. The chamber is

made by forming carbon fiber to the

chamber geometry and using

chemical vapor infiltration where

gases are diffused into the molded

carbon. The C/SiC CMC material

also has great tensile strength and

its density is less than that of

typically used refractory metals.

The C/SiC combustion chamber

and throat sections are 3 mm in

thickness and are primarily

radiatively cooled. The nozzle is

tapered in thickness from 3 mm to 1

mm since the nozzle is subjected to

less internal pressure; shedding

weight off the engine. The 3 mm

chamber thickness is designed to

withstand the structural strains

resulting from the chamber

pressure as well as the temperatures of the hot combustion gasses.

The key design parameters for the combustion chamber are the characteristic length, L* and the chamber Mach, Mc.

For LOX/hydrocarbon engines, a common L* is between 0.8 and 1.0 meters. The efficiency of the vaporization,

burning, and mixing also depends on the propellant choice. Engines with larger L* values have greater mixing

efficiencies, but are larger in size and have greater pressure and heat losses, however engines with L* values too low

poorly mix and burn propellants. Here the L* was selected to be a slightly aggressive at 0.75 m, just below the

common range. When selecting a design chamber Mach number, the goal is to design the combustion chamber

pressure as close as possible, ideally equal, to the stagnation pressure. A smaller Mach number is ideal because the

low velocity allows more time for the propellants to mix and chemically react. Here a chamber Mach number, Mc, of

0.15 was selected.13

Figure 7 shows the process for the design of the thrust chamber geometry following design

methods from Space Propulsion Analysis and Design13

.

Due to the composite CMC chamber, interfacing the injector is difficult and is a critical process which will

consist of brazing the two components using Cusil-ABA.18

This braze will consist of a 10 micron braze foil that will

be heated in a vacuum furnace to a temperature above the liquid temperature of the braze foil and create a seal and

join the CMC chamber to the injector

D. Igniter

The ignition method uses triethylaluminum/triethylborane (TEA/TEB), a pyrophoric liquid combination, to react

with LOX starting combustion. This system features a cartridge that injects the TEA/TEB into the chamber through

the center of the injector; this method is a lightweight system and provides the TEA/TEB with direct impingement

of the LOX for more reliable combustion. This is done independently of the propylene feeds because TEA freezes at

-50 C20

, and TEB freezes at -93 C21

, and the densified propylene is at -150 C.

III. Thermal Analysis

Based Despite their high temperature capabilities, radiatively cooled CMCs still require film cooling when used with high performance bipropellants. A preliminary Computational Fluid Dynamics analysis was performed to

estimate the amount of film cooling required to keep the CMC below 2,473 K (2,200 C) in most locations, allowing

for localized temperatures up to 2,673 K (2,400 C).

Figure 7. Thrust chamber top down design


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A. CFD Model Setup

A commercial Computational Fluid Dynamics program from Metacomp Technologies, CFD++, is used

primarily to verify that the FFC percentage will be sufficient to keep the chamber cool during operation. The CFD

model uses an axisymmetric configuration, consisting of three different groups of meshes; one for the flow of the

core engine combustion gas, one for the pure propylene film cooling, and one for the solid CMC thrust chamber

thickness. The model assumes no reaction or combustion; instead it models a single combustion gas injected at the combustion temperature, corresponding to the O/F and chamber pressure. The FFC is injected as pure gaseous

propylene at the point where the film cooling impinges with the chamber, and this gaseous layer mixes with the

combustion gas through the entire length of the chamber. The inlet conditions inject the combustion core combustion gas and FFC at their corresponding temperatures and mass flow rates. Convective heat transfer in the

core and FFC is assumed, with conduction through the CMC, and radiation on the outer boundary of the chamber.

The radiative outer surface models the space thermal environment and uses an emissivity of 0.9 for the CMC.16

The

solution is steady state.

B. Preliminary CFD results

Figure 8 shows the computed Mach number distribution in the nozzle for a solution run with 12.5% FFC. The

calculated exit Mach number from the CFD model is 4.3, very close to the exit Mach of 4.26 calculated from the 1-

D analytical top down engine analysis mentioned above. The thrust calculated at the exit of the engine proved to be

close to the 1-D analytical analysis as well. A detailed quantitative analysis, however, is not performed to verify the

model since CFD++ assumes perfect combustion and is based on a frozen chemistry at combustion chamber

conditions while the analytical model with NASA CEA is based on shifting equilibrium conditions. Figure 9 shows

the temperature distribution across the chamber and Figure 10 shows a comparison of results with varying FFC

levels in with the mass flow rates were varied in the model to reflect the amount of propylene injected. The

maximum throat temperatures seen in the results over a range of FFC levels from 5% to 15% range from 2880 K to

2420 K. The results for FFC level between 12.5% and 15% yields a chamber wall temperature within the acceptable

range of the CMC material.

Figure 8. 12.5% FFC solution: Mach

Number Distribution


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Figure 10. CMC Throat Wall temperatures

at 5%, 10%, 12.5% and 15% FFC (in

degrees Kelvin)

Figure 9. 12.5% FFC solution, Temperature

Distribution


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C. Isp Losses from FFC

The actual vacuum Isp calculated from the top down engine design does not take into account the slight

performance losses associated with diverting fuel away from the combustion to be used as liquid fuel film cooling.

In order to characterize the losses associated with the FFC, we use Equation 1, where W is mass flow rate, the

subscript g is for the core gas, and the subscript c is for the gaseous propylene film cooling.17

Isp=(WIsp)g + (WIsp)c/(Wg + Wc) (1)

To calculate the Isp of the coolant gas, a hot propylene temperature of 2200K was assumed from the CFD model

to find the characteristic exhaust velocity, c*, which was then used to determine its corresponding Isp knowing its

thermochemical properties at engine operating conditions described in earlier sections. The Isp of the core gas was

calculated taking into account a mixing efficiency from the split triplet injector of 97%, based on test data from

similar style injectors.14

With these two calculated specific impulses, and the corresponding flow rate, Equation 1

was used to obtain the actual Isp of the engine with associate losses due to FFC show in Table 3.

Table 3. Isp Losses for various FFC %

FFC % Isp with FFC losses (s)

5% 351

10% 341

12.5% 337

15% 333

Since FFC levels between 10.0% and 12.5% yield a CMC temperature at operating conditions mostly well below

2,473 K (2,200 C) and only very locally near 2,673 K (2,400 C), the Isp of the engine would be between 337 and

341s which just falls short of the targeted Isp of 342 s. Since this analysis is based on preliminary CFD results,

further analysis with a more refined model that takes into account combustion of propellants to better capture the

real physics and heat transfer of inside the engine.

IV. Proof-of-Concept Engine A. Overview

In addition to the engine presented above, a proof-of-concept engine is developed for sea level testing to verify

critical engine interfacing. The proof-of-concept test engine will have relaxed requirements that will be used for a

sea level static fire test. These requirements differ because the engine will be tested in an atmospheric environment

with propylene fuel at ambient temperature rather than being chilled (densified) like the primary design. Since the

proof-of-concept engine nozzle is truncated much sooner (expansion ratio of 8 instead of 70) the chamber pressure

will be increased and the thrust requirement will be relaxed to 2,225 N. Also, since the test will use ambient

temperature propylene as fuel, its vapor pressure will be higher than that of cryogenic propylene.7 This leads to the

requirement of higher chamber pressure than the primary design, such that the pressure of the fuel does not drop

below its vapor pressure, causing what is known as cavitation, which can place extra stresses on the injector and

negatively impact performance.

Table 4 below shows the characteristics of both the primary and proof-of-concept test engines. The combustion

chamber will be a customer supplied chamber with a set geometry and thus the injector was designed around its

geometry, while still utilizing a similar split-triplet injection method and interfacing with a high temperature braze

between the injector and CMC chamber.


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Table 4. Engine Characteristics

Characteristic Reference engine Proof engine Units

Chamber Pressure, Pc 1.03 2.07 MPa

Fuel Temperature Cryogenic Ambient -

Fuel Vapor Pressure at Operating

Conditions


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B. Water Flow Testing

After manufacturing of the injector has been completed, the injector is then subjected to a waterflow test. Water is flowed though the injector into both the fuel and oxidizer ports to visually inspect the impingement of the

propellants (Figure 13). Water is also flowed separately through each propellant orifice set (LOX and fuel) to

determine the mass flow rate for a set time interval. This is used to determine the injectors discharge coefficient, which helps characterize all the pressure losses across the injector as well as verify the O/F.

Figure 13. Representative

Water Flow Test of Injector

Figure 12.Proof of Concept Engine Chamber and Injector


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C. Static Fire Testing

Static fire testing is conducted to verify the structural integrity of the engine, verify engine meets its required

performance and to ensure all engine components work properly as a system. Several static fire test stands built and

used by students at California State University, Long Beach utilize a regulated pressure system to conduct tests of

engines that range in size from 130 lbf to 4,500 lbf. During testing, data is recorded from pressure transducers at

several different locations in the propulsion system and chamber, with thermocouples used to analyze heat transfer

rates and combustion temperatures. A load cell connects the engine assembly to the static fire test stand to measure

the amount of thrust that the engine produces. Figure 14 shows the configuration of the test stand and a hot fire test

of a sea level prototype LOX/propylene rocket engine7.

Figure 14. Sea Level LOX/Propylene Static Fire Test Conducted by CSULB Students in Nov. 20087

V. Conclusion

With technology constantly growing, the demand for small nanosatellites to be launched is increasing while only expensive large launch opportunities exist in the market. The development of high performance rocket engines, such

as the upper stage rocket engine designed here, using LOX and chilled propylene as propellants provides the

opportunity to launch satellites at a fraction of what launches would cost.

Acknowledgements

The authors would like to thank their peers in the Mechanical and Aerospace Engineering department at CSULB and

their colleges in the CALVEIN rocket program for all their help.


14

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Design and Analysis of a CMC Radiation Cooled LOX Densified Propylene Engine

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