PLUME AND PERFORMANCE MEASUREMENTS ON A PLUG NOZZLE FOR SUPERSONIC BUSINESS JET APPLICATIONS

A Thesis

Submitted to the Faculty

of

Purdue University

by

Alexander Michael Sandroni

In Partial Fulfillment of the

Requirements for the Degree

of

Master of Science in Aeronautics and Astronautics

May 2009

Purdue University

West Lafayette, Indiana

UMI Number: 1469911

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ii

ACKNOWLEDGMENTS

Thanks are in order for a number of people:

First, I cannot thank Scott Meyer and Yu Matsutomi enough for their constant mentoring

in all aspects of being an engineer. Without them, my time at Purdue would have been

much less fulfilling.

Thanks are also due to Dr. Steve Heister for his constant guidance, wisdom, and most

importantly, patience over the past two years.

Rob McGuire deserves credit for a myriad of things. Simply enough, the BANR would

not be operational if it weren’t for Rob’s expertise in making hardware and fixing our

mistakes. His craftsmanship is second to none. Michelle Kidd and Joan Jackson were

also instrumental to the project’s success.

Three fellow graduate students also deserve special recognition. Thanks to Adam Trebs

for allowing me to ride his coattails, Chase Cummings for being so capable as to need no

coattails whatsoever, and John Tapee for being the catalyst needed to make this work

what it is.

Finally, thanks to the United States Navy for giving me the opportunity to take advantage

of one of the best universities with the best people in the field.

iii

TABLE OF CONTENTS

Page

LIST OF TABLES ............................................................................................................. vi

LIST OF FIGURES .......................................................................................................... vii

ABSTRACT ....................................................................................................................... xi

CHAPTER 1. INTRODUCTION ........................................................................................1

1.1. Motivation and Requirements for a Supersonic Propulsion System ............................1

1.2. Candidate Nozzle Configurations .................................................................................3

1.2.1. Forced Mixer-Ejector Nozzles ........................................................................... 4

1.2.2. Center-Body Plug Nozzles ............................................................................... 13

1.3. Mixing Phenomena and Noise Generation .................................................................16

1.3.1. Exhaust Plume Mixing ..................................................................................... 17

1.3.2. Noise Sources ................................................................................................... 18

1.3.3. Noise Mitigation ............................................................................................... 19

1.4. Objectives for Present Work .......................................................................................20

1.5. Initial and Follow-On Investigations ..........................................................................21

CHAPTER 2. FACILITY AND TEST ARTICLE DEVELOPMENT .............................22

2.1. BANR Facility Overview ...........................................................................................23

2.1.1. Core and Bypass Streams ................................................................................. 24

2.1.2. Operational Limits ........................................................................................... 27

2.2. Rig Operations ............................................................................................................29

2.2.1. Code for Predicting and Setting Conditions ..................................................... 29

2.3. Pressure Instrumentation .............................................................................................35

2.4. Traversing Rake System for Plume Diagnostics ........................................................40

2.4.1. Rake Design ..................................................................................................... 42

iv

Page

2.4.2. Traversing System Design ............................................................................... 45

2.4.3. Pressure Measurement Performance ................................................................ 53

2.4.4. Motion Control Program .................................................................................. 59

2.4.5. Other Uses ........................................................................................................ 60

2.5. Force Measurement .....................................................................................................60

2.5.1. Resolving Forces and Moments ....................................................................... 61

2.5.2. Load Cell and Data Acquisition Calibration .................................................... 66

2.5.3. Vertical Thrust Anomaly .................................................................................. 67

CHAPTER 3. RESULTS ...................................................................................................72

3.1. Facility Performance ...................................................................................................73

3.1.1. Feed Pressures, Temperatures, and Mass Flows .............................................. 73

3.1.2. Charging Station Conditions ............................................................................ 81

3.1.3. Forces ............................................................................................................... 89

3.1.4. Correlation of Desired Conditions with Actual Conditions ............................. 94

3.2. Nozzle Performance Results .......................................................................................96

3.2.1. Axial Thrust ..................................................................................................... 97

3.2.2. Off-Axial Forces .............................................................................................. 97

3.2.3. Efficiency and Discharge Coefficient Assuming Unmixed Streams ............... 99

3.2.4. Efficiency and Discharge Coefficient Assuming Perfectly Mixed Streams .. 111

3.3. Plume Rake Data .......................................................................................................115

3.3.1. Noise Concerns .............................................................................................. 116

3.3.2. Plume Maps .................................................................................................... 117

CHAPTER 4. CONCLUSIONS AND RECOMMENDATIONS FOR FUTURE

WORK .............................................................................................................................122

4.1. Concluding Remarks on Plug Nozzle Investigation .................................................122

4.2. Future Work ..............................................................................................................123

4.3. Final Thoughts and Lessons Learned .......................................................................124

LIST OF REFERENCES .................................................................................................127

v

Page

APPENDICES

Appendix A. Rig Setting Code ........................................................................................132

Appendix B. Plume Data Reduction Code ......................................................................138

vi

LIST OF TABLES

Table Page

Table 2-1 TMS Load Specifications ................................................................................. 61

Table 3-1 Cold Flow Test Condition Comparison ............................................................ 95

Table 3-2 Hot Fire Test Condition Comparison ............................................................... 95

vii

LIST OF FIGURES

Figure Page

Figure 1-1 Gulfstream SSBJ Concept Aircraft with Nose Boom Extended ....................... 2

Figure 1-2 Plug Nozzle Model with Full Contoured Center-body and Shroud .................. 4

Figure 1-3 Nozzle with Confluent Splitter.......................................................................... 5

Figure 1-4 Nozzle with Forced Mixer ................................................................................ 6

Figure 1-5 Olympus 593 Engine Nozzle ............................................................................ 7

Figure 1-6 Lobed Mixers with No Scalloping, Moderate Scalloping, and Deep Scalloping

.................................................................................................................................... 11

Figure 1-7 Example of a Full Scale Nozzle Utilizing Chevrons ...................................... 12

Figure 1-8 Plug Nozzle Flowfield Phenomena ................................................................. 14

Figure 1-9 Gap in Plug Nozzle Literature......................................................................... 16

Figure 2-1 The Bi-Annular Nozzle Rig, BANR ............................................................... 23

Figure 2-2 Bypass Air Circuit ........................................................................................... 24

Figure 2-3 Combustor and Torch Ignitor .......................................................................... 25

Figure 2-4 Charging Station.............................................................................................. 26

Figure 2-5 Aft Interior View of Charging Station ............................................................ 27

Figure 2-6 Pressure Rating of Combustor ........................................................................ 28

Figure 2-7 Predictive Program Functional Schematic ...................................................... 29

Figure 2-8 Sample Output of Required Rig Settings ........................................................ 33

Figure 2-9 ESP Module Suite ........................................................................................... 36

Figure 2-10 Core Charging Station Rake Drawing ........................................................... 37

Figure 2-11 Bypass Charging Station Rake Drawing ....................................................... 37

Figure 2-12 Photograph of Bypass and Core Charging Station Rakes ............................. 38

Figure 2-13 Plug and Shroud Tap Locations .................................................................... 39

viii

Figure Page

Figure 2-14 HFJER Plume Traversing System ................................................................. 41

Figure 2-15 Plume Rake Traverse System........................................................................ 42

Figure 2-16 Plume Rake Model ........................................................................................ 44

Figure 2-17 Pertinent Rake Dimensions in Inches ........................................................... 44

Figure 2-18 Basic Frame of Plume Traverse System ....................................................... 46

Figure 2-19 Parker HD-Series Actuator Used in Actuator System .................................. 47

Figure 2-20 Rake Free Body Diagram with Distributed Drag Load and Reaction Loads 48

Figure 2-21 Drag on Rake vs. Nozzle Pressure Ratio ...................................................... 49

Figure 2-22 Jackscrew Assembly for Providing Vertical Motion .................................... 50

Figure 2-23 Fully Assembled Plume Rake and Traverse System ..................................... 50

Figure 2-24 Plume Rake Control System Architecture .................................................... 51

Figure 2-25 Schematic of Pressure Response Problem .................................................... 54

Figure 2-26 Free Body Diagram of the Fluid Element ..................................................... 55

Figure 2-27 Frequency Response vs. Length for 0.125” Tubing ...................................... 59

Figure 2-28 Live Ring with Sign Convention Used ......................................................... 62

Figure 2-29 Live Ring Free Body Diagram with Measurement Load Cells Depicted ..... 63

Figure 2-30 Free Body Diagram of Force Measurement System with Calibration Loads

Depicted ..................................................................................................................... 65

Figure 2-31 Bypass Steerhorn and its Deflection While Pressurized ............................... 68

Figure 2-32 Vertical Force Generated by Bypass Horn Pressurization ............................ 69

Figure 2-33 Steerhorn Flange Supports ............................................................................ 70

Figure 2-34 Vertical Force Transmitted to Rig After Supports Installed ......................... 70

Figure 3-1 Gulfstream Plug Nozzle .................................................................................. 73

Figure 3-2 Time History of Feed Pressures ...................................................................... 74

Figure 3-3 Time History of Torch Ignitor Chamber Pressure .......................................... 75

Figure 3-4 Time History of Fuel System Pressures .......................................................... 76

Figure 3-5 Air Supply Temperature During A Typical Test ............................................ 77

Figure 3-6 Frosted Shroud and Rig with Hot Plug ........................................................... 78

Figure 3-7 Time History of Air Flow Rates from Turbine Flowmeters and Orifices ....... 79

ix

Figure Page

Figure 3-8 Core Spool Thermal Growth and Resulting Bypass Orifice Area Variation .. 80

Figure 3-9 Charging Station Quadrant Definition, Looking Upstream ............................ 81

Figure 3-10 Time History of Bypass Rake Pressures, Quadrant I .................................... 82

Figure 3-11 Time History of Bypass Rake Pressures-Quadrant II ................................... 82

Figure 3-12 Time History of Bypass Rake Pressures-Quadrant III .................................. 83

Figure 3-13 Time History of Bypass Rake Pressures-Quadrant IV .................................. 83

Figure 3-14 Time History of Core Rake Pressures-Quadrant I ........................................ 84

Figure 3-15 Time History of Core Rake Pressures-Quadrant II ....................................... 84

Figure 3-16 Time History of Core Rake Pressures-Quadrant III ...................................... 85

Figure 3-17 Time History of Core Rake Pressures-Quadrant IV ..................................... 85

Figure 3-18 Time History of All Charging Station Rakes-CR1 is Core Rake Quadrant I 86

Figure 3-19 Time History of Bypass Stream Total Temperature ..................................... 87

Figure 3-20 Time History of Core Stream Total Temperature ......................................... 88

Figure 3-21 Time History of Core Temperatures: Azimuthal Variation .......................... 89

Figure 3-22 Time History of Forces ................................................................................. 90

Figure 3-23 Time History of Combustor Pressure for Hot-Fire Test ............................... 91

Figure 3-24 Time History of Combustor Pressure for Cold Flow Test ............................ 92

Figure 3-25 Low Frequency Content of Axial Load Cells ............................................... 93

Figure 3-26 Low Frequency Content of Off-Axial Load Cells ........................................ 94

Figure 3-27 Axial Thrust vs. NPR for All Tests ............................................................... 97

Figure 3-28 Side and Vertical Force vs. NPR for All Tests ............................................. 98

Figure 3-29 Off-Axial Forces as a Percentage of Axial Thrust for All Tests ................... 99

Figure 3-30 Hot Stream Properties ................................................................................. 100

Figure 3-31 Nozzle Thrust Efficiency vs. NPR for All Tests Assuming Unmixed Flow

.................................................................................................................................. 102

Figure 3-32 Schlieren Image of NPR = 2.59 .................................................................. 103

Figure 3-33 Schlieren Image of NPR = 5.01 .................................................................. 104

Figure 3-34 Nozzle Resultant Force Efficiency vs. NPR for All Tests Assuming Unmixed

Flow .......................................................................................................................... 105

x

Figure Page

Figure 3-35 Notional Misalignment of Total Force ........................................................ 107

Figure 3-36 Dual Streams Sharing a Common Throat and Static Pressure .................... 108

Figure 3-37 Discharge Coefficient vs. NPR Assuming Unmixed Streams .................... 110

Figure 3-38 Thrust Efficiency vs. NPR for All Tests Assuming Perfectly Mixed Flow 112

Figure 3-39 Ratio of Measured Thrust Coefficient to Theoretical Thrust Coefficient ... 113

Figure 3-40 Discharge Coefficients ................................................................................ 115

Figure 3-41 Electrical Noise Induced by Stepper Motors and Drives on a Plume

Temperature Measurement ....................................................................................... 116

Figure 3-42 Magnetic Ferrite and Installation ................................................................ 117

Figure 3-43 2-D Total Temperature Map of Nozzle Exit Plane ..................................... 118

Figure 3-44 Total Temperature in Core Charging Station During Plume Mapping ....... 119

Figure 3-45 2-D Total Pressure Map of Nozzle Exit Plane ............................................ 120

Figure 3-46 Charging Station Total Pressures During Plume Mapping ......................... 121

xi

ABSTRACT

Sandroni, Alexander Michael. M.S.A.A.E., Purdue University, May, 2009. Plume and Performance Measurements on a Plug Nozzle for Supersonic Business Jet Applications. Major Professor: Stephen Heister.

The motivation for a supersonic propulsion system is primarily the faster travel

times that such a system would enable, but several conflicting requirements must be

satisfied in order for a nozzle configuration to be viable. In short, any candidate nozzle

configuration must be capable of generating high thrust levels across a large flight

envelope, and must also effectively integrate into a supersonic airframe. However, noise

requirements for operating over land and at commercial airports present a significant

challenge to designing an acceptable nozzle.

A facility has been designed to evaluate the performance of supersonic propulsion

nozzle concepts. The facility has been made operational and a plug nozzle has been

tested at a wide range of conditions. The specific tools developed for the facility include

a suite of pressure instrumentation, a traversing rake system for measuring temperatures

and pressures in the exhaust plume, and a six-axis force measurement system.

The facility has demonstrated reliable and consistent operation at all the test

conditions investigated. In terms of facility performance, the conditions entering the test

article were seen to be uniform in each stream at the test article interface. Also, the force

data from the plug nozzle shows the expected trends. The efficiency and thrust

coefficient of the plug nozzle is evaluated, and it is seen that the formulation of thrust

coefficient typically used in rocket applications is more appropriate for this particular

nozzle. Also, the discharge coefficient is evaluated and seen to have the expected trend,

although the actual values are somewhat low. Finally, plume data is shown to illustrate

xii

that the plume rake traversing system is functional, as well as to verify the mixing

assumptions used in the analysis of the nozzle.

1

CHAPTER 1. INTRODUCTION

1.1. Motivation and Requirements for a Supersonic Propulsion System

Perhaps the main motivation for consideration of a supersonic business jet (SSBJ)

lies in the fact that the cruise speed of this platform would be approximately twice that of

conventional business jets and airliners. To put this speed increase in perspective, it

could enable an individual to leave New York at 8 am, spend a full eight hour workday in

California, and return to New York by 10 pm. Additionally, the increased speed would

allow a traveler to be anywhere in the world in 10 hours with a single stop1. While the

time value of such a platform is obvious, the design of a viable SSBJ presents numerous

challenges in the areas of aerodynamics, airframe integration, and the propulsion system.

Several general requirements must be met in order for an SSBJ to be a viable option for

commercial air transportation. First, the airframe and propulsion system must be

effectively integrated to create a configuration suitable for supersonic flight.

Additionally, the propulsion system must be both capable and efficient across a large

flight envelope. Finally, any aircraft must satisfy Federal Aviation Administration (FAA)

noise regulations, including generating acceptable noise levels at takeoff and approach,

and the mitigation of any sonic booms created during supersonic cruise.

Conners et al2 describes the integration challenges for a Quiet Supersonic Jet.

From an overall airframe standpoint, the most challenging aspect are the shockwaves

generated as the aircraft approaches the local sonic velocity and then continues to

accelerate until reaching its target supersonic cruise speed. Historically, supersonic flight

of aircraft has been restricted to take place over unpopulated regions of land or water.

The Concorde was able to operate under these restrictions, operating at supersonic

velocities only while on the trans-Atlantic portion of its flight as well as operating under

less stringent noise requirements. In order to mitigate the issue of sonic boom noise,

2

techniques of airframe shaping and airframe morphing will be used to reduce the shock

strength during supersonic flight. In the Gulfstream model shown in Figure 1-1, the

aircraft is optimally shaped to reduce the strength of shock waves created during

supersonic flight. Additionally, the aircraft will utilize Quiet Spike™ technology, using

an extendible nose boom to extend in front of the aircraft in order to create weak oblique

shocks instead of strong oblique or normal shock waves. Ideally, these measures will

overcome aerodynamic and regulatory hurdles to a practical SSBJ.

Figure 1-1 Gulfstream SSBJ Concept Aircraft with Nose Boom Extended2

The propulsion system of an SSBJ is subject to various conflicting requirements

for performance. The need to propel the aircraft to high speeds and maintain these for

extended periods implies a system with high thrust capability. However, in direct

competition with the need for thrust is the requirement that the noise generated by the

propulsion system remain below acceptable levels for commercial use. Traditionally, a

high thrust capability might imply a high-bypass turbofan with a large-diameter fan, or

alternatively a low-bypass turbofan or turbojet with a higher jet exit velocity. However,

in an SSBJ application, both these options are unattractive for several reasons. A

subsonic aircraft is able to accept lower specific thrust engines such as high-bypass

turbofans in order to minimize takeoff noise and fuel consumption. While this propulsion

system provides high levels of thrust, integrating a large fan into an aircraft designed for

supersonic flight proves to be difficult, if not impossible because of the greater

3

aerodynamic and installation penalties at transonic and supersonic speeds. Increasing the

jet velocity can potentially generate the required level of thrust, but it also significantly

increases noise and fuel consumption. Unfortunately, because the noise generated at

takeoff and power cutback is an equally important issue, any system with a relatively

high exit velocity for these phases will not be viable for an SSBJ application.

Specifically, any viable nozzle must generate exhaust velocities on the order of 2000 feet

per second during cruise, and reduce this by approximately half to meet noise

requirements for operating in the vicinity of commercial airports.

These conflicting requirements on the propulsion system underscore the general

design guidelines that must be followed. First, the system must be reasonably sized in

order to integrate effectively into a supersonic airframe. Second, the propulsion system

must generate high levels of thrust across the entire flight envelope at a variety of nozzle

pressure ratios (NPRs). Third, the propulsion system must be “quiet” enough to avoid

negatively impacting the overall perceived noise level of the aircraft in flight, and to

satisfy applicable noise regulations.

1.2. Candidate Nozzle Configurations

The demanding requirements placed on the propulsion system of an SSBJ require

a number of compromises to be made between size, weight, complexity, and

performance. In order to achieve maximum thrust performance at cruise, the nozzle must

be designed to the ideal jet expansion ratio expected in the cruise configuration.

However, after the design is optimized for the cruise condition, it must also generate

sufficient thrust at takeoff and other regimes where the NPR might be only twenty

percent of the design cruise NPR. In all cases, any negative effects on the overall jet

noise level must be minimized. For SSBJ applications, two main types of nozzles have

been identified as viable candidates. Whurr3 describes several advanced nozzle concepts

including the forced mixer-ejector, which uses a shaped exhaust splitter to mix core and

fan engine streams and an ejector to entrain ambient air.

The other viable nozzle concept is the center-body plug nozzle, which uses a

contoured center body as an expansion surface in place of the traditional converging

4

nozzle walls. Figure 1-2 shows a plug nozzle drawn by John Tapee, demonstrating a full

length center-body contour surrounded by a cylindrical shroud.

Figure 1-2 Plug Nozzle Model with Full Contoured Center-body and Shroud

Both concepts have advantages and disadvantages in terms of thrust performance, noise

generation, and aircraft integration. While the mixer-ejector type nozzle has been the

subject of significant work in the past, the plug nozzle also offers several unique

advantages as far as pressure matching and aircraft integration. In either case, the

particular nozzle type chosen will depend on acoustic and aerodynamic considerations as

well as the actual vehicle application.

1.2.1. Forced Mixer-Ejector Nozzles

Forced mixer-ejector nozzles satisfy the requirements of an SSBJ nozzle through

several different mechanisms. First, for a conventional nozzle to generate the required

5

thrust, a convergent-divergent (CD) design must be used in order to increase the exhaust

velocity above Mach 1, which is the maximum for a conventional business jet or airliner

operating with a converging-only nozzle. Second, while the exhaust velocity must be

relatively high at cruise, it must be significantly lower at takeoff to mitigate noise. Third,

the exit area must be varied to provide pressure matching, maximizing thrust at all flight

conditions and minimizing the existence of shocks that exist in improperly expanded

supersonic flows.

A forced mixer-ejector nozzle primarily uses area variation and mixing to achieve

the desired performance. First, the mixing between the core stream and the fan stream is

enhanced by a forced mixer. The two flows mix inside the engine duct, and what results

is an exhaust jet of increased temperature and velocity uniformity. Additionally, the

detrimental noise effects from the high-speed and high temperature core stream are

reduced more than if the two streams were simply separated by a confluent splitter.

Figure 1-3 shows a sample convergent nozzle with a confluent splitter and Figure 1-4

shows the same nozzle with a forced mixer.

Figure 1-3 Nozzle with Confluent Splitter4

6

Figure 1-4 Nozzle with Forced Mixer4

Ideally, a properly designed mixer will significantly enhance mixing while incurring

minimal losses. This increased uniformity significantly reduces jet mixing noise and

improves specific fuel consumption. The second feature of the nozzle is an ejector that

entrains ambient air into the exhaust flow. For an SSBJ, these ejector “buckets” perform

two functions. With the upstream edges pivoted outwards, low speed, low temperature

ambient air is entrained into the jet exhaust. The mass flow of cold ambient air mixes

with the exhaust flow to further reduce the overall speed and temperature of the exhaust

exiting the nozzle. The increased mass flow provides a measure of thrust augmentation,

but more importantly it further reduces the noise generated at takeoff. After takeoff, the

upstream edges of the buckets are pivoted inwards to form a CD nozzle with an area ratio

designed to match the appropriate pressure ratio. This matching maximizes the thrust

generated and reduces the presence of plume shocks.

The most significant attempt at a commercial supersonic aircraft was the

Concorde, which utilized the Rolls-Royce Olympus 593 engine. This engine is a turbojet

so it did not include a forced mixer. However, Figure 1-5 shows the functionality of the

ejector buckets. These ejectors functioned in the manner described previously, beginning

with the buckets pivoted in a manner to entrain ambient air and progressing to fully open

7

buckets forming a CD nozzle at the cruise condition. An additional benefit of these

ejectors is the ability to pivot them fully closed to provide thrust reversing during

landing.

Figure 1-5 Olympus 593 Engine Nozzle5

Over the past several decades, most turbofan nozzle research has focused on the

design and performance of forced mixers, and several fundamental features of forced

mixers have been identified. In general, the greater the penetration of the mixer lobes

into the different streams, the more rapid the mixing. However, flow separation from

excessive turning and the associated performance loss must be avoided. Different mixer

geometries have been investigated ranging from simple splitters, to high penetration

lobed mixers with or without scallops, to chevrons as a means for achieving a high degree

of mixing in the shortest possible distance. Regardless of the mechanism generating the

increased mixing, the result is lower peak jet velocities leading to less jet mixing noise,

increased thrust, and improved thrust specific fuel consumption (TSFC). As an added

benefit, well-mixed exit flows are achieved after 1-2 mixing duct diameters when

including a forced mixer, a significant improvement over the 5-7 diameters required by

conventional ejectors. Using the ejector and forced mixer concepts together, core and fan

streams can be mixed more rapidly inside the engine duct and mixed further with ambient

air entrained by an ejector, resulting in higher performance and less nozzle weight.

One of the first major investigations of mixer geometry began in 1976 when

NASA initiated the Aircraft Energy Efficient Program in order to spur development of

8

more efficient aircraft and propulsion systems. This resulted in a detailed study of mixer

design and geometry for increasing performance and efficiency. Kuchar and

Chamberlain6 investigated different features of the mixer geometry, including the number

of lobes, perimeter-to-lobe ratio, and penetration. Also, lobes were “scalloped,” or cut

back to modify the mixers further in some cases. Total pressure and total temperature

were measured upstream of the nozzle and at the exit plane using rakes, and the mixer

and nozzle itself were instrumented with static pressure taps. For the test, cruise

temperature ratio was matched for the core and fan streams, and a pressure ratio of 2.4

was utilized to match the desired cruise condition of the notional aircraft.

The results of this particular study supported the understanding of how mixer

geometry actually affects the mixing process. It was seen that the penetration of the

mixer lobes has the most significant effect on performance and pressure loss. To a point,

increased penetration increases the performance until the flow turning became too great,

resulting in separation and associated pressure loss. The overall pressure loss was seen to

depend mostly on this flow turning and to a lesser degree on wall friction losses. The

number of lobes and the perimeter had little effect, while scalloping of the lobes was

found to increase mixing, especially with fewer lobes. Finally, the reduction of the gap

between the mixer and nozzle centerbody was seen to be the most important factor,

especially for higher bypass ratio engines.

Presz et al7 studied different mixer designs for use in augmenting ejector

performance. They found that at the low pressure ratios investigated, a more than 100%

increase in ejector performance was achievable through the use of forced mixers,

resulting in more complete mixing in significantly shorter ducts. Additionally, they

found that relatively aggressive designs could be used without resulting in separation.

Tillman et al8 also performed a detailed survey of different supersonic nozzles at

elevated exhaust temperatures of around 1000 degrees F, which is notable in that it

simulated a realistic turbine exit temperature. In order to assess the mixing of the

different nozzle designs, the experiment utilized a combination probe to measure total

temperature, total pressure, and static pressure. Additionally, a laser velocimetry system

was utilized to determine flow velocity in three dimensions. This study found that in

9

relation to baseline confluent nozzles, splayed nozzles displayed greatly enhanced mixing

due to the large scale axial vortices induced by the geometry. The potential core length

with the mixer nozzle was found to be approximately half that of the baseline nozzle,

demonstrating the increased mixing rate present in the exhaust. Detailed total

temperature and total pressure profiles were found for the baseline nozzles in this study,

but not for the more complex mixer geometries.

Tillman and Presz9 also investigated the overall thrust performance of lobed

mixer nozzles in comparison to standard circular nozzles. One of the initial concerns in

relation to thrust loss is the flow non-uniformity that is inherently necessary to create

large vortices. However, it was analytically determined that the thrust lost due to non-

uniformity from the mixer would be on the order or 0.1%. It was also seen that for both

un-choked and choked nozzle conditions, no appreciable thrust loss resulted in

comparison to a conventional circular nozzle. Thrust augmentation through the use of an

ejector was also investigated. It was seen that a properly sized ejector could result in

thrust gains for the mixer-ejector nozzle on the order of 10%. However, it was also seen

that improperly configured mixer-ejectors could actually lead to a decrease in thrust

relative to a normal ejector. Overall, it was found that it was possible to create mixer-

ejector configurations that were similarly effective at mixing, but could vary significantly

in their thrust performance.

Sokhey10 has investigated the use of ventilated mixers as a means for achieving

greater mixing with less total pressure loss. Greater mixing can generally be achieved by

greater vortex generation, which in turn is achieved by greater penetration between

streams. However, the greater penetration results in greater flow turning, which

eventually results in separation. The ventilation concept involves making slots in the

mixer lobes at appropriate places in order to energize the boundary layer and prevent

flow separation. Tests were conducted at various nozzle pressure ratios ranging from 1.5

to 2.5. Overall, it was concluded that the ventilated mixer concept significantly reduced

the pressure loss and also enhanced the mixing effectiveness. It was also determined that

for a high bypass turbofan, a ventilated mixer could provide a 0.4% gain in TSFC over an

10

unventilated mixer of the same geometry. Finally, it was seen that the noise reduction

potential was greater than that of a standard splitter nozzle.

Abolfadl and Sehra11 performed an investigation focused on the thrust

performance of mixers and how different geometric factors affect the performance of

forced mixers. Consistent with previous findings, they saw that TSFC, mixing, and

overall performance could be significantly improved through the use of a longer mixing

duct, scalloping, and increased lobe number. Abofadl et al 12 also specifically

investigated lobe geometry, including the effects of lobe height, wavelength, and

penetration angle. Several specific conclusions were reached regarding lobe geometry. It

was seen that the measure of mixedness used increased with increasing lobe height to a

maximum and a lobe height-to-wavelength ratio of unity. Mixedness also increased as

lobe penetration angle increased, up to a maximum at 20 degrees. These findings are

consistent with those found previously by other investigators.

More recent work by Presz and Werle13 has investigated the use of multi-stage

ejectors to improve ejector system performance. Ejector stages are placed in series

resulting in a device that generates higher diffusion rates, better wall cooling, and more

efficient flow mixing. Results showed that multi-stage designs were capable of higher

diffusion rates and greater thrust augmentation than single-stage counterparts. Mixing

was seen to be enhanced through the pumping of multiple streams of secondary flow.

Additionally, the multi-stage design was seen to exhibit better nozzle cooling and greater

insensitivity to aircraft installation effects such as blockages or other system losses.

Specific applications for this technology center around the need for thrust augmentation

with added IR signature reduction or added cooling. Accordingly, similar designs were

utilized on the Comanche attack helicopter and have been considered for rocket

applications to increase nozzle cooling.

Mengle14 has investigated the reduction of jet noise by modifying the scalloping

of lobed mixers, and also more recently, by using “chevron” geometry as opposed to

lobes in the mixer geometry. Scalloping, or removing sections of the lobe sidewalls, has

been seen to increase mixing effectiveness because of earlier interaction between the two

streams and the generation of smaller secondary vortices that merge with the main axial

11

vortices to increase their overall strength. These geometries are shown in Figure 1-6.

Scalloped lobe walls were also seen to reduce the overall noise in the more annoying

intermediate frequencies. However, the scalloping also resulted in more non-uniformity

in flow at the exit plane, which tends to be an undesirable side effect of mixers that

generate very complex internal flows. Overall, it was found that proper shaping and

sizing of the scalloped area resulted in decreased effective perceived noise levels

(EPNL), a benefit that was seen to increase as the overall thrust increased. However, an

overall decrease of thrust was found to be associated with scalloping, resulting in a trade-

off between decreased noise and increased thrust.

Figure 1-6 Lobed Mixers with No Scalloping, Moderate Scalloping, and Deep Scalloping14

The use of chevrons in mixers has also been investigated for its potential to

further decrease jet mixing noise. Mengle15 found that a combination of properly

designed chevrons placed on the mixer and nozzle lip can result in lower noise levels

than lobe mixers. This configuration is shown in Figure 1-7.

12

Figure 1-7 Example of a Full Scale Nozzle Utilizing Chevrons16

Additionally, it was found that chevrons only on the mixer also result in noise levels

below those of simple splitter mixers with no complicated geometry. The chevrons

reduce noise in a different manner than lobes, reducing low frequency noise while not

creating any significant high frequency “lift.” Lobes accept this high frequency “lift”

while reducing the low frequency noise to a greater degree. Essentially, the chevrons

were seen to produce a more gentle mixing process because the secondary flows induced

by chevrons were weaker in comparison to typical lobed mixers. This then is thought to

result in less internal high-frequency jet noise, resulting in less far-field noise.

Although the literature contains investigations performed by a number of

researchers under a range of conditions, a gap remains in the area that would be

representative of an SSBJ’s NPR range and size. While studies of NPRs between 3 and 4

and Mach numbers up to 1.5 have been performed, there is a lack of data at higher

conditions. Additionally, cold flow tests are significantly more common than hot-fire

tests with temperatures representative of turbine exit temperatures. Conveniently, the

new Bi-Annular Nozzle Rig (BANR) facility provides an ideal setup to study mixing at

both the exit plane and downstream distances. With the use of a total temperature and

total pressure measuring rake, the gap in data on higher flow conditions can be filled. In

turn, the effectiveness of mixers in SSBJ applications can be evaluated. The design of

the system used in the BANR facility will be discussed in detail in the next chapter.

13

1.2.2. Center-Body Plug Nozzles

Originally, the motivation for studying plug nozzles came from the desire to

design a single stage to orbit launch vehicle (SSTO). An SSTO requires maximizing

thrust across the entire flight envelope while simultaneously minimizing mass and

complexity. These factors directly translate to SSBJ applications where high levels of

thrust and easy airframe integration are of primary importance. Although the plug nozzle

is attractive as a potential nozzle concept, its merits are somewhat different than the

conventional convergent or CD nozzles. Although a nozzle with movable ejector buckets

can adjust to provide the desired exit area, the required area divergence can potentially

cause issues with airframe integration. These movable ejector buckets also add

significant complexity and weight to the propulsion system. The plug nozzle shown

below demonstrates the simplicity and ease of integration that such a nozzle offers. The

nozzle essentially consists of a cylindrical shroud and the plug expansion surface. While

this plug can translate to vary the throat for engine matching, its major benefit is that it

allows the flow through the nozzle to self-adjust to the ambient pressure. Ideally, what

results is a uniform flow without the periodic shock trains associated with improperly

expanded exhaust. Conners17 describes perhaps the most attractive feature of the plug

nozzle. While a conventional nozzle will have to open significantly to achieve the

desired area divergence, the engine nacelle footprint of a plug nozzle remains a constant

“trash can,” which in turn is easier to deal with from an engine integration standpoint.

Like a conventional nozzle, the plug nozzle is designed to a specific NPR, and as

a result, operates in three different regimes according to the existing pressure ratio.

Hagemann et al18 describe the flow phenomena experienced by plug nozzles. At pressure

ratios below design, the flow is expanded along the plug and interacts with the shear

layer, adapting to the ambient pressure through a series of compression and expansion

waves. At the design pressure ratio, the shear layer is parallel to the nozzle centerline and

the flow becomes essentially uniform, although perfect uniformity is not realistic for a

multi-dimensional flow. Finally, at pressure ratios above design, the plug nozzle

essentially operates as a conventional nozzle with the flow initially expanding to match

the lower ambient pressure, and then recompressing through a series of shocks. These

14

conditions are depicted in Figure 1-8 for typical SSBJ flight regimes. In general, the plug

nozzle results in relatively uniform flow at the nozzle exit with the exit pressure

approximately equal to the ambient pressure.

Figure 1-8 Plug Nozzle Flowfield Phenomena

Some of the early fundamental work done to develop plug nozzles involved the

design of an isentropic plug surface for expansion instead of a simple conical contour.

Lee19 and Angelino20 independently developed methods for determining the isentropic

contour for a planar flow in a plug nozzle. Angelino then further developed a more

robust design approach utilizing the method of characteristics. While these isentropic

contours expand the flow with minimal losses, they also require significant length for

high NPRs. For rocket applications with large pressure ratios, these contours become

extremely long and heavy. As a result, significant work has also been done to investigate

the effect on thrust when the plug is truncated to some degree. Hageman et al21 and

15

researchers at DLR performed extensive subscale experiments attempting to characterize

the effect of the flow over a truncated plug. Tomita et al22 also investigated this

transition region for conical and contoured plug surfaces over a large range of pressure

ratios.

For airbreathing applications, the most relevant plug nozzle investigations took

place under what originally was the Supersonic Transport Program and later the

Supersonic Cruise Research Program. A first generation low-angle plug nozzle was

extensively tested, examining internal performance, external performance, and

installation effects. Additionally, a second generation co-annular plug nozzle utilizing an

inverted velocity profile was also tested. Results showed that the low angle plug

demonstrated excellent thrust efficiency, while the co-annular plug’s efficiency was

somewhat less. In general, both nozzles showed promise, but needed additional work in

order to satisfy the goal of the study23.

As described, there has been significant attention paid to plug nozzles for rocket

and large supersonic transport applications. However, there is a distinct lack of research

that is directly applicable to SSBJ applications. This gap is depicted in . Although the

previous work on plug contours, thrust performance, and noise generation provides a

basis for comparing future designs and experimental results, data at representative

pressure ratios and scales will help to fill in a large gap in experimental results.

Fortunately, the pressure ratios and scales under consideration for SSBJs allow a plug

nozzle design to be directly tested at various conditions of interest with only minimal

scaling.

16

Figure 1-9 Gap in Plug Nozzle Literature

1.3. Mixing Phenomena and Noise Generation

Aeroacoustics studies the generation of noise by aerodynamic mechanisms. This

field is important for SSBJ applications as it pertains to the operation of commercial

aircraft, specifically the compliance with FAA FAR Part 36 noise regulations24.

Although noise generation is not a direct subject of the present study, it does provide

background for why mixer-ejector and plug nozzles are attractive for use in SSBJ

applications. As noted previously, the noise generated by any SSBJ is critical to its

overall viability. As a result, any future propulsion system must effectively minimize jet

noise at takeoff and approach. In order to do this, the jet exit velocity must be reduced as

much and as quickly as possible. Furthermore, supersonic shock noise from the exhaust

17

plume must be mitigated in order to reduce the intense noise emissions caused by

periodic shock systems that exist in improperly expanded plumes.

1.3.1. Exhaust Plume Mixing

In the case of an SSBJ, the fundamental problem to consider is the turbulent

mixing of two co-annular compressible jets. The mixing of core and fan flows represents

one case while the mixing of ambient air with the mixed exhaust represents another.

Ideally, a forced mixer combined with a mixing duct of appropriate length will result in

exhaust that is of a relatively uniform temperature and velocity, maximizing thrust and

efficiency of the nozzle. This exhaust then interacts with ambient air, which can have a

relative velocity anywhere between static and supersonic and be at a wide range of

temperatures. Depending on the nozzle design, exit conditions, and ambient conditions,

the exhaust will spread and mix with the ambient air in some manner, and a plume

boundary can be defined according to chosen criteria.

Plume mixing and spreading is usually characterized by velocity and temperature

decay as well as the physical spreading of the plume described by the jet half radius.

Extensive analytical and experimental work has been done in an attempt to develop

correlations and predictions for spreading and decay of the plume based on nozzle exit

Mach number and exit temperature or density. Chu25 describes several correlations that

have been developed to predict centerline velocity decay for both static and freestream

conditions. In general, it is seen that non-static ambient condition tends to reduce jet

mixing and increase potential core length, and the correlation between static and moving

freestream conditions is significant because of the relative ease of gathering static

freestream data versus simulated forward flight conditions. In addition to the velocity

difference between the freestream and the jet, it is seen that the density ratio plays a

significant role when the streams are of different composition. Total temperature can

also be related to the centerline velocity to predict the centerline total temperature decay.

Finally, through conservation of momentum, an expression for the jet half radius can be

expressed in terms of known temperatures and velocities.

18

The expressions described by Chu provide a baseline prediction of plume mixing

and spreading. However, they do make several simplifying assumptions, namely velocity

and temperature uniformity of separate streams. Most practical applications though

involve more complexity for a variety of reasons. In the case of a turbofan engine, the jet

exhausting into the freestream is rarely perfectly uniform in temperature or velocity.

Additionally, if a forced mixer or ejector is present, secondary flows will most likely be

present at the nozzle exit which will affect external plume mixing and decay. Finally,

recently gathered data has suggested that asymmetric nozzles can improve plume mixing

over axisymmetric designs26. These factors can be significant depending on the actual

configuration of the nozzle under consideration. Accordingly, analytical expressions are

sometimes unable to predict mixing and spreading as accurately as desired, requiring

experimental data to provide a basis for comparison. The investigations performed at the

BANR facility initially will not have the capability to directly observe velocity decay.

However, temperature decay will be directly observed with the added benefit of

accommodating streams of different temperatures.

1.3.2. Noise Sources

Louis et al27 describes the various noise sources that aircraft generate. For

subsonic commercial aircraft, the main noise concern comes from the general turbulent

mixing of core and fan streams with each other and with the ambient air. Shock-

associated noise is of lesser concern and sonic boom noise from subsonic aircraft is

obviously not applicable. For an SSBJ, all of these factors are concerns, and generally to

a greater degree than for subsonic aircraft with lower specific thrust propulsion systems.

The higher nozzle pressure ratios necessitated by the supersonic propulsion application

results in higher energy, higher speed, and higher temperature exhaust flows mixing with

the ambient air. Additionally, shock associated noise has the potential to be significantly

greater because of the supersonic exhaust jets. Finally, although not specifically related

to the propulsion system, the boom noise associated with supersonic travel becomes a

concern if the aircraft is to be operated over land and populated areas. While boom and

shock noise are major concerns during cruise, noise generated by the mixing of the

19

exhaust with ambient air is dominant during takeoff and power cutback/approach. In

general, at any particular flight condition, the noise from a supersonic propulsion system

can come from any number of mechanisms including Mach wave radiation, nozzle lip

radiation, shock turbulence radiation, shock unsteadiness, and turbulent mixing.

It is worth noting that historically there has been a significant volume of theories

and methods put forth for predicting jet noise, both in subsonic and supersonic plumes.

While many of the noise-generating phenomena are easily observed, a model to

accurately characterize the different mechanisms as far as their contribution to noise and

interaction with one another has yet to be fully accepted. However, these mechanisms

are commonly regarded as the major contributors to the overall level of noise produced

by supersonic exhaust, and consequently are of major concern when attempting to design

a propulsion system for an SSBJ or other supersonic aircraft.

1.3.3. Noise Mitigation

The different noise sources call for different techniques to reduce their effects on

the overall noise signature of the propulsion system. Jet mixing noise refers to the noise

generated by the turbulent interaction and mixing of fluids of different properties. Sir

James Lighthill28,29 developed the seminal relationship for the strength of jet mixing noise

as a function of jet velocity to the eighth power, and this relationship has been used

extensively as a baseline for predicting noise from jet mixing. To reduce the turbulent

mixing noise radiated outside the exhaust nozzle, the mean jet exit velocity and plume

velocity must be minimized, which is one of the primary goals of mixer-ejector systems.

Tam30 also describes the sources of noise generated from imperfect expansion of

exhaust plumes. Shock associated noise is the major distinguishing factor between

supersonic and subsonic jet noise and results in the large degree of directivity and

spectral content that is observed from supersonic exhaust. In a supersonic plume that is

imperfectly expanded, a train of shock waves will exist that is nearly periodic and decays

at a rate determined by the turbulent mixing layer. These shocks produce high levels of

acoustic emissions through two major mechanisms, broadband shock noise and jet

screech. Broadband shock noise and jet screech occur when large scale turbulent

20

structures in the plume interact with shocks in the plume. In the case of jet screech, the

disturbance is propagated upstream and excites the mixing layer, generating a distinct

audible tone. For shock associated noise, the most effective way to reduce its

contribution to the overall noise is to eliminate the presence of the periodic shock

structure in the exhaust plume. Accordingly, a mechanism must be incorporated to

appropriately expand the exhaust flow to an appropriate exit pressure to avoid a pressure

mismatch at the nozzle exit plane. Both variable exit geometry nozzles and plug nozzles

are helpful in achieving this condition.

1.4. Objectives for Present Work

The literature contains a significant volume of work relating to forced mixers,

thrust augmentation, noise generation, and plug nozzles, but it is clear that there is a need

for additional data pertinent to SSBJ applications. Most mixing investigations have been

performed at pressure ratios below those applicable for future high performance SSBJ

propulsion systems. Additionally, hot-fire data with realistic temperature conditions is

significantly less common than cold flow tests. Although scaling relationships can be

used to relate cold flow data to hot-fire data, the realistic conditions of a hot-fire test

allow for direct observation and evaluation of thrust performance and mixing.

Additionally, very little data exists in the literature with regards to plug nozzles for

airbreathing applications, both in the area of thrust performance and plume development.

The objectives of the initial SSBJ experiments were twofold. First, a facility had

to be developed to assess overall nozzle performance. Primarily, this includes equipment

to measure the thrust generated by the nozzle, but it also involves developing appropriate

diagnostic techniques for assessing mixing and plume characteristics. The second

objective was to actually perform experiments with scaled SSBJ nozzle hardware at

appropriate temperatures, pressure ratios, and flowrates. These experiments ideally

would yield data on thrust, mixing, and unsteadiness to allow for reasonable evaluation of

the particular nozzle under consideration.

21

1.5. Initial and Follow-On Investigations

The first hot-fire tests to be performed were on a plug nozzle scale model

developed by Gulfstream. This nozzle was tested across a range of nozzle pressure

ratios, ranging from approximately 2 at takeoff to 6.2 at the supersonic cruise condition.

Of prime interest was the internal flow characteristics of the nozzle, as well as any flow

unsteadiness generated. The thrust performance of the nozzle was also measured and is a

significant portion of the data analyzed in this thesis. Overall, the data gathered will be

used to determine the plug nozzle’s viability as a potential SSBJ nozzle, identify any

areas for further study, and to validate CFD results.

Following completion of the plug nozzle test matrix, data will be gathered on a

two dimensional IR signature reduction nozzle. Of primary interest in this investigation

will be the thrust performance and temperature distribution in the exhaust plume.

Although this nozzle is not an SSBJ concept, the BANR facility is ideally suited for

gathering the desired data. Finally, a forced mixer-ejector design from Rolls-Royce will

be evaluated. This nozzle will also be tested across a similar range of nozzle pressure

ratios to simulate operation of an SSBJ. For this particular configuration, mixing

effectiveness will be significant as well as thrust performance. Like the plug nozzle

investigation, the data gathered will be used to evaluate the performance of the nozzle,

including the particular forced mixer design and ejector contour. This will then allow for

a similar evaluation of the nozzle’s viability, and validation of CFD results.

The following chapters describe both the development of the facility and present

results from initial investigations. In Chapter 2, the development of rig operations is

detailed as well as actual hardware to measure plume mixing and thrust. Chapter 3

contains results from two sets of testing. Thrust data is presented from the plug nozzle

investigations that took place in January and February 2009. Initial plume mixing data

from investigations in March and April 2009 is also presented. Finally, Chapter 4

concludes the document with a summary of the knowledge and experience gained

through these first investigations with the BANR, and recommendations for future work

and improvements.

22

CHAPTER 2. FACILITY AND TEST ARTICLE DEVELOPMENT

To obtain relevant experimental data for a particular nozzle configuration, flow

must be provided that is either representative of realistic turbine exit conditions, or can be

correlated with such conditions. Purdue’s BANR, shown in Figure 2-1, was developed to

simulate realistic nozzle conditions representative of SSBJ engine cycles. With

representative conditions simulated, the thrust generated and other phenomena such as

plume mixing and spreading can be evaluated and correlated to full scale applications.

To obtain accurate information relating to the internal nozzle flow, plume characteristics,

and generated thrust, several diagnostic tools have been developed and implemented.

These include a suite of electronically scanned pressure (ESP) modules, a six-axis force

measurement system for measurement of forces and moments generated by the nozzle,

and a computer-controlled two-axis positioning system with measurement rake for spatial

mapping of total pressures and total temperatures. In general, these are the BANR’s

primary tools used for evaluating nozzle performance and are considered to be standard

equipment for any test. Additionally, the plug nozzle investigation required

supplementary instrumentation including high-frequency pressure transducers,

accelerometers, and Schlieren imaging. This required adding some temporary and some

permanent capabilities to the facility.

23

Figure 2-1 The Bi-Annular Nozzle Rig, BANR

The following sections describe the BANR and the tools that have been developed

and implemented to use it for nozzle investigations. These tools include a predictive

method for overall operation of the facility, pressure instrumentation to characterize

nozzle flow, a plume rake to map exhaust plumes, and a force measurement system to

resolve the forces generated by the nozzle.

2.1. BANR Facility Overview

Trebs31 describes the design methodology for the BANR in his M.S. thesis. In

general, the facility was designed to simulate nozzle inlet conditions of a turbofan engine.

Accordingly, it is capable of supplying separate heated and unheated streams to test

articles of varying sizes and configurations. The facility was designed to provide flows at

24

NPR’s ranging from 2 to 10, and core stream temperatures up to 1500 degrees F. The

current air system at Purdue’s High Pressure Laboratory (HPL) enables run times as high

as 6 minutes for low flow rate and low NPR conditions, decreasing to around 1 minute

for the highest flow test cases.

2.1.1. Core and Bypass Streams

The BANR provides the appropriate conditions to the test article with two

independently regulated air circuits. Both circuits are controlled with tunable PID-

controlled regulators with large flow coefficients. The bypass stream is shown in Figure

2-2, and the core uses similar components.

Figure 2-2 Bypass Air Circuit

With these regulators it is possible to set a desired condition (pressure or mass flow) and

utilize the control loop to maintain this condition throughout a test regardless of upstream

25

conditions. Downstream of these control valves, the flow in both circuits is set by critical

orifice plates. To minimize momentum contributions on force and moment

measurements from the flowing air, U-tubes are used on the core stream, and a horn with

flexlines is used on the bypass stream. The core circuit can provide heated flow using a

Rolls-Royce 501K combustor liner. This combustor is initially ignited with a propane-air

torch ignitor, and uses standard jet fuel to support combustion. The combustor, ignitor,

and fuel injector are shown in Figure 2-3.

Figure 2-3 Combustor and Torch Ignitor

The bypass air is unheated. Both streams enter the charging station where they pass

through flow conditioning screens and area contractions to ensure the cleanest flow

profile possible. The charging station also includes rakes to measure total temperature,

total pressure, and static pressure to verify the flow profile. The charging station is

shown in Figure 2-4.

Orifice

Plate

26

Figure 2-4 Charging Station

The streams remain separate up to the test article interface. This interface can

accommodate a splitter or forced mixer, a center body, and either separate nozzles or a

single nozzle. Figure 2-5 shows the interfaces as well as provides a clear view of the

flow conditioning screens and charging station rakes.

Orifice

Plate

Location

27

Figure 2-5 Aft Interior View of Charging Station

2.1.2. Operational Limits

The design of the BANR imposes several limitations on the maximum

temperatures, pressures, and flow rates that can be provided to a test article. As analyzed

by Trebs, the combustor imposes the most significant limit on NPR. When the

combustor operates at lower temperatures such as 800 F, the NPR can be up to 10.

However, as the temperature is increased to 1500 F, the NPR is limited to approximately

4 because of the lowered pressure rating of the combustor case at the elevated

temperature. These pressure limitations are shown in Figure 2-6.

28

Figure 2-6 Pressure Rating of Combustor31

Overall operational temperatures in the core are limited to approximately 1500 F because

of the limitations of the stainless steel parts. Fuel flow is limited by the fuel pump in use.

With the smaller fuel pump that was initially sized for the facility, fuel flow was limited

to approximately 1 gallon per minute at 600 psi. With a larger pump, the maximum fuel

flow is significantly increased to 4.7 gallons per minute at 1500 psi, which is sufficient

for all tests currently anticipated. Run time is primarily limited by air supply capacity,

although the large flow coefficients of the control valves allow for significant flexibility

in test length especially at low flow rates. Finally, flow rates and pressures are also

limited by the plumbing and orifice plates in use. Due to the large variation in flow

conditions desired, it may be necessary to change the core stream metering orifice so that

the required supply pressure is not higher than the limitations on the plumbing. For the

core stream, the maximum pressure is approximately 750 psi, and the bypass limitation is

approximately 400 psi. Fortunately, the core metering orifice can be changed easily.

29

Unfortunately, the bypass metering orifice requires complete disassembly of the rig.

However, the bypass orifice is sized such that it can provide mass flows of greater than

30 lbm/s, which is sufficient for the testing planned in the near term.

2.2. Rig Operations

In order to prepare for actual operation of the BANR, an analytical tool was

needed to convert a desired test condition into the actual rig settings required to achieve

the condition. In addition to the required settings, an approximate prediction of

conditions throughout the rig was also desirable to ensure safe operation. Finally, a set of

procedures needed to be developed for easy and consistent use of the BANR.

2.2.1. Code for Predicting and Setting Conditions

A MATLAB script was developed to translate a desired test condition into

information useful for operations. This code takes inputs of NPR, bypass ratio (BPR),

and desired core temperature and outputs the required orifice pressures and fuel circuit

settings to achieve the desired condition. In general, it is a tool that quickly performs a

one dimensional analysis that is intended to provide a baseline for operating at a desired

test point. Figure 2-7 shows the operation of the program in simple schematic form.

Figure 2-7 Predictive Program Functional Schematic

30

2.2.1.1. General Assumptions

This code makes several assumptions in order to simplify the analysis. In general,

there are a significant number of unknowns taken into account, and with more experience

operating the rig, these unknowns can be better characterized, resulting in a more refined

predictive tool. First, the pressure in the tank is assumed to be constant, which is

reasonable for a relatively short duration test, but less so for a longer test or one at very

high NPR, as the tank pressure will decrease significantly as well as the air supply

temperature. Second, all properties of air and jet fuel are assumed to be constant. Across

the range of temperatures and pressures experienced, this is reasonable although it does

introduce an error when a hot-fire test is being considered as the ratio of specific heats is

less accurate. Third, the blowdown of the air supply tanks is assumed to be a polytropic

process from the tanks to the nozzle throat. The process is assumed to be between

isothermal and adiabatic, with an appropriate polytropic constant. As tests are completed

and more data on this blowdown process becomes available, this constant can be

modified to more accurately represent the actual process that is taking place, even taking

into account differences between the core and bypass streams. Fourth, the two streams

are assumed to be unmixed. Finally, the code assumes a value for the discharge

coefficient of both the orifice plates and nozzle. The discharge coefficients for both the

orifice plates and nozzles should be the subject of close scrutiny as more test data is

gathered and analyzed.

2.2.1.2. Execution

The actual code used can be found in Appendix A. With the test day ambient

conditions given as well as the starting air tank supply pressure, the code first calculates

the required total pressure (Ptotal) at the throat using the desired nozzle pressure ratio

(NPR) and ambient pressure (Pa):

t aP =NPR×P Eq. 2.1

31

Next, the temperature at the throat (Tthroat) is calculated using a polytropic relationship

that incorporates the tank pressure (Ptank), tank temperature (Ttank), and total pressure:

n-1

nt

throat tanktank

PT =T ×

P

Eq. 2.2

The polytropic constant (n) describes the blowdown process, with a value of 1.0 being an

isothermal process and a value equal to the ratio of specific heats being an adiabatic

process. From empirical data, this value has been slightly more than 1.0. In the case of a

cold-flow test, the throat temperature calculated in Eq. 2.2 is used as the mean

temperature. For a hot-fire test, the core temperature is assumed to be the desired core

temperature. These two temperatures are then averaged using the desired bypass ratio

(BPR), core stream temperature (Tcore) and bypass stream temperature (assumed to be

equal to Tthroat) to get the mean throat temperature (Tmean):

mean core throat

1 BPRT = T +T

1+BPR 1+BPR

Eq. 2.3

Using this temperature, the desired total pressure, the throat area (Athroat), an assumed

discharge coefficient (Cdnozzle), and gas properties (γ and R), the total required mass flow

(•

totalm ) through the throat is calculated (with gc=32.2 (lbm-ft)/(lbf-s2) needed for

consistent units):

- γ+1• 2 γ-1

ctotal t throat nozzle

mean

γg γ+1m =P A Cd

RT 2

Eq. 2.4

This total mass flow is then used with the BPR to determine flows through the core

(•

corem ) and bypass streams (•

bypassm ).

32

••

totalcore

mm =

1+BPR

Eq. 2.5

• • •

bypass total corem = m -m

Eq. 2.6

Using these mass flows and temperatures, the orifice areas (Aorifice,core, Aorifice,bypass,

Aorifice,combustor), an assumed discharge coefficient for orifices (Cdorifice), and gas

properties, the respective bypass orifice and core orifice pressures (Porifice,bypass, Porifice,core)

required to reach the condition can be found, as well as the pressure in the combustor

(Pcombustor). For the orifice temperatures (Torifice,core, Torifice,bypass), a polytropic relationship

is again used between the tank and the orifices, requiring several iterations to achieve the

correct value. This allows for separate modeling of streams and the associated heat

transfer during the blowdown process.

-1γ+1•

2 γ-1orifice,corecoreorifice,core

orifice orifice,core c

RTm γ+1P =

Cd A γg 2

Eq. 2.7

-1γ+1•

2 γ-1core combustor

combustororifice orifice,combustor c

m RT γ+1P =

Cd A γg 2

Eq. 2.8

-1• γ+1

2 γ-1bypass orifice,bypassorifice,bypass

orifice orifice,bypass c

m RT γ+1P =

Cd A γg 2

Eq. 2.9

A hot-fire test adds more complexity to the prediction. The code utilizes NASA’s

Chemical Equilibrium with Applications (CEA)32 and iterates on a desired combustor

temperature to determine the appropriate fuel-air ratio and also the corresponding

combustor pressure and fuel mass based on the core air flow. The last portion of the code

calculates the pressure drops through the various components of the fuel supply system,

resulting in the required settings for the regulator and valve that control the fuel circuit.

33

The code summarizes this process and outputs all the input values and the pertinent

output values on the MATLAB desktop, as shown in Figure 2-8.

Figure 2-8 Sample Output of Required Rig Settings

2.2.1.3. Concerns and Potential Improvements

From initial testing, the code has already demonstrated good functionality and

accuracy. However, the predictive capability of the code can ultimately be improved by

several factors. First, the discharge coefficients of the orifice plates need to be evaluated.

This coefficient can then become a parameter in the code that is known with certainty.

As different nozzles are tested, their individual discharge coefficients will also need to be

determined. Next, the blowdown process of the rig needs to be carefully evaluated.

From empirical evidence, the rig tends to frost when flowing larger mass flows even after

only a short run, and air temperatures decrease significantly during a test. If this is not

properly accounted for, the code can miss on its predictions by a significant amount.

Perhaps the most work lies in the fuel supply circuit. Significant effort was put into

understanding the original fuel injector procured for use. However, it was seen that this

injector was incapable of maintaining combustion. Accordingly, an alternate injector

with different characteristics was used temporarily to support testing. The original

injector was then modified slightly and will require more analysis to verify its flow

characteristics. Also, the original valve in the fuel supply circuit was oversized, resulting

in poor control at lower flowrates. The problem is further complicated by the fact that

34

the pump that pressurizes the fuel does not have a characteristic as flat as that suggested

by the manufacturer. All these factors combine to make the performance of the fuel

circuit difficult to characterize. A more capable pump and a smaller control valve will

provide more useful and consistent operation of the circuit across all anticipated supply

needs.

There are also a number of concerns with the operation of the rig that do not have

straightforward fixes. First, as detailed in Trebs’ thesis31, screen stacks are used as flow

conditioners in both the core and bypass streams. The literature suggests methods for

predicting an appropriate loss coefficient for a screen in a similar setup. However, the

stacks were custom-made with six layers of mesh oriented at different angles, making it

difficult to make an actual prediction of the loss coefficient based on existing methods.

Also, it is impossible to calculate the pressure drop across the stack for a given condition

without actually measuring it. To this end, a port in the bypass stream immediately

upstream of the screens will provide information on pressure upstream of the stack. This

information will be used with the pressure upstream of the orifice and the charging

station rake pressures to calculate pressure drops across the orifice plate and screen stack.

Although it is impossible to do this for the core stream in the current configuration of the

rig, the data from the bypass stream will provide information on the pressure drop the

stack is seeing, and in turn the loads on the screen assemblies. This knowledge will

hopefully help to avoid any catastrophic event where a piece of the screen stack is blown

downstream, causing damage to the charging station rakes and the test article.

Finally, the separation of the bypass and core streams at their mixing or interface

plane poses a challenge to accurately determining NPRs. Theoretically, the mass flow

for a cold test needs to be divided according to the stream area ratio in order for the NPR

of both streams to be matched. This same concept applies for a hot-fire test as well,

although the increase in stagnation temperature and change in ratio of specific heats must

be accounted for33. As a result, it is more complicated to characterize the NPR in hot-fire

tests due to the range of temperatures that can be seen during a test. From data already

gathered, it can be seen that in some cases, the flow from one stream can influence the

other, resulting in uniform pressures but not necessarily the desired mass flows.

35

Although the temperatures, pressures, and mass flows of the two streams can be carefully

accounted for and matched relatively closely, some degree of mismatch is unavoidable.

Therefore, to effectively describe the overall NPR, a mass weighted average can be used

to combine two slightly different stream NPRs (NPRcore and NPRbypass) into a single value

( NPR )34 using their respective flow rates:

• •

core bypasscore bypass

•

total

m NPR + m NPRNPR=

m

Eq. 2.10

This value can then be used to correlate relevant data such as thrust, mixing, and other

parameters.

2.3. Pressure Instrumentation

The main source of pressure data for the BANR is a suite of ESP modules,

manufactured by Esterline Pressure Systems. The facility currently has 6 modules, each

capable of 16 simultaneous pressure measurements over a range of pressures. This suite

is shown in Figure 2-9. There are several benefits to utilizing these modules instead of

individual transducers. First, these modules are more flexible across a range of pressures.

They measure differential pressure with respect to a chosen reference pressure that can

easily be changed, allowing them to be used at most absolute pressures despite their finite

differential range. Accordingly, a higher accuracy can be achieved than with absolute

transducers as the full scale of the instrument will be less. The suite also has three of the

modules that are bi-directional, allowing sub-atmospheric pressures to be measured

without the use of a vacuum reference. Communication with the modules is through an

ethernet connection, allowing for simultaneous data transfer at up to 500 Hz per channel

without tying up all the analog data channels in the facility. To verify the accuracy of the

modules, high-accuracy Druck transducers are used to verify accurate reference pressure

36

measurements. Finally, the cost per channel for the data is significantly less than an

equivalent number of high-accuracy individual transducers operating simultaneously.

Figure 2-9 ESP Module Suite

In the current configuration, these modules primarily measure three groups of

pressures. First, they measure the total and static pressures in both the core and bypass

streams as measured by the charging station rakes. These rakes each have 5 total

pressure and one static pressure measurement, as well as 5 total temperature

measurements. As shown in Figure 2-10 and Figure 2-11, the probes are spaced such that

they are at the center of annuli of equal area, allowing the mean pressure or temperature

in each stream to be computed with a simple average of the measurements. The actual

rakes (uninstalled) are shown in Figure 2-12.

37

Figure 2-10 Core Charging Station Rake Drawing

Figure 2-11 Bypass Charging Station Rake Drawing

38

Figure 2-12 Photograph of Bypass and Core Charging Station Rakes

Additionally, the modules were set up to measure the static pressures along the plug

nozzle expansion surface and shroud. The layout of these taps was created by John

Tapee35 and a diagram is shown in Figure 2-13.

39

Figure 2-13 Plug and Shroud Tap Locations

Finally, a specific module is used to measure the total pressures from the plume rake.

This module has the largest range in the suite in order to account for the wide range of

pressures the rake can potentially see as it is traversed from quiescent air to flow at a

stagnation pressure of approximately 150 psi. As an added benefit, the pressure

connections to all the modules are made with plastic tubing and can be easily changed.

The majority of pressure measurements for the facility are “steady state,” or less

than 500 Hz. However, the plug nozzle investigation required high-frequency pressure

measurements taken between 20 kHz and 100 kHz. To accommodate this, high

frequency signal conditioners were added to the facility in order to interface with HPL’s

portable high-speed data acquisition system. Additionally, the necessary wiring and

connections were made for communication with a high-speed camera. A further

description of these capabilities can be found in John Tapee’s work35.

40

2.4. Traversing Rake System for Plume Diagnostics

The overall objective for performing plume diagnostics in the BANR facility is to

obtain quantitative data on the nozzle plume characteristics. This can include nozzles

with static mixers, mixer/ejector systems utilized in advanced nozzle concepts, transition

nozzles used in low infrared signature applications, or any other nozzle of interest.

Methods such as high frequency pressure measurement, hot-wire anemometry, or PIV

were considered as options for evaluating unsteadiness and turbulence in the plume.

However, each of these methods posed significant challenges. High frequency pressure

transducers available for use have a large cross section due to the required cooling system

for use in high temperature exhaust representative of advanced engine cycles. Hot-wire

anemometry was considered, but deemed to lack the necessary robustness for the high

temperatures and high flow velocities and flow rates36. Quantitative laser imaging

provides useful data on the velocity fields in the plume. However, depending on the

technique, this could require rig modifications to enable seeding the flow, expensive laser

components, and precise setup requirements. Accordingly, this diagnostic technique is

seen as a future capability to be developed. Although quantitative data is not currently a

capability, HPL does have experience with high-speed Schlieren imaging, and it was used

to obtain qualitative images of the nozzle and plume flow characteristics.

Therefore the selected source of quantitative plume data is a rake that measures

total temperature and total pressure. This concept was deemed to be sufficient to achieve

the stated goals by providing steady-state mixing characteristics of the plume, as well as

being relatively simple and low-cost. Additionally, a similar concept with total pressure

and total temperature measurements is used in the High Flow Jet Exit Rig (HFJER) at

NASA Glenn Research Center37. The major difference is that the system used on the

HFJER utilizes several rakes with 40 probes on each measurement rake, and is also

significantly larger, as shown in Figure 2-14. These rakes are mounted on a large

structure on rails and are able to be translated horizontally as well as axially. With the

large number of rakes and probes, the HFJER system is capable of achieving a

measurement resolution of 0.25-0.5 inches.

41

Figure 2-14 HFJER Plume Traversing System37

Due to cost and space restrictions, it was determined that the BANR facility

should attempt to satisfy a similar measurement resolution using a single rake with a

more capable traversing system. Accordingly, the rake design and was straightforward,

but the traversing system was a more significant challenge. The final design for the

BANR system is shown in Figure 2-15. It is significantly smaller than the equipment at

the HFJER, but is able to achieve comparable measurement resolution and is capable of

completely mapping all plumes from nozzles currently slated for testing on the BANR.

42

Figure 2-15 Plume Rake Traverse System

2.4.1. Rake Design

Due to the wide variety of nozzles anticipated to be tested using the BANR, it was

necessary to design a rake that can be useful for all anticipated nozzles. To do this, the

rake was designed with circular/rectangular transition nozzles in mind. Since one aim of

this particular nozzle type is the rapid spreading and mixing of the exhaust plume, it was

taken as the most extreme case of plume spreading. A nozzle of this type with a

rectangular exit plane of 6.32 in. by 1.58 in. (equivalent diameter of 4.45 in.) was used as

the design case because of the availability of previously gathered empirical data, as well

as the desire to conduct additional investigations in the near future using the BANR.

Although the BANR can potentially accommodate nozzles of approximately 10 inch exit

diameter, this particular transition nozzle was determined to be the most extreme case of

near-field plume spreading that is anticipated to be tested in the BANR facility.

43

Aside from runtime and scaling limitations, the facility also imposed a limitation

on how far downstream measurements could be taken. In general, the exit plane of the

nozzle will be approximately 4 feet from the overhead door aft of the BANR.

Accordingly, for the largest nozzle under consideration, this results in a maximum

downstream measurement point of approximately 4.5 nozzle diameters. Aft of this, it is

likely that flow measurement could be significantly affected by interference from the

Annex walls and weather conditions outside of the building.

2.4.1.1. Measurement Area

In Tang’s Ph.D. dissertation26, plume spreading data is presented for several

design variations of a transition nozzle. This data shows that at x/De = 5, the plume

width/De is approximately 3. Accordingly, for a nozzle De of 4.45 inches such as the one

under consideration for testing in the BANR, the plume width should be approximately

13.35 inches. If this spreading rate is extrapolated further downstream to the back wall of

the Annex, the plume width should be approximately 24 inches at 48 inches downstream.

As a result, the measurement area of the rake was designed to be 24 inches while the

overall length of the rake was determined to be 36 inches. This includes end supports

and mounting flanges, and will also ensure that any deviations of the plume from rig

centerline or major asymmetries do not result in hot exhaust impinging on any of the

structure or actuators used to support and translate the rake.

2.4.1.2. Construction

The final rake design is shown in Figure 2-16. The overall shape is rectangular

with a 30 degree wedge at the leading edge. The rake utilizes 20 probes (10 total

pressure and 10 total temperature) spaced evenly and alternately over the 24 inch

measurement area. As a result, there is approximately 1.25 inches between each probe,

and 2.4 inches between adjacent pressure or temperature measurements. The actual

probes are 0.125 inch diameter tubes with either pressure tubing or thermocouples

44

attached and fed out through the mounting flanges of the rake on each end. Some of the

pertinent rake dimensions are also shown in Figure 2-17.

Figure 2-16 Plume Rake Model

Figure 2-17 Pertinent Rake Dimensions in Inches

45

ATK GASL was chosen as the designer and manufacturer of the rake because of

their extensive experience designing and fabricating aerospace test hardware, including

similar rake designs. A general analysis was performed in-house to approximate the

aerodynamic loading on the rake for the purpose of designing the actuator system, but the

detailed finite element analysis was performed by ATK GASL. The harshest condition

was given to their analyst as a stagnation pressure of 147 psi, a stagnation temperature of

1300 F, and a Mach number of 2.15, which encompasses all the worst-case conditions

that are anticipated from BANR testing. For this condition, it was determined that a

stainless steel rake would not provide a sufficient safety margin. As a result, it was

decided to utilize Inconel 625 because of its high temperature capabilities and strength.

The analysis performed by ATK GASL shows stresses well below the allowable levels

for Inconel at the given temperatures38. The choice of Inconel as the material resulted in

significantly higher cost due to the material and machining costs, but was necessary to

ensure the ability to use this rake in all anticipated conditions.

2.4.2. Traversing System Design

The overall goal of the frame and actuator system is to allow for motion of the

plume rake in two dimensions (vertically and horizontally) during a test, and axially

between test runs. The motion during a test needs to be remote and automated, although

the axial motion can be done manually. The rake needs to be able to move approximately

36 inches in the stream-wise direction and 4 inches in the vertical direction in order to be

able to map the entire plume area for all anticipated nozzles. The actuators must be able

to control the rake in 0.25 inch increments at speeds on the order of 1 inch/second. The

frame needs to support all the required actuators to move the rake and any other

necessary instrumentation.

The system basically consists of three levels. The first level is a rectangular base

made of welded square tubing. This level is stationary in two dimensions, but has casters

on it to translate axially. Mounted to this level are screw-jack actuators used to provide

vertical motion. The translating ends of these screw-jacks are mounted to the middle

level, made of welded rectangular tubing. This level only moves vertically, but has a

46

linear actuator and several linear bearing rails mounted to it to provide horizontal motion.

The top level is a welded assembly of square tubing shaped similar to a soccer goal. This

assembly has the rake mounted to it and translates across the exhaust stream. This design

allows for precise, de-coupled positioning of the rake in three axes. These three main

levels are depicted in Figure 2-18.

Figure 2-18 Basic Frame of Plume Traverse System

2.4.2.1. Actuators

As mentioned previously, the actuators represented a significant design challenge

because of the range of motion required and the forces reacted by the plume rake. A

linear positioner is used to move the rake across the stream. This linear actuator is the

47

Parker HD-Series actuator with an overall travel distance of approximately 40 inches, and

is shown in Figure 2-19. This actuator must be able to react the drag force that the rake

experiences.

Figure 2-19 Parker HD-Series Actuator Used in Actuator System

To estimate the drag force, an analysis was performed assuming a range of NPRs

from 2 to 10. It was also assumed that the entire 36 inch length of the rake was exposed

to a uniform flow, and that the wave drag coefficient of the rake was that of a 30 degree

wedge at a Mach number of 2.2, corresponding to the actual design of the rake and the

most extreme case of aerodynamic loading. The drag coefficient (cd) depends on the drag

force ( 'D ), dynamic pressure (q1), and the chord length (c), where subscript 1 denotes

upstream of the oblique shock and 2 denotes downstream of the shock):

'

d1

Dc =

q c

Eq. 2.11

Using the geometry of the wedge and the definition of the dynamic pressure, the drag

coefficient can be found in terms of the wedge half-angle (θ), Mach number (M1), and

pressure ratio (p2/p1) across the oblique shock:

48

21 1

γq= p M

2

Eq. 2.12

2d 2

1 1

p4tanθc = -1

γM p

Eq. 2.13

Using the Mach number, the pressure ratio can be found using oblique shock tables, thus

determining the drag coefficient. For this particular case, the value was found to be

approximately 0.1939. These simplifications yielded a simplified but conservative

estimate of the total drag force on the rake due to the fact that the entire 36 inch rake

should never actually be exposed to the highest flow condition. According to this

analysis, the actuator should be able to operate under a maximum load of 225 pounds.

However, the actuator will also be assisted by two linear bearing rails to provide

additional support and stability in the axial direction, minimizing the loads and moments

on the actuator as much as possible. Figure 2-20 shows a free body diagram of the rake

depicting the loads imparted on it, and Figure 2-21 shows the drag force experienced by

the rake as a function of NPR.

Figure 2-20 Rake Free Body Diagram with Distributed Drag Load and Reaction Loads

49

2 3 4 5 6 7 8 9 1040

60

80

100

120

140

160

180

200

220

240Drag on Rake vs. Nozzle Pressure Ratio

NPR

Dra

g (lb

s)

Figure 2-21 Drag on Rake vs. Nozzle Pressure Ratio

The requirement on the vertical actuators is to lift the weight of the frame, rake,

and horizontal actuators, a weight of approximately 200 pounds. It also must be lifted

evenly, ensuring that the measurement plane of the rake remains reasonably orthogonal to

the floor. The actuators chosen are a set of jack-screws with motor, gear box, and

couplings to coordinate motion. This assembly is built by Nook Industries and is shown

in Figure 2-22. Although these jackscrews do not operate reliably with large side-loads

imparted on them, the screws of the assembly chosen for use are significantly more

robust than would otherwise be necessary to lift the load. This limitation is not

anticipated to be a problem, but any difficulties in providing the desired motion should be

immediately obvious. Figure 2-23shows the fully assembled system just downstream of

the nozzle exit plane.

50

Figure 2-22 Jackscrew Assembly for Providing Vertical Motion

Figure 2-23 Fully Assembled Plume Rake and Traverse System

Vertical

Actuators

Horizontal

Actuator

Plume

Rake

51

2.4.2.2. Control Architecture

The overall architecture for positioning the plume rake and gathering data is

depicted in Figure 2-24. It is essentially an open control loop, starting with a

programmed sequence and ending with the actual rake position being read into the data

acquisition system.

Figure 2-24 Plume Rake Control System Architecture

In order to move both the linear positioner and jack-screw assembly, motors and

motor drives are required to translate computer control signals into actual torque. As

noted, two motors are required in the system; one to move the rake across the stream and

one to power the jack-screw assembly. Conveniently, the loads and speeds required for

both actuators are of the same order. This similarity allows both common motors and

common drives to be used. The motors and drives were chosen from the product line of

Applied Motion Products. A high-torque stepper motor was determined to provide the

necessary torques, speeds, and smoothness of motion for both axes, and STAC-6 stepper

drives are used to power and control the motors. The STAC-6 is actually a standalone

52

drive that can either be programmed or provide real-time control. In this application, a

motion control program is written using the drive software, uploaded and stored on the

drive, and then executed through the use of a relay switch. The program will execute but

will not detect any motor stalls or other faults, although the drives are capable of this

when properly equipped with encoders and limit switches.

In order to reliably determine the actual position of the rake, position information

of some kind is necessary. The drives chosen for this application actually have the

capability to ensure that the motors they control achieve a very precise position through

various feedback and control loops. If for some reason they are unable to achieve the

precise position commanded, the drives remember and display any faults incurred while

executing their program. Although these advanced capabilities are useful in many

motion control applications, these features are not helpful for the rake positioning

application because of the difficulty that arises in synching rake data with position data in

the data acquisition system. The BANR facility is set up to record synchronous analog

data, which includes pressure measurements, temperature measurements, valve positions,

load cells, or any device that outputs 0 to 10 VDC. Unfortunately the STAC-6 drive does

not provide this type of output, nor do any other drives that were considered. As such, a

more direct way was necessary to measure the position of the rake that would be synched

in time with the actual temperature and pressure measurements taken by the rake in order

to correlate position with data and create a spatial map of exhaust characteristics.

To measure position, linear potentiometers are used. Linear potentiometers are

devices that have one translating end that effectively changes the resistance of the device

as it extends or retracts. With this method, the potentionmeter is included in a voltage

divider circuit with another resistor. A voltage is applied across the circuit, and the

resulting voltage drop across the potentionmeter can be used to determine its resistance.

This resistance value is then correlated to a position. Both the voltage across the entire

circuit and the voltage drop across the potentiometer are measured in order to achieve the

most accurate measurement possible. Overall, these linear potentiometers provide

several advantages. Perhaps the most beneficial is the fact that a linear potentiometer

does not rely on magnetic inductance such as a linear transducer. Also, the function of

53

the device allows for the translating ends to be placed in convenient locations on the

system.

2.4.3. Pressure Measurement Performance

Another important factor to consider when using the plume rake is its frequency

response. This response is governed principally by the physical characteristics of

associated tubing, ports, and transducers. Ideally, the transducer would be mounted flush

with a port of minimum depth, and the electronics would be directly connected to the

transducer itself, eliminating the need for any tubing. This type of transducer can be

useful for obtaining high frequency pressure data. Unfortunately, the high temperatures

of an exhaust plume require that a transducer used has a cooling system of some kind. In

turn, this cooling system results in greater size and cost. For the BANR facility, it was

determined that steady-state (on the order of 100 Hz) pressure measurements were

sufficient for evaluating mixer performance, allowing standard response transducers to be

used and not mounted in the actual exhaust plume. However, locating standard

transducers remotely from the actual measurement location inherently limits the

frequency response of the overall system. When there is an appreciable volume in the

port, transducer, and tubing, any pressure disturbances has to be transmitted through the

medium to the actual transducer. Accordingly, important variables are the length of

tubing, diameter of tubing, port volume, transducer volume, and sound speed of the

medium.

An analysis was performed for the current configuration of the BANR using the

following method adapted from several different references on measurement

techniques40,41. Two assumptions are made, namely that the fluid is assumed to behave

as an ideal gas, and the tubing is assumed to be rigid. The physical situation is depicted

in Figure 2-25. In the schematic, a transducer with internal dead volume, V, experiences

some measured pressure Pm. This transducer is connected by some tubing of length, L,

inner diameter Dt, and volume Vt. The system is driven by the pressure that is acting at

the entrance of the system, Pa. Initially, the pressure measured by the transducer, Pm is

assumed to be the same as Pa. Thereafter, the pressure acting on the system is a function

54

of time, and the pressure measured by the transducer will also be a function of time,

though it will lag the driving pressure according to the measuring system characteristics.

Figure 2-25 Schematic of Pressure Response Problem

Figure 2-26 is a free-body diagram of an element of fluid in the tubing. Time-dependent

pressure changes will act on the fluid to disturb and move it back and forth in the tubing.

The forces on the fluid are the driving pressure over the area of the tubing, a damping

force from fluid shear forces, and a compression-restoring force.

55

Figure 2-26 Free Body Diagram of the Fluid Element

In this case, ρ is the density of the fluid, x’’ is the acceleration of the fluid, x’ is the

velocity of the fluid, and μ is the viscosity of the fluid. The adiabatic bulk modulus of

elasticity (Em) can be found as follows (for the general case when the tube volume is

much less than the internal dead volume) from the ratio of specific heats (γ) and ambient

pressure:

m a

dpE =- =γP

dVV

Eq. 2.14

If the fluid element in the tube moves a certain distance, the resulting volume change

(dV) can be written in terms of the tubing volume (Dt) and distance moved (x):

(x/4)πDdV 2t Eq. 2.15

This change in volume results in a pressure excess (Pm) defined in terms of the adiabatic

bulk modulus of elasticity, tube diameter, distance moved, and system volume (V):

56

4Vx/DπEP 2tmm Eq. 2.16

Newton’s second law is then applied resulting in the following second order ordinary

differential equation (where L is the length of tubing):

2 2 4 2' ''a t m t tπP D π E D πD Lρ

-8πμL X - X= X4 16V 4

Eq. 2.17

Applying Eq. 2.16 to Eq. 2.17, the final form is attained:

'' 'm m m a2 4

m t m t

4LρV 128μLVP - P +P =P

πE D πE D

Eq. 2.18

This damped system second-order system can be modeled as follows:

KF(t)yyω

2y

ω

1

n2

n

Eq. 2.19

The resulting natural frequency (ωn) and the damping ratio (ζ) being defined as follows:

t mn

D πEω =

2 LρV

Eq. 2.20

ρπE

VL

D

32μζ

m3

t

Eq. 2.21

57

In cases where the fluid temperature is given, the acoustic wave speed of a perfect gas (a)

can be used:

cλRTga Eq. 2.22

Accordingly, Eq. 2.20 and Eq. 2.21 become:

2t

n

D πaω =

2 LV

Eq. 2.23

3t

32 VL=

a D

Eq. 2.24

From Eq. 2.14 and the ideal gas law, it is clear that Em and ρ both vary with pressure

changes, making the system non-linear. Because of this, it is assumed that the changes in

pressure are only small variations about an equilibrium pressure, resulting in changes that

are close to linear.

In cases where volume contained inside the tubing (Vt) is significantly larger than

the internal dead volume of the transducer, it has been shown that slightly different

formulas are appropriate due to the fact that the system behaves more like an operating

organ pipe42:

)4(V/V0.5L

aω

t

n

Eq. 2.25

)4(V/V0.5ρaD

L16ζ t2

t

Eq. 2.26

In either case, it is clear that a length that goes to zero will result in a natural frequency

that increases rapidly. This is handled by considering an equivalent length (Le) that

58

assumes that even though the transducer may be flush-mounted with L=0, there is still

some air present in the vicinity of the transducer diaphragm:

L

D

3π

81LL t

e Eq. 2.27

For under-damped, second-order systems, the response to a step function input can be

characterized additionally by the time the system takes to reach its first response peak

(Tp):

2n

pζ1ω

πT

Eq. 2.28

ns ωζ

4T

Eq. 2.29

With these expressions, the time it takes for the pressure measuring system to respond to

a step change in pressure can be found, as well as the time it takes for the measurement to

be damped to within ± 2% of the steady state value (Ts).

For the example of the plume rake configuration, the fluid properties were taken

to be that of ambient air for standard day values. The tubing was assumed to have a

0.085 inch inner diameter and an overall length of 20 feet. With these characteristics, the

volume contained in the tubing drives the frequency response as opposed to the volume

of the transducer or port having more influence. This calculation yields the result that the

system has a frequency response of approximately 10 Hz. As shown in Figure 2-27, the

length of the tubing has a significant effect on the frequency response. The discontinuity

at a tubing length of 13 feet is the point where the calculation methodology changes and

the volume in the tubing begins to have more influence than the volume of the port and

transducer.

59

0 2 4 6 8 10 12 14 16 18 200

20

40

60

80

100

120

140

160Frequency Response vs. Tubing Length

Tubing Length

Fre

quen

cy (

Hz)

Figure 2-27 Frequency Response vs. Length for 0.125” Tubing

The pressures measured by the rake are measured by an ESP Module that is

capable of taking measurements at up to 500 Hz. However, the maximum frequency that

can actually be physically experienced is on the order of 10 Hz as calculated. Therefore,

it is useless to measure at more than twice the estimated frequency, assuming it is

desirable to be able to perform a spectral analysis. If frequency content is unimportant,

then pressures can be scanned at whatever rate deemed appropriate for the rest of the data

being taken by the data acquisition system.

2.4.4. Motion Control Program

This frequency response also plays a role in determining how and at what speed

the plume rake can be traversed across the plume. If the area of interest is chosen to be a

set of discrete points, the loiter time at each point is determined by how many data points

60

are required to achieve a desired uncertainty level. If the system is only responding at 10

Hz, the loiter time required to get a set number of data points is clearly much greater than

if the system was responding at 50 Hz. However, the loiter time is also limited by the

overall amount of test time available at a given flow condition. Accordingly, the motion

control program for a particular test might need to execute simultaneous moves in both

directions, perhaps utilizing the maximum speed of the actuators. However, if more data

points are desired at only a few locations, the program can be more simple and the loiter

time can be longer. While a general motion profile can be developed to position the rake

in two axes, individual users will need to modify it in order to take data in the most

optimal manner.

2.4.5. Other Uses

Initially, this frame and motion system will be used to position the plume rake

designed specifically for evaluating mixer performance and plume characteristics.

However, the design is flexible in that with slight modifications, a wide variety of

instrumentation can be mounted and positioned in the exhaust plume of the rig. As

mentioned, the drives chosen have significant functionality beyond what is currently

being used. The motors are also sized such that they have the capability to handle more

loading. Accordingly, this system will be invaluable in both the short term and long term

for obtaining data on exhaust plumes.

2.5. Force Measurement

In order to assess the performance of a nozzle, the measurement of axial thrust is

critical. However, also of interest are forces in other directions, as well as moments

generated. To obtain this data, an integrated system consisting of multiple load cells was

incorporated into the BANR facility design. The thrust measurement system (TMS) in

use was designed, manufactured, and certified by Force Measurement Systems of

Fullerton, CA. It consists of a fixed ground frame with a flexurally supported live bed.

The live bed is designed to absorb the thrust loads with six data load strings that are

61

mounted between the live bed and ground frame. The load strings themselves consist of

a load cell with universal flexures on each end. In general, they measure axial thrust, side

thrust, vertical thrust, roll moment, yaw moment, and pitch moment. In addition to being

able measure six loads, the system is also able to perform in situ calibration on five of

these loads, lacking only the ability to calibrate the pitch moment. This calibration is

done through the use of high-precision load cells mounted in tension with universal

flexures. Screw jacks are then used to apply known loads to the calibration load cells.

Overall performance specifications are presented in

Table 2-1. In general, the thrust measurement system is designed to measure

axial thrust with an accuracy of 0.25% of full scale, resulting in a measurement that is

accurate to within 7.5 pounds. At full scale of 3000 pounds of axial thrust, the total

system deflection is approximately 0.05 inches. This minimal deflection will result in

better alignment and accuracy across the range of thrusts being measured.

Table 2-1 TMS Load Specifications

Axial Thrust 3000 lbs. Side Thrust 300 lbs.

Normal Thrust 300 lbs. Test Article Weight 500 lbs

Test Article C.G. 24 in. aft of live ring

2.5.1. Resolving Forces and Moments

The force measurement system incorporates multiple load cells at different

locations and orientations for measuring all loads imparted to the live bed. Accordingly,

62

each load cell measurement has to be resolved into orthogonal components to calculate

the resultant forces and moments in the axial, side, and vertical directions.

For all force and moment calculations the origin and coordinate axes are defined

as in Figure 2-28. When downstream looking forward at the interface of the live ring, the

positive z-direction points at the observer (downstream). The positive y-direction points

up and the positive x-direction points to the right (towards the center of the facility).

Note that axial thrust will accordingly be the opposite of the force in the positive z-

direction. Moments are calculated to be positive in the direction determined by the right-

hand rule. The XY-plane actually lies 2.425 inches upstream from the live ring interface,

which is the plane in which the load cells act. The load cells are configured to provide a

positive signal when in tension and a negative signal when in compression.

Figure 2-28 Live Ring with Sign Convention Used

2.5.1.1. Measurement Load Cells

The measurement capabilities of the thrust stand come from the 6 load cells

designated 0 through 5. These load cells react forces as depicted in Figure 2-29. Load

cells 0, 1, and 2 measure axial force (Z–direction) exclusively. Load cells 3 and 5

63

measure force in both the Y-direction and the X-direction, but not the Z-direction. Load

cell 4 measures force only in the X-direction.

Figure 2-29 Live Ring Free Body Diagram with Measurement Load Cells Depicted

Moments are also resolved using these six load cells. Since the point where the load is

being reacted is known for each load cell, moment arms can be determined. Load cells 0,

1, and 2 are used to resolve the moment about the X-axis. Load cells 1 and 2 are used to

resolve the moment about the Y-axis. Load cells 3, 4, and 5 are used to resolve the

moment about the Z-axis.

Using the given geometry of the live ring and load cells, equations for both the

forces and moments can be readily determined. In the moment equations, moment arms

are given in terms of absolute distance from the coordinate system origin along the

orthogonal axes and radii in the case of the Z-moment.

o ox 3 5 4F =F cos 60 +F cos(60 )-F Eq. 2.30

64

o oy 3 5F =-F sin(60 )+F sin(60 ) Eq. 2.31

z 0 1 2F =F +F +F Eq. 2.32

x 0 0 2 2 1 1M =F y -F y -F y Eq. 2.33

y 2 2 1 1M =F x -F x Eq. 2.34

z 3 3 4 4 5 5M =-F r -F r -F r Eq. 2.35

2.5.1.2. Calibration Load Cells

The ability to calibrate the BANR thrust stand in 5 axes comes from 5 calibration

load cells with screw-jack actuators that are used to apply various calibration loads to the

live ring. Load cells 6 and 7 apply force axially (in the Z-direction). Load cells 8 and 9

apply side force (in the X-direction). Load cell 10 applies force vertically (in the Y-

direction). These forces are depicted in Figure 2-30.

65

+Y

+X+ZF9

F6

F7

F8

F10

Figure 2-30 Free Body Diagram of Force Measurement System with Calibration Loads Depicted

Force can be calibrated in all three axes, but moment calibration is possible in only two

of the three moment axes. Calibration of moment about the X-axis is not possible with

this configuration because of the inability to create a pure couple about the X-axis. A

moment about the X-axis can be calculated using F10, but this will result in a force in the

Y-direction that cannot be countered. The moment about the Y-axis is calibrated using

load cells 6 and 7. The moment about the Z-axis is calibrated using load cells 8 and 9.

Similarly, the calibration load cells can be resolved to yield the resultant forces in

orthogonal directions:

x 8 9F =-F -F Eq. 2.36

y 10F =-F Eq. 2.37

66

z 6 7F =-F -F Eq. 2.38

x 10 10M =-F z Eq. 2.39

y 6 6 7 7 8 8 9 9M =F x -F x +F z +F z Eq. 2.40

z 8 8 9 9M =F y -F y Eq. 2.41

2.5.2. Load Cell and Data Acquisition Calibration

Calibration for forces and moments should be performed over the range of values

expected during a particular test. The maximum expected thrust should be used as a

baseline for axial thrust. Side and vertical forces should be calibrated to approximately

10% of this range. Likewise, moments should be scaled similarly to calibration loads.

First, the appropriate loads for each calibration load cell need to be determined.

Assuming a maximum expected axial thrust of 3000 pounds, the following loads should

be applied:

To calibrate load in X-axis: 8 9F =F =150 lbs.

To calibrate load in Y-axis: 10F =300 lbs.

To calibrate load in Z-axis: 6 7F =F =1500 lbs.

To calibrate moment in Y-axis: 6 7F =-F =1500 lbs.

To calibrate moment in Z-axis: 8 9F =-F =150 lbs.

After these calibrations are performed, a calibration constant for each load and moment

can be determined as follows:

calibrationt

measurement

Fk =

F

67

This calibration constant can then be applied to the measured thrust values from tests to

obtain the most accurate thrust and moment values for the particular rig configuration in

use.

Calibration procedures have also been performed on the signal conditioning for

each individual load cell to aid in providing the most accurate force measurement

possible. This calibration was performed with the signal conditioning system powered on

for 30 minutes and an empty stand (no hardware was mounted to the live ring). First, the

output from each load cell was measured at the Annex patch panel using a precision

voltmeter. Each load cell was then zeroed using the adjustment screws on its front face

to 0 V +/- 1 mV. After zeroing, the span was adjusted for each load cell. At the thrust

stand, each load cell cable was disconnected from its respective load cell and connected

to a precision voltage source that was used to supply 30 mV. The output of each load cell

signal conditioning unit was then measured at the data acquisition system’s patch panel

and the span was adjusted on each unit to yield an output of 10 V +/- 1 mV.

Scale factors were also calculated to accurately determine loads based on input

voltage to the data acquisition system. Each load cell was factory calibrated yielding a

load cell scale of approximately 3 V/mV at 2000 lbf. The excitation voltage of each load

cell was recorded and multiplied by the load cell scale and the signal conditioning gain

applied (10V/30mV = 333.333). This yielded notional data acquisition system input

voltages for full scale loads of 2000 lbf. A scale factor for each load cell was then

calculated by dividing the full scale load by the notional full scale voltage input, yielding

a scale factor of approximately 200 lbf/V.

To ensure the best performance of the load cells and associated data acquisition

system, these procedures should be performed periodically, but especially after major

changes to the rig or installed test article, or the data acquisition system.

2.5.3. Vertical Thrust Anomaly

During pressure calibrations and initial testing, a large vertical force was seen to

be exerted on the rig. After verifying that the force was not a result of faulty data

acquisition or reduction, further investigation revealed that this force increased fairly

68

linearly with increased pressure. Initially, several causes were suggested. First, it was

thought that the flow through the bypass horn was imparting enough momentum to

generate the force. However, it was determined that the flow would have to be traveling

at very high velocities to have such a large effect, and this was determined to be unlikely.

The pressurization of the bypass flexlines was also considered to be a possibility, but

again was considered to incapable of producing such a dramatic effect. Finally, the

pressurization of the bypass steerhorn was investigated. A micrometer was placed on one

of the vertical flange faces of the steerhorn. At a pressure of only 100 psi, this flange was

seen to deflect approximately 0.050”. The steerhorn and flange deflection are depicted in

Figure 2-31.

Figure 2-31 Bypass Steerhorn and its Deflection While Pressurized

As shown in Figure 2-32, this pressurization resulted in a vertical force nearly three times

that anticipated under maximum flow conditions.

69

10 20 30 40 50 60 70 80 90 100 110-1000

-800

-600

-400

-200

0

200Vertical Force vs. Pressure

Pressure (psi)

For

ce (

lbs)

Figure 2-32 Vertical Force Generated by Bypass Horn Pressurization

Accordingly, the pressurization of the steerhorn and the resulting deflection of the flanges

was identified as the source of the large vertical force.

To mitigate this, supports were constructed to support the vertical flanges of the

bypass steerhorn. Additionally, the flexlines were removed and reinstalled in an attempt

to eliminate any preloads being imparted by the flexlines. A close-up view of the

supported flange and flexline is shown in Figure 2-33.

70

Figure 2-33 Steerhorn Flange Supports

These efforts significantly reduced the vertical force transmitted to the rig, as shown in

Figure 2-34.

10 20 30 40 50 60 70 80 90 100-20

-18

-16

-14

-12

-10

-8

-6

-4

-2

0Vertical Force vs. Pressure

Pressure (psi)

For

ce (

lbs)

Figure 2-34 Vertical Force Transmitted to Rig After Supports Installed

71

The latest testing data shows that the vertical force generated now is less than 10% of the

axial thrust, as anticipated. With this solution implemented, the BANR facility is now

capable of measuring forces and moments at all anticipated test conditions.

72

CHAPTER 3. RESULTS

The data presented in this chapter was gathered over the course of several months

in the spring of 2009. First, over 70 tests, including both cold flows and hot-fires, were

performed on the Gulfstream plug nozzle shown in Figure 3-1. These tests were intended

to gather data relating to the internal nozzle flow, unsteadiness, qualitative plume shock

structure, and thrust performance. This group of tests was used to present the overall

conditions throughout the rig as well as interpret the overall performance of the plug

nozzle, and the results are presented here. Other data such as the results relating to

internal flow, unsteadiness, and qualitative plume shock structure can be found in John

Tapee’s thesis35, and are referenced occasionally in this chapter. An additional-fire test

was then performed to investigate the mixing present in the plug nozzle plume, utilizing

the plume rake and traversing system. This data is presented both to show a small

portion of quantitative mixing data for the plug nozzle, and to demonstrate the

functionality of the plume rake and frame system.

73

Figure 3-1 Gulfstream Plug Nozzle35

3.1. Facility Performance

A representative hot-fire test was chosen to illustrate the different aspects of

overall rig performance. These aspects include feed pressures and temperatures, mass

flow rates, internal charging station conditions as measured by the internal flow rakes,

and forces generated The test in question was intended to simulate the plug nozzle cruise

condition at nozzle pressure ratio (NPR) of approximately 6.2 and hot stream temperature

of 1000 degrees F. This is the highest pressure/flow condition that has been tested using

the BANR, and is close to the highest condition possible for the current rig configuration.

3.1.1. Feed Pressures, Temperatures, and Mass Flows

As described in Chapter 2, the air mass flows in the rig are set by critical orifice

plates (the locations of which are depicted in Figure 2-3 and Figure 2-4), and the fuel

74

flow is set by the pressure drop provided across the fuel injector. Accordingly, the

pressures and temperatures upstream of the orifices and fuel injector are the key

parameters for setting the flow conditions to the test article. Figure 3-2 shows a time

history of the pressure upstream of each orifice in the rig.

0 10 20 30 40 50 60 70 80 900

50

100

150

200

250

300

350

400

450

500

Time (s)

Pre

ssur

e (p

sia)

Bypass

Core

Vitiator

Figure 3-2 Time History of Feed Pressures

As annotated in Figure 3-2, a typical test consists of three phases. The first phase

consists of flow adjustments to achieve the desired air flow condition. These adjustments

are made by setting the desired rig pressures upstream of the orifice plates and allowing

the control valves to use their closed-loop control to open or close as needed. This ramp-

up period is somewhat affected by the upstream supply pressure and the set pressures, but

can also be changed by adjusting the control loop settings. The torch ignitor is then

utilized to ignite the combustor. The torch is operated for several seconds to ensure

ignition of the combustor, and then extinguished. The torch can be configured to operate

Ramp-Up

To Condition

On-Condition

Test Time

Longer

Shutdown To

Cool Rig

Ignition

Occurs

75

over a range of chamber pressures, although for most tests it is operated at approximately

120 psi chamber pressure, as shown in Figure 3-3.

0 10 20 30 40 50 60 70 80 900

20

40

60

80

100

120

140

160

180

Time (s)

Pre

ssur

e (p

sia)

Figure 3-3 Time History of Torch Ignitor Chamber Pressure

Once on condition, the actual test time can be anywhere from several seconds to several

minutes in duration, depending on the condition and the amount of data desired. Notable

in Figure 3-2 and Figure 3-3is the jump in pressure that occurs upon ignition. Finally, the

test concludes with a shutdown period, which for hot-fire tests is extended in order to

provide cooling air for internal rig components. On the suggestion of Adam Trebs, this

cool down period is extended long enough to reduce rig temperatures to approximately

100-200 degrees F. Although temperatures will never reach a level where the

components’ structural integrity will be significantly degraded, cooling the rig as

gradually and evenly as possible will hopefully reduce any deformation of parts or

Ignition

Occurs

Torch

Operation

Torch is off but

chamber is

backpressured from

combustor

76

fasteners. The utility of this practice will be verified in the future when the rig is

completely disassembled for the first time.

Figure 3-4 shows the important fuel system pressures.

0 10 20 30 40 50 60 70 80 900

200

400

600

800

1000

1200

Time (s)

Pre

ssur

e (p

sia)

Injector

Supply

Figure 3-4 Time History of Fuel System Pressures

Several aspects of the current fuel supply system are demonstrated in Figure 3-4. First,

the supply pressure decreases when actually flowing fuel due to the inability of the pump

to supply the high flow rate and maintain the set pressure. Additionally, another

significant pressure drop is evident across the fuel fire valve. In the current rig

configuration, these pressure drops have been difficult to accurately characterize,

resulting in the some difficulty in achieving desired fuel flows and combustion

temperatures. This problem will eventually be solved by the installation of a more

capable fuel pump and smaller control valve. The larger pump will supply any

conceivable fuel flow rate with no appreciable pressure drop under flow, and the control

Pump

Pressure Drop

When Flowing

Pressure Drop

Across Fuel

Fire Valve

77

valve will be able to use its entire range to provide the required pressure drop to achieve

the desired fuel flow.

Also important to accurately setting air flows is the temperature of the air being

supplied to the rig. Due to the blowdown process within the air storage tanks feeding the

rig, feed temperatures can vary significantly depending on the set pressures, supply tank

pressure, and ambient temperature. Shown in Figure 3-5 is the temperature of the air

being supplied to the rig. It provides a useful reference of the overall decrease in air

temperature over the course of a test.

0 10 20 30 40 50 60 70 80 9020

40

60

80

100

120

140

160

180

200

Time (s)

Tem

pera

ture

(F

)

Figure 3-5 Air Supply Temperature During A Typical Test

Aside from the period of electrical noise induced by the spark ignition of the torch

ignitor, the temperature starts at the ambient temperature in the Annex. As the test

proceeds and the flow is maintained, the temperature drops between 20 and 30 degrees to

a point below freezing. After the test, the temperature slowly begins to climb back to its

Spark

Energized for

Torch Ignition

Shutdown

78

starting value as the cold internal air is heated by the warmer pipe and external air. In

many cases, the rig itself frosts over, as shown in Figure 3-6. It shows the result of a hot-

fire test where the bypass stream tends to frost even when there is hot gas flowing

through the core, suggesting that the bypass stream is not heated significantly by the core

stream.

Figure 3-6 Frosted Shroud and Rig with Hot Plug

Mass flow data was useful for verifying that a given condition was achieved and

troubleshooting if it was not. Specifically, the data was used to better refine the

predictive code for setting conditions in the rig. Also, the mass flow data is used in the

following section to calculate the ideal thrust of the nozzle and compare it to the actual

measured thrust, and calculate the discharge coefficient of the nozzle. In addition to the

flowmeters, mass flows were calculated based on orifice areas and measured upstream

79

pressures and temperatures. Figure 3-7 compares the low-pass filtered measurements

from both air flowmeters, and the orifice flow rate calculations.

0 10 20 30 40 50 60 70 80 90-5

0

5

10

15

20

25

30

35

Time (s)

Mas

s F

low

(lb

m/s

)

Bypass Flowmeter

Core Flowmeter

Bypass Orifice Calculation

Core Orifice Calculation

Figure 3-7 Time History of Air Flow Rates from Turbine Flowmeters and Orifices

Both figures show excellent agreement and stability for the core stream with the

exception of the electrical noise-induced excursions associated with the torch ignition

voltage command between 23 and 28 seconds. However, there is up to 10% variation

between the bypass flowmeter and orifice calculation over the length of a hot-fire test.

This variation changes throughout the test, reaching a maximum of about 10% at the end

of the test. This effect is most likely a result of the design of the rig. The core spool sits

inside the bypass orifice plate, requiring a thin gap between the two surfaces to allow for

assembly and thermal growth. This area of this annulus is approximately 30% of the

overall orifice plate area under normal conditions. As the core spool heats up during a

80

test, it expands and eliminates part of this annulus. This effect can be modeled as

follows:

initialΔX=αX ΔT Eq. 3.1

where X is the circumference, α is the coefficient of thermal expansion, and T is

temperature. Using appropriate values for stainless steel, the dimensions of the core

spool, and an assumed temperature difference of 1000 degrees, it was seen that the core

spool could conceivably expand to completely eliminate the starting annulus. Therefore,

in any hot-fire test, the temperature increase will decrease the bypass orifice area to some

degree, in turn reducing the overall flow rate since the upstream pressure is held constant.

This scenario is depicted in Figure 3-8. This theory is further supported by the fact that

this effect is not seen during a cold flow test.

Figure 3-8 Core Spool Thermal Growth and Resulting Bypass Orifice Area Variation

81

3.1.2. Charging Station Conditions

The most important rig conditions to consider are those in the charging station, as

this component provides the flow that feeds directly to the test article. As noted in

Chapter 2, the charging station utilizes rakes with total pressure and total temperature

measurements to verify flow conditions entering the nozzle. Each rake has 5 total

temperature measurements, 5 total pressure measurements, and a static pressure

measurement. The measurements are taken at locations that are the theoretical centers of

annuli of equal areas. Each stream has four rakes, offset by 90 degrees to provide full

360 degree azimuthal coverage. The four quadrants housing the rakes are defined in

Figure 3-9. The conditions measured by these rakes are used to determine the overall

NPR and core stream temperature achieved during a given test.

Figure 3-9 Charging Station Quadrant Definition, Looking Upstream

Figure 3-10 through Figure 3-17 show the pressures achieved in both the bypass

rakes and core rakes, respectively.

QUADRANT

I

QUADRANT

II

QUADRANT

III

QUADRANT

IV

82

-10 0 10 20 30 40 50 60 70 80 9010

20

30

40

50

60

70

80

90

100

Time (s)

Pre

ssur

e (p

si)

Wall StaticCenterline

Second from Center

Middle

Second From WallWall

Figure 3-10 Time History of Bypass Rake Pressures, Quadrant I

-10 0 10 20 30 40 50 60 70 80 9010

20

30

40

50

60

70

80

90

100

Time (s)

Pre

ssur

e (p

si)

Wall StaticCenterline

Second from Center

Middle

Second From WallWall

Figure 3-11 Time History of Bypass Rake Pressures-Quadrant II

83

-10 0 10 20 30 40 50 60 70 80 9010

20

30

40

50

60

70

80

90

100

Time (s)

Pre

ssur

e (p

si)

Wall StaticCenterline

Second from Center

Middle

Second From WallWall

Figure 3-12 Time History of Bypass Rake Pressures-Quadrant III

-10 0 10 20 30 40 50 60 70 80 9010

20

30

40

50

60

70

80

90

100

Time (s)

Pre

ssur

e (p

si)

Wall StaticCenterline

Second from Center

Middle

Second From WallWall

Figure 3-13 Time History of Bypass Rake Pressures-Quadrant IV

Malfunctioning

Pressure Probe

84

-10 0 10 20 30 40 50 60 70 80 9010

20

30

40

50

60

70

80

90

100

Time (s)

Pre

ssur

e (p

si)

Wall StaticCenterline

Second from Center

Middle

Second From WallWall

Figure 3-14 Time History of Core Rake Pressures-Quadrant I

-10 0 10 20 30 40 50 60 70 80 9010

20

30

40

50

60

70

80

90

100

Time (s)

Pre

ssur

e (p

si)

Wall StaticCenterline

Second from Center

Middle

Second From WallWall

Figure 3-15 Time History of Core Rake Pressures-Quadrant II

85

-10 0 10 20 30 40 50 60 70 80 9010

20

30

40

50

60

70

80

90

100

Time (s)

Pre

ssur

e (p

si)

Wall StaticCenterline

Second from Center

Middle

Second From WallWall

Figure 3-16 Time History of Core Rake Pressures-Quadrant III

-10 0 10 20 30 40 50 60 70 80 9010

20

30

40

50

60

70

80

90

100

Time (s)

Pre

ssur

e (p

si)

Wall StaticCenterline

Second from Center

Middle

Second From WallWall

Figure 3-17 Time History of Core Rake Pressures-Quadrant IV

86

The figures both show the trends of a typical test as discussed before. Additionally, the

pressures vary radially by only 1 or 2%, demonstrating excellent uniformity across the

streams. Finally, the wall static pressure can be seen to be only slightly less than the total

pressure in the stream, indicating the low Mach number design of the charging station.

In addition to radial uniformity, excellent azimuthal uniformity is demonstrated in

Figure 3-18. The pressures in each rake were averaged and compared to each other,

representing the entire 360 degrees of the bypass and core streams. Although there is

some initial non-uniformity between streams, the pressures during the test are uniform to

within approximately 2%.

-10 0 10 20 30 40 50 60 70 80 9010

20

30

40

50

60

70

80

90

100

Time (s)

Tot

al P

ress

ure

(psi

a)

CR1

CR2CR3

CR4

BR1

BR2BR3

BR4

Figure 3-18 Time History of All Charging Station Rakes-CR1 is Core Rake Quadrant I

While the total pressures in both streams follow similar trends, the total

temperatures do not. For the bypass stream, total temperature decreases throughout the

test. For the core stream, total temperature rises sharply as combustion is initiated and

87

then continues to climb slowly over the length of the test. These trends in the bypass and

core streams are shown in Figure 3-19 and Figure 3-20.

0 10 20 30 40 50 60 70 80 90-50

0

50

100

150

200

Time (s)

Tem

pera

ture

(F

)

Centerline

Second from Center

MiddleSecond From Wall

Wall

Supply

Figure 3-19 Time History of Bypass Stream Total Temperature

The total rake temperatures in the bypass stream as well as the air supply temperature

start at approximately ambient temperature. The figure clearly shows an initial radial

variation, with the wall being approximately 10 degrees colder than the centerline. This

was due to a brief test being run previously which reduced the temperature of the outer

wall. As the air begins to flow, the temperature drops significantly at first. There is

significant electrical noise introduced from the energizing of the spark for the torch

ignitor. After ignition, the temperatures remain reasonably stable despite the overall drop

in supply temperature, owing largely to the heat added in the core spool and transferred to

the bypass stream. After the test, the flow rate is reduced, and the temperatures begin

climbing.

Torch Ignitor

Operation

88

0 10 20 30 40 50 60 70 80 900

100

200

300

400

500

600

700

800

900

Time (s)

Tem

pera

ture

(F

)

Centerline

Second from Center

MiddleSecond From Wall

Wall

Figure 3-20 Time History of Core Stream Total Temperature

The core rake temperatures shown in Figure 3-20 behave quite differently. They initially

start slightly above 100 degrees due to thermal soakback from the previously run test. As

the air begins to flow, the temperatures decrease significantly until the combustor is lit.

As the torch ignitor is lit and the combustor initiates, the temperature spikes until the

torch ignitor is extinguished. As the test proceeds, the temperatures rise gradually,

varying slightly in the radial direction with the center being the hottest. After the test, the

temperatures fall very sharply back to a level close to their initial values. If any of the

rake temperatures, orifice temperatures, combustor liner temperatures, or skin

temperatures were seen to be significantly more than approximately 150 degrees F,

additional air was provided after the test to aid in the cooling process. In addition to

radial uniformity, the flow in the core stream shows good azimuthal uniformity,

indicating that the flow conditioning screen stack works as intended to eliminate any

Torch Ignitor

Operation

89

variation in the conditions exiting the combustor. This can be seen in Figure 3-21, which

presents the middle probe of each core rake and their azimuthal locations.

0 10 20 30 40 50 60 70 80 900

100

200

300

400

500

600

700

800

900

1000

Time (s)

Tem

pera

ture

(F

)

45 degrees

135 degrees

225 degrees

315 degrees

Figure 3-21 Time History of Core Temperatures: Azimuthal Variation

These pressure and temperature measurements show that the flow being provided to the

test article is uniform in both the radial and azimuthal directions. It also shows that the

Mach number in both streams is low, as designed.

3.1.3. Forces

As stated previously in Chapter 2, the BANR’s force measurement system is

capable of accurately resolving approximately 3000 pounds of axial thrust and 300

pounds of vertical and side force in its current configuration. Figure 3-22 shows the

forces generated during a typical test. Notable is the significant jump in axial thrust upon

90

ignition of the combustor, as well as the lack of jump in the off-axial forces. This

suggests that the nozzle itself is not the main source of the off-axial forces measured.

The axial force is seen to be approximately 2300 pounds, approximately 75% of the stand

capacity. Accordingly, this particular test article could safely be tested up to an NPR of

approximately 8. Additionally, the side and vertical forces are within 10% of the axial

thrust, also fitting appropriately within the capabilities of the force measurement system.

0 10 20 30 40 50 60 70 80 90-500

0

500

1000

1500

2000

2500

Time (s)

For

ce (

lbs)

Side Force

Vertical Force

Axial Force

Figure 3-22 Time History of Forces

Also notable in the force data is the fact that the signal is relatively noisy, especially in

the axial direction during combustion. Figure 3-23 shows the pressure in the combustor,

suggesting that the majority of the noise in the force data comes from the actual

combustion. For the hot-fire test, the amplitude of the noise is approximately 8 psi.

Ignition of

Combustor

91

0 10 20 30 40 50 60 70 80 900

20

40

60

80

100

120

140

160

Time (s)

Pre

ssur

e (p

sia)

Figure 3-23 Time History of Combustor Pressure for Hot-Fire Test

Figure 3-24 shows the combustor pressure during a cold flow test. When combustion is

absent, the amplitude of the noise is significantly less, on the order of 1-2 psi.

92

0 10 20 30 40 50 60 70 8010

20

30

40

50

60

70

80

90

100

Time (s)

Pre

ssur

e (p

sia)

Figure 3-24 Time History of Combustor Pressure for Cold Flow Test

Fast Fourier transforms were also performed to determine the frequency content of the

data in an attempt to identify if any low-frequency structural modes were present. This

analysis was performed on each load cell to identify modes in the axial direction as well

as off-axial direction. Figure 3-25 shows the fast Fourier transform results for the three

axial load cells.

93

0 10 20 30 40 500

2000

4000

6000

8000

10000

Frequency (Hz)

Pow

er (

lbf2 /H

z)

0 10 20 30 40 500

2000

4000

6000

8000

10000

Frequency (Hz)

Pow

er (

lbf2 /H

z)

0 10 20 30 40 500

2000

4000

6000

8000

10000

Frequency (Hz)

Pow

er (

lbf2 /H

z)

Figure 3-25 Low Frequency Content of Axial Load Cells

Each axial load cell shows a noticeable peak between 18 and 20 Hz, indicating that the

stand is not particularly stiff in the axial direction. The results for the off-axial load cells

are shown in Figure 3-26. In the first two, there seem to be distinct peaks around 20 and

30 Hz, however the magnitude of these peaks is approximately an order of magnitude

less that the peaks seen in the axial load cells. As such, there does not seem to be any

major low-frequency modes present in the off-axial direction.

94

0 10 20 30 40 500

1000

2000

3000

4000

5000

Frequency (Hz)

Pow

er (

lbf2 /H

z)

0 10 20 30 40 500

1000

2000

3000

4000

5000

Frequency (Hz)

Pow

er (

lbf2 /H

z)

0 10 20 30 40 500

1000

2000

3000

4000

5000

Frequency (Hz)

Pow

er (

lbf2 /H

z)

Figure 3-26 Low Frequency Content of Off-Axial Load Cells

3.1.4. Correlation of Desired Conditions with Actual Conditions

The predictive code described in Chapter 2 has demonstrated its usefulness across

a wide variety of conditions, including varying tank supply pressures, ambient

temperatures, and desired NPR. In general, the set points generated by the code always

result in a test NPR within 10% of that desired. As noted, the prediction of fuel flow and

resulting core temperatures is the largest handicap of the system, but enough empirical

data has been gathered at this point to also achieve temperatures within 10% of the

desired condition.

95

Table 3-1 and Table 3-2 show comparisons between predicted conditions and

actual conditions for a representative cold flow test and the hot-fire test under

consideration, respectively. The major difference between cold flow and hot fire tests is

the inaccuracy in the fuel setting, which results in an error that is significantly more than

that in any other condition.

Table 3-1 Cold Flow Test Condition Comparison

Parameter Desired Achieved Difference Core NPR 4.5 4.38 2.7 %

Bypass NPR 4.5 4.55 1.1 % Bypass Ratio 3.0 3.0 0.0 % Core Airflow

(lbm/s) 9.1 8.75 3.8 %

Bypass Airflow (lbm/s)

27.3 26.1 4.4 %

Table 3-2 Hot Fire Test Condition Comparison

Parameter Desired Achieved Difference Core NPR 6.23 6.08 2.4 %

Bypass NPR 6.23 6.12 1.8 % Bypass Ratio 3.0 3.01 0.33 % Core Airflow

(lbm/s) 9.9 10.0 1.0 %

Bypass Airflow (lbm/s)

29.8 28.5 4.4 %

Jet Fuel (lbm/s) 0.166 0.135 18.7 % Average Core

Temperature (deg F) While On Condition

1000 880 12.0 %

As shown, both hot and cold test conditions can be achieved with reasonable accuracy.

Repeatability has also been seen to be reasonable. One thing that can be done to slightly

improve the matching ability is to predict set conditions in real time instead of populating

an entire test matrix beforehand. This allows the test operator to take into account

changes in ambient temperature and tank pressure from day to day, or throughout the

96

course of a single day. However, this is not necessary if accuracy to within a few percent

is sufficient. Additionally, accuracy better than a few percent is probably not realistic

due to the myriad of effects on the rig and air supply that cannot be accurately accounted

for.

Practically speaking, the two major factors in achieving desired conditions are the

variation in bypass air flow due to thermal expansion, and the actual behavior of the

control valves used to regulate the air flow rates. Both issues can be addressed by

modifying the operation of the control valves. First, instead of using pressure to set a

condition, the actual flow rate can be set, eliminating the variation of the bypass flow

during a hot-fire from thermal expansion. Also, the valves tend to perform most

effectively with intermediate supply pressures and pressure settings corresponding to

NPRs above 2. At NPRs lower than 2, experience has shown that setting the pressure to

overshoot slightly and ramp down to the desired condition improves the performance of

the valves. If this is not done, the valves tend to slowly “chase” the condition, never

actually reaching the desired downstream set pressures. This step helps to minimize the

amount of tuning needed for the valves’ control loops.

3.2. Nozzle Performance Results

The data presented in this section was gathered over the course of several test

days from January 28, 2009 to February 23, 2009. The overall test campaign consisted of

79 tests, 58 of which were successful tests at discrete conditions. These tests were further

divided into 21 hot-fire tests and 37 cold-flow tests. The reason for the high number of

tests is that test points were often repeated in order to adjust the Schlieren imaging

system, change camera frame rates, or modify the high-frequency data acquisition. Also,

there is a higher number of tests at low NPRs due to the interest in aerodynamic

performance around NPRs of 1.6 to 2.0. Regardless of the configuration of any other

hardware, force, temperature, pressure, and mass flow data was taken for every test.

Although the NPRs in each stream are slightly different for each test, the core and bypass

NPRs were mass-averaged using Eq. 2.10 to obtain a single value more useful for data

presentation.

97

3.2.1. Axial Thrust

The measured axial force generated as a function of NPR is shown in Figure 3-27.

1 2 3 4 5 6 70

500

1000

1500

2000

2500

NPR

Thr

ust

(lbs)

Cold

Hot

Figure 3-27 Axial Thrust vs. NPR for All Tests

As shown in the figure, the axial thrust climbs linearly with pressure ratio. Also notable

is the fact that the thrust data for the cold flow tests matches that from the hot-fire tests

for the same NPR, verifying that the temperature of the flow affects the thrust only by

varying the total pressure in the nozzle.

3.2.2. Off-Axial Forces

Although axial thrust is of primary interest, side and vertical forces are also

important for determining the overall performance of the rig and feed system. In terms of

rig performance, this off-axis thrust data will be examined more in the future to better

98

characterize the contributions of the rig and feed system to force measurement.

However, at this time, the data is used to ensure the rig is operating safely and there are

no unusual off-axial contributions from the test article. Figure 3-28 shows the off-axial

forces for all the tests conducted. In almost all cases, the forces are negative, meaning

that they exert a force downward and to the port side when looking upstream.

1 2 3 4 5 6 7-250

-200

-150

-100

-50

0

50

NPR

Thr

ust

(lbs)

Side Thrust - Cold

Vertical Thrust - ColdSide Thrust - Hot

Vertical Thrust - Hot

Figure 3-28 Side and Vertical Force vs. NPR for All Tests

There is a fair amount of scatter in all the data presented, probably indicating that a

hysteresis effect caused by the bypass flexlines does contribute significantly to the forces

measured. This hysteresis effect is also demonstrated in Figure 3-29, which shows the

off-axial forces as a percentage of axial thrust. This data also shows a great deal of

scatter, further suggesting that there is probably significant hysteresis present in the

bypass flexlines.

99

1 2 3 4 5 6 70

2

4

6

8

10

12

14

16

18

NPR

Per

cent

Side/Axial - Cold

Vertical/Axial - ColdSide/Axial - Hot

Vertical/Axial - Hot

Figure 3-29 Off-Axial Forces as a Percentage of Axial Thrust for All Tests

3.2.3. Efficiency and Discharge Coefficient Assuming Unmixed Streams

This particular plug nozzle was not designed specifically to generate significant

mixing. Accordingly, the performance of the nozzle was first analyzed assuming two

unmixed streams. The thrust efficiency, one of the key performance parameters

comparing the measured axial thrust to the ideal jet thrust, was used in addition to the

discharge coefficient. For the bypass stream, the ratio of specific heats was assumed to

always be 1.4 while the gas constant was assumed to be 53.3 (ft-lbf)/(lbm-R). For the hot

stream, CEA32 was run for various cases to determine the variation in the ratio of specific

heats and molecular weight. The variation in molecular weight was seen to be minimal,

but the variation in the ratio of specific heats was seen to be as much as 3% as shown in

Figure 3-30. Also shown is the theoretical variation in equivalence ratio with combustor

temperature, assuming a stoichiometric fuel-to-air ratio of 0.067843. To simplify the

100

reduction of the data, an average value of 1.348 was assumed for the ratio of specific

heats for the hot stream. This assumption was seen to only affect the calculation of thrust

efficiency on the order of 0.1%.

700 750 800 850 900 950 1000 1050 1100 1150 12000.12

0.14

0.16

0.18

0.2

0.22

0.24

0.26

Desired Adiabatic Flame Temperature (deg. F)

Inlet Temperature = -20 F

Inlet Temperature = 0 FInlet Temperature = 20 F

0.12 0.14 0.16 0.18 0.2 0.22 0.24 0.261.33

1.335

1.34

1.345

1.35

1.355

1.36

1.365Ratio of Specific Heats vs. Equivalence Ratio

Inlet Temperature = -20 F

Inlet Temperature = 0 FInlet Temperature = 20 F

Figure 3-30 Hot Stream Properties

101

First, the theoretical exit velocity (Ve) of each stream was calculated using the ratio of

specific heats (γ), gas constant (R), stream total temperature (Ttotal) and NPR44.

1γ-1 2γ

totale

2γRT 1V = 1-

γ-1 NPR

Eq. 3.2

The ideal thrust (Fideal) was then determined by mass weighting the exit velocities with

the respective mass flow rates ( corem and bypassm ) as measured by the turbine flowmeters,

which are able to take into account the area variation in the bypass stream.

ideal core,ideal bypass,ideal core e,core bypass e,bypassF = F +F =m V +m V Eq. 3.3

The axial thrust efficiency (ηaxial) was then defined as the measured axial thrust (Fz)

divided by the ideal thrust:

zaxial

ideal

Fη =

F

Eq. 3.4

This data is presented in Figure 3-31. Additionally, the vector sum (Ftotal) of all the

forces (Fx, Fy, Fz) was found and the total force efficiency (ηtotal) was calculated assuming

that side forces were a result of misalignment of the nozzle relative to the force

measurement system.

2 2 2total x y zF = F +F +F Eq. 3.5

totaltotal

ideal

Fη =

F

Eq. 3.6

102

This data is shown in Figure 3-34. Also, the misalignment angle (α) between the

resultant force and the axial direction was computed:

-1 z

total

Fα=cos

F

Eq. 3.7

1 2 3 4 5 6 70.8

0.85

0.9

0.95

1

1.05

1.1

1.15

1.2

NPR

Cold

Hot

Figure 3-31 Nozzle Thrust Efficiency vs. NPR for All Tests Assuming Unmixed Flow

The thrust efficiency data shows an interesting difference between the low NPR

cases (NPR < 3) and high NPR cases (NPR > 3). For the low NPR cases, the efficiency

tends to be clustered around 0.9, with a scatter between 5 and 10%. In general, the hot-

fire tests tend to be more efficient at these low NPRs. However, there appears to be a

non-linear jump in efficiency at approximately NPR = 3. According to Tim Conners of

Gulfstream45, this jump is a result of the internal flow of the nozzle becoming fully

103

supersonic. This can be seen qualitatively in the Schlieren images shown in Figure 3-32

and Figure 3-33. At NPR = 2.59, the primary shock still extends inside the shroud.

However, at NPR = 5.01, this shock is completely external to the nozzle. After this point,

the efficiencies remain fairly constant around 1.05 until the cruise condition where it is

slightly higher.

Figure 3-32 Schlieren Image of NPR = 2.59

Shroud

Boundary

Shocks

Still Inside

Shroud

104

Figure 3-33 Schlieren Image of NPR = 5.01

The total resultant force data shows a similar trend and its magnitude is almost

identical to the axial force data due to the fact that the off-axis forces are only a small

percentage of the axial forces.

Shocks

Completely

Outside

Shroud

105

1 2 3 4 5 6 70.8

0.85

0.9

0.95

1

1.05

1.1

1.15

1.2

NPR

Cold

Hot

Figure 3-34 Nozzle Resultant Force Efficiency vs. NPR for All Tests Assuming Unmixed Flow

Probably the most interesting and curious feature of the thrust efficiency is the

fact that it is greater than unity for many of the cases. There are several sources of error

that could be responsible for this. First, there is the potential for error in the calculation

of the ideal thrust. Eq. 3.2 shows a dependence on gas properties such as R (with

molecular weight included) and γ. These values are well characterized for the bypass

stream, but less so for the core stream when combustion is present. In general, these

contributions will be small because any error is under the square root in the equation.

The other major source of error in the actual data could be from the mass flow

measurements. A bias error of 5% in these measurements would directly correlate to a

5% error in the calculated efficiency. The measurements from the turbine flowmeters are

probably closer to within 1 or 2% accurate based on comparisons to orifice calculations,

106

but this correlation should be monitored closely in the future, as recalibration of the

flowmeters will be necessary at some point. The other sources of error could be from the

actual nozzle flow and rig phenomena. First, the assumption that the two streams remain

completely separated throughout the nozzle is not necessarily accurate. In general, this

assumption is probably not the main contributor due to the fact that both cold and hot

tests show efficiencies greater than unity, but could still have some small effect. This

issue is examined further in the next subsection and the section on plume rake data. A

significant source of error could be from the pressurization of the bypass flexlines. This

pressurization effect has been seen in the side and vertical force data to be as high as 10%

of the axial force, with a hysteresis effect of up to 100% in some cases. Accordingly, it is

not unreasonable to expect a similar contribution to the axial force. Although the

pressure calibrations that have been performed on the rig show only modest contributions

to the axial force from the flexlines, these calibrations were limited to 100 psi and do not

take into consideration any non-linear pressurization effects that might occur at higher

pressures. Finally, as this expression for theoretical thrust does not include any pressure

thrust, it is possible that the complex flow generated in the nozzle and along the plug

surface could be contributing to the axial thrust measured, leading to efficiency values

greater than unity.

Finally, the notional misalignment angle of the resultant force was seen to be

between 2 and 11 degrees as shown in Figure 3-35. A physical misalignment of the same

amount would be readily apparent when looking at the rig. No such misalignment has

been observed, suggesting that the off-axis forces are largely due to the asymmetric load

contributions of the bypass flexlines and the contributions from other rig components.

107

1 2 3 4 5 6 72

3

4

5

6

7

8

9

10

11

NPR

Cold

Hot

Figure 3-35 Notional Misalignment of Total Force

The other key nozzle performance parameter is the discharge coefficient. As

noted before, the assumption is made that the two nozzle streams remain unmixed

throughout the nozzle. These two streams adjust to a common static pressure at the throat

and are then expanded downstream of the throat along the plug surface and shroud. A

schematic of this scenario is shown in Figure 3-36.

108

Figure 3-36 Dual Streams Sharing a Common Throat and Static Pressure

Although the discharge coefficient is usually defined as the ratio of the actual mass flow

to the theoretical mass flow, the notional split between core and bypass flow areas at the

throat needed to calculate mass flow are not known a priori. Accordingly, it was

desirable to define the discharge coefficient in terms of properties that are well known.

The properties that are known are the total pressures and temperatures from the charging

station measurements, the mass flows from the turbine flowmeters, and the physical

throat area. In the data used, the total pressure in the bypass stream is always slightly

higher than the total pressure in the core. From this, it is assumed that the bypass stream

is choked at the throat. Knowing the Mach number of the bypass stream (M8bypass) is

unity at the throat, the mass flow parameter (MFP) can be calculated:

Station 8 

Charging Station Rakes

109

γ+1

2 γ-12

γ MMFP=

R γ-11+ M

2

Eq. 3.8

Knowing the mass flow parameter, the theoretical area of the bypass stream (A8bypass) can

be calculated:

•

bypass t,bypass

bypassbypass t,bypass

m TA8 =

MFP(M8 )P

Eq. 3.9

Also, the static pressure at the throat can be calculated using an isentropic relationship

knowing the Mach number of the bypass stream is unity:

t,bypass8 γ

γ-1

PP =

γ+1

2

Eq. 3.10

The Mach number of the core stream (M8core) can then be calculated, as well as the mass

flow parameter, using Eq. 3.11 as before:

γ

γ-1t,core 2core

8

P γ-1= 1+ M8

P 2

Eq. 3.11

The theoretical area of the core stream (A8core) can be calculated:

•

core t,core

corecore t,core

m TA8 =

MFP(M8 )P

Eq. 3.12

Finally, the two theoretical stream areas can be added to get a total theoretical throat area

(Atheoretical):

110

theoretical bypass coreA =A8 +A8 Eq. 3.13

The nozzle discharge coefficient (Cd) is then defined:

theoreticald

throat

AC =

A

Eq. 3.14

This discharge coefficient is shown in Figure 3-37.

1 2 3 4 5 6 70.8

0.85

0.9

0.95

1

1.05

NPR

Cd

Cold

Hot

Figure 3-37 Discharge Coefficient vs. NPR Assuming Unmixed Streams

A notable feature of Figure 3-37 is the fact that the there appears to be different trends for

the cold flow and hot-fire data. As would be expected, this suggests that there is some

effect in the hot-fire tests that is causing the discharge coefficient to differ from cold

111

flows, in some cases even making it higher. Additionally, the discharge coefficient data

shows a conflicting trend with the nozzle efficiencies. For both hot and cold tests, the

discharge coefficient decreases as the NPR increases. This indicates more loss occurring

in the nozzle at the higher NPRs than at the lower, which is the opposite of the trends

shown in the nozzle thrust efficiencies in Figure 3-31 and Figure 3-34. This

contradiction seems to further indicate that error is being introduced from either some

type of geometric variation, the pressurization of the bypass flexlines, or both.

3.2.4. Efficiency and Discharge Coefficient Assuming Perfectly Mixed Streams

For comparison with the unmixed analysis, the thrust efficiency and discharge

coefficient were also calculated assuming perfectly mixed flow. Although the total

pressure (Pt) varies slightly between streams, it can be mass weighted to get an average

value ( tP ).

core t,core bypass t,bypasst

core bypass

m P +m PP =

m +m

Eq. 3.15

The geometric throat area (Athroat) is also fixed. Accordingly, the thrust efficiency

(defined as ηm to differentiate between the previous formulation):

zm

total throat

F=

P A

Eq. 3.16

Also, the theoretical thrust efficiency can be calculated using γ (mass averaged according

to flowmeter data), NPR, exit pressure (Pe), ambient pressure (Pa), total pressure, exit

area (Ae) and throat area:

112

γ+1 γ-12 γ-1 γ

e a em

t throat

P -P A2γ 2 1= 1-

γ-1 γ+1 NPR P A

Eq. 3.17

In Eq. 3.10, the exit area is not defined for the plug nozzle. Additionally, ideal operation

of a plug nozzle will result in no pressure mismatch between exit pressure and ambient

pressure. Accordingly, this term is negated in the calculation. The actual thrust

efficiency data is compared to the theoretical data in Figure 3-38.

1 2 3 4 5 6 7

0.4

0.5

0.6

0.7

0.8

0.9

1

1.1

1.2

1.3

NPR

f

Cold-Measured

Hot-Measured

Cold-Theoretical

Hot-Theoretical

Figure 3-38 Thrust Efficiency vs. NPR for All Tests Assuming Perfectly Mixed Flow

The data collapses onto a single line for both hot and cold tests, as would be expected

from the axial thrust data presented earlier. Additionally, the data shows the expected

trend of nozzle performance as a function of NPR. At low NPR, the thrust coefficient is

low, indicating less efficient nozzle operation due to the shock structures present. As the

113

NPR increases towards the design point of approximately 6.2, the thrust coefficient

continues to increase until it gradually levels off at its maximum value. There is also a

difference between the theoretical and actual values. This indicates that there is actually

a pressure mismatch to some degree, leading to a contribution from the pressure area

term. However, the magnitude of the theoretical data indicate that the calculations made

using measured thrust and total pressure are reasonable. The ratio of the actual thrust

coefficient over the theoretical thrust coefficient is shown in Figure 3-39.

1 2 3 4 5 6 70.78

0.8

0.82

0.84

0.86

0.88

0.9

0.92

0.94

0.96

NPR

Cf/

Cf th

eory

Cold

Hot

Figure 3-39 Ratio of Measured Thrust Coefficient to Theoretical Thrust Coefficient

As expected, Figure 3-39 shows a general trend towards more efficient nozzle operation

at the higher NPRs. However, the magnitude of the ratio also indicates that the nozzle

operation is different from that of an ideal plug nozzle at low NPRs. It also could suggest

that there is some geometrical variation taking place in the vicinity of the throat due to

the increase in temperature and pressure. Overall though, it appears to be more

114

appropriate to represent the efficiency of the nozzle by using the perfectly mixed

assumption.

The discharge coefficient found assuming unmixed streams can then be compared

to the discharge coefficient assuming perfectly mixed flow according to the measured

bypass ratio. This is defined as the measured mass flow (•

measuredm ) divided by the ideal

mass flow (•

idealm ):

•

measured

d •

ideal

mC =

m

Eq. 3.18

The ideal mass flow can be defined in terms of gas properties, total pressure, total

temperature, and throat area. For the calculation, the ratio of specific heats, total

pressure, and total temperature are mass averaged according to the measured bypass ratio.

γ+1

• 2 γ-1t

ideal throat

t

Pγ 2m = A

R γ+1 T

Eq. 3.19

These discharge coefficients for all the tests are shown in Figure 3-40. Like the discharge

coefficient for the unmixed analysis, the trends for hot and cold tests are different, again

indicating that additional effects are present in hot-fire tests. However, they are also

somewhat conflicting in that they likewise indicate less efficient operation at higher

NPRs.

115

1 2 3 4 5 6 70.84

0.86

0.88

0.9

0.92

0.94

0.96

0.98

1

1.02

1.04

NPR

Cd

Cold

Hot

Figure 3-40 Discharge Coefficients

3.3. Plume Rake Data

The plume rake and traverse system proved to be the most challenging aspect of

the instrumentation, both in terms of design as well as its implementation. Making all the

required connections for remotely controlling the motors and drives, as well as the limit

and home switches, was simple and straightforward. The application used for

programming the drives with motion control routines proved to be simple and user-

friendly (although a machine loaded with Windows Vista refused to communicate with

the drives). As such, the motion control aspect of the plume rake system was ready in

relatively short order. However, when sample data was taken, the plume temperature

data was seen to be corrupted with large amounts of noise. Although some noise was

expected with the use of the electric motors, the actual magnitude of the noise was

somewhat disconcerting.

116

3.3.1. Noise Concerns

Several tests were performed to attempt to better identify the specific source of

the noise. First, data was gathered with the motors actually in motion, as well as in an

idle state. Figure 3-41 demonstrates the effect of powering the stepper drives and motors.

Initially, both drives are on, leading to a significant zero offset and a significant noise

level. When one of the drives is turned off, the offset decreases by approximately 50%,

suggesting that the effects of each of the two drives and motors is somewhat equal.

Finally, when both drives are powered down, the abnormal noise level and offset are

eliminated.

0 1000 2000 3000 4000 5000 6000 700070

80

90

100

110

120

130

140

150

160

170

Sample

Tem

pera

ture

(F

)

Figure 3-41 Electrical Noise Induced by Stepper Motors and Drives on a Plume Temperature Measurement

After speaking with the manufacturer of the motors and drives, a specific remedy

was suggested involving the use of ferrite toroids46. These magnetic “doughnuts,” seen

Both

Drives OFF

Horizontal

Drive OFF

Both

Drives ON

117

in Figure 3-42, are intended to mitigate electro-magnetic interference radiated from motor

power leads. The power leads are first twisted (positive with negative), and then looped

around one or more ferrites between 3 and 5 times. This essentially constrains the

radiated magnetic field within the loop, ideally eliminating the radiated noise.

Figure 3-42 Magnetic Ferrite and Installation

The installation of these ferrites was seen to improve the quality of temperature data to a

small degree. However, the major driver in the amount of noise was still seen to be the

amount of current being used to power the stepper motors. As such, the current used was

reduced as much as possible to minimize this effect.

Although there is most likely room for further incremental improvements in terms

of noise mitigation, the current levels were deemed acceptable for demonstrating the

functionality of the plume rake traverse system, as well as gaining insight into the mixing

processes in the plug nozzle.

3.3.2. Plume Maps

A single test of 120 seconds was used to gather data to create partial maps of the total

temperature and total pressure at the exit plane of the nozzle. These maps are shown in

Figure 3-43 and Figure 3-45.

Unfortunately, the actuators did not work precisely as desired due to the lowered

current levels used to power the step motors. It also appears that possibly one or two of

118

the total pressure probes malfunctioned. However, the partial maps do show some

interesting features of the exhaust flow. The temperature map shows the anticipated

temperature distribution, with hot flow in the center and cold flow on the outside.

However, it also seems to indicate that there is some mixing present.

-2 -1 0 1 2 3 4 5 6-4

-3

-2

-1

0

1

2

3

4

X (in.)

Y (

in.)

-100

0

100

200

300

400

500

600

700

800

900

1000

Figure 3-43 2-D Total Temperature Map of Nozzle Exit Plane

When compared to the charging station conditions shown in, it is seen that the hottest

temperature observed by the rake is less than that of the core stream. Although the

maximum temperature seen in the core stream is approximately 1100 degrees F, the

maximum temperature measured by the plume rake is less than 1000, indicating that

mixing between streams is most likely taking place to some degree.

Nozzle

Shroud

Plug

Throat

o F

119

0 20 40 60 80 100 120 140 160 180 2000

200

400

600

800

1000

1200

Time (s)

Tem

pera

ture

(F

)

45 degrees

135 degrees225 degrees

315 degrees

Figure 3-44 Total Temperature in Core Charging Station During Plume Mapping

The pressure map suggests that the total pressure is largely axisymmetric at the exit plane

at this condition. Also, the total pressure distribution from centerline to the shroud

diameter is not clearly stratified, suggesting there is significant shock generated

instability at this condition. Although Gulfstream does suggest that asymmetries have

appeared in CFD results at low NPR conditions such as that tested due to instabilities47,

no significant asymmetries were observed at the exit plane at this condition.

120

-6 -4 -2 0 2 4 6-6

-4

-2

0

2

4

6

X (in.)

Y (

in.)

15

20

25

30

35

Figure 3-45 2-D Total Pressure Map of Nozzle Exit Plane

Another feature to note is the magnitude of the total pressures observed by the rake. At

all locations, they appear to be somewhat lower than the total pressure in the charging

station, shown in Figure 3-46. This is likely due to the shock structures that are present

inside the shroud.

Nozzle

Shroud

Plug

Throat

PSIA

121

0 20 40 60 80 100 120 140 160 180 20014

16

18

20

22

24

26

28

30

Time (s)

Pre

ssur

e (p

sia)

Core Total Pressure

Bypass Total Pressure

Figure 3-46 Charging Station Total Pressures During Plume Mapping

The data gathered and displayed in these maps seems to indicate that the two streams in

the nozzle do mix to some degree. The plume rake and traverse system has also

demonstrated functionality, although it is clear that significant work is involved in

completely mapping an exhaust plume and its development.

122

CHAPTER 4. CONCLUSIONS AND RECOMMENDATIONS FOR FUTURE WORK

4.1. Concluding Remarks on Plug Nozzle Investigation

Several things can be said about the plug nozzle investigation undertaken in the

beginning of 2009. First, a significant gap in existing plug nozzle data has begun to be

addressed. A great deal of information has been gathered about a turbofan nozzle

intended for actual flight, and at NPRs and scales representative of an SSBJ. This data

includes quantitative information about the internal flow, forces generated, and a small

amount of plume mixing data. The newly gathered data also includes quantitative

information about plume shock structures in the form of high-speed Schlieren footage.

This data essentially fits in at the lower end of the plug nozzle database. Significant work

by NASA and DLR cover rocket pressure ratios and nozzle sizes. The research by

NASA under the Supersonic Transport Program23 addresses large airbreathing

applications. The current investigation continues stepping down to airbreathing pressure

ratios and more practical business jet sizes.

Although the force data itself was not critical for Gulfstream’s investigation, it is

intriguing in that for two different formulations of thrust efficiency, the nozzle

demonstrates excellent performance at the design condition where the NPR is

approximately 6.2. Additionally, the nozzle maintains this relatively high performance

until the NPR is decreased to the range where the internal nozzle flow becomes unsteady

and not fully supersonic. Although there is a significant drop in nozzle performance at

these low pressure ratios, this deficiency could be offset in the overall trade study by the

easier aircraft integration enabled by the plug nozzle. The plume mixing data is also

interesting in that aside from demonstrating the utility of the plume rake system, it also

was useful in evaluating the major assumptions used to analyze this nozzle, namely that

the two streams either remain unmixed, or perfectly mixed.

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Aside from the huge amount of data collected, the experience gained throughout

the test campaign is invaluable to the future success of the BANR. Operational

procedures were written, implemented, and refined to allow for smooth and efficient

operation of the rig at all desired test conditions. Experience was gained using high-

frequency instrumentation and its associated data acquisition. A wealth of knowledge

relating to Schlieren imaging and high-speed camera use was obtained as well, albeit

painfully at times. Finally, the overall behavior of the rig under flow was characterized

across a wide operating envelope. Most importantly, the BANR can now be operated to

gather various types of data at any flow condition that the hardware can physically

support.

4.2. Future Work

Although over 60 successful tests were conducted during the plug nozzle test

campaign constituting a huge volume of data, a significant amount of work can still be

done to better characterize the nozzle, and more importantly the BANR itself. With

regards to the plug nozzle, more data points in the region between NPRs of 3 and 6 would

help to better understand and verify CFD results on the performance of the nozzle in the

climb portion of its flight regime. Also, initial plume mixing data has been gathered to

create a partial plume map at the exit plane. However, this map only shows part of the

mixing of the plume at the exit plane and at one test condition. If desired, significant

work could be done to fully map the plume generated by the plug nozzle, including

gathering data at multiple axial stations downstream of the exit plane, as well as at

different test conditions. More experience with operating the traverse system will also be

invaluable in the future. In the near term, the rake will be used to measure plume

development in an IR reduction nozzle. However, interest has been expressed in testing

mixer-ejector nozzles in general, as well as various specialized mixer designs.

The most important area of work in terms of long term operation of the rig is the

characterization of the rig as it relates to measuring forces. As shown in the data, there

seems to be a significant amount of non-repeatability, as well as force levels greater than

expected given the test conditions. It is suspected that the bypass flexlines are the

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primary culprits. For future investigations where accurate force measurement is critical,

it will be crucial to know what forces result from the pressurization of these flexlines. As

important will be the repeatability of this pressurization effect. The initial data presented

here does not seem to have good repeatability, however a more extensive and meticulous

study of this effect could prove otherwise. Specifically, future work should focus on

determining what effect the bypass flexlines have on the force measured in all three

directions and the repeatability of this effect. This will give insight into the overall

accuracy of force measurements made on the rig in its current configuration, and whether

or not significant modifications are necessary to achieve a desired accuracy level.

Fortunately, the tools required to resolve this issue have already been procured or

are in place. A nozzle with an ASME standard contour has already been built and tested

briefly before the plug nozzle. Because this nozzle generates a known amount of purely

axial force given a set of conditions, it will be possible to identify what forces are being

generated by the rig itself as opposed to the nozzle. In addition, the calibration load cells

on the force measurement system will be useful for applying known loads to the thrust

stand, even during rig operation.

4.3. Final Thoughts and Lessons Learned

In general, the amount of time and effort required to make the BANR functional

was far greater than anticipated. Fortunately, there were no major part modifications or

redesigns required. However, several lessons were well learned over the course of 18

months. First, never assume a component is correctly assembled unless you want to have

rapid unplanned disassembly of that particular component. Second, establish design

goals and then work to meet them, otherwise the design effort will take much longer and

creep continuously. Third, some things are extremely difficult to procure, namely metal

e-seals and fuel injectors. Finally, things will almost always take 50% longer than

anticipated in the long run no matter how much planning is done. Do not underestimate

the time it will take to do tedious things such as wiring and plumbing.

Specifically, several things affected the fabrication, assembly, and operational

timetable of the BANR. First, although all the parts and fasteners were carefully

125

designed and specified, the reality of actual assembly ended up requiring a certain degree

of customization and adaptation, especially with fasteners. For the BANR, the assembly

process was inherently a serial operation. As a result, issues with seals and fasteners

were encountered one at a time and could not be dealt with simultaneously.

Unfortunately, each time a new fastener was needed, no less than a day was added to the

assembly process. In the case of some of the advanced seals used in the core stream,

several weeks were added to either re-machine a part or get new seals custom made.

Although a day or two is usually not critical, these slips tend to add up, especially with

weekends and holidays added in. The other major detractor from the schedule was the

sheer number of man hours required to build a facility from scratch. Over 100 wires and

connectors had to be made, connected, run, terminated, and checked for functionality.

Aside from wires, nearly 100 separate pneumatic tubes had to be cut and connected. The

most tedious instrumentation challenge was making connections for the metal-sheathed

thermocouples used in the charging station rakes and hot rig parts. Although a wealth of

experience was gained in doing this type of instrumentation, it came at a high price in

man hours. Finally, the lack of experience with the rig and the simple fact that all the

parts had never been fully assembled necessitated a slow and deliberate assembly and

shakedown process.

At this point, all assembly issues and most operational issues have hopefully been

identified. Barring any major damage or changes to rig components from operation,

disassembly and reassembly should be relatively routine. However, to aid in reassembly,

fasteners must be saved and kept with their respective components. A large and

organized fastener inventory, including all fasteners called for (as well as sizes slightly

larger and smaller) would also aid in the rig assembly. Overall, the most valuable tool for

avoiding a slipping schedule is efficient use of manpower. When there is only a single

specialized task that needs to be done (such as design work, coding, data reduction, etc),

there is no reason to devote more than one or two people. However, when a large amount

of work needs to be completed that is either non-specialized (wiring, plumbing, wrench-

turning, etc) or a combination of general and specialized tasks, sheer manpower is

invaluable. This effect was most evident in the winter of 2008/2009. Previously, there

126

had never been more than two people working full-time on making the rig operational.

However, a third and sometimes fourth person were added in the winter and the progress

leading up to the Gulfstream test campaign was remarkable.

Despite the challenges faced, the Gulfstream test campaign was extremely

successful. Although a working fuel injector was not installed until the day the campaign

started, a huge amount of data was collected efficiently and professionally, even at test

conditions initially thought to be unattainable. The test team encountered and overcame

challenges to gain the knowledge and experience to operate the BANR safely and

effectively. Hopefully this test campaign is only the first in what will be a long and

storied lifetime for the BANR.

LIST OF REFERENCES

127

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128

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17. Conners, Tim, E-mail Correspondence, Gulfstream Aircraft Corporation, April 3,

2008.

18. Hagemann, G., Immich, H., Nguyen, T. V. and Dumnov, G. E., “Advanced Rocket

Nozzles,” Journal of Propulsion and Power, pp620-634, Vol. 14, No. 5, September-

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19. Lee, C. C., “Fortran Programs for Plug-Nozzle Design,” NASA TN/R-41, NASA,

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21. Hagemann, G., Immich, H. and Terhardt M., “Flow Phenomena in Advanced Rocket

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22. Tomita, T., Tamura H. and Mamoru T., “An Experimental Evaluation of Plug Nozzle

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23. Stitt, L. E., “Exhaust Nozzles for Propulsion Systems With Emphasis on Supersonic

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27. Louis, J. F., Letty, R. P. and Patel, J. R., “A Systematic Study of Supersonic Jet

Noise,” AIAA Paper 72-641, 1972.

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30. Tam, C. K. W., “Supersonic Jet Noise,” Annual Review of Fluid Mechanics, pp17-43,

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35. Tapee, John, “Experimental Aerodynamic Analysis of a Plug Nozzle for Supersonic

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131

37. Soeder, R., Wnuk, S. and Loew, R., “Aero-Acoustic Propulsion Laboratory Nozzle

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APPENDICES

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Appendix A. Rig Setting Code

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% % Program title: BANR_Flow_Analysis.m % Author: Alex Sandroni % Last Modified: 26 January 2009 % Modified By: Alex Sandroni %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %This program takes a given test condition and calculates the various set points and expected flow conditions throughout the BANR facility. The final outputs are required flowrates, orifice set pressures, and control valve set points. %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% clear all close all clc format compact format short g %% Test Day Conditions P_amb = 14.7; %Atmospheric pressure in psi T_tank_F = 30; %Ambient temperature in F, taken from www.weather.com T_tank = T_tank_F + 460; %Air tank temperature in Rankine T_tank_total = T_tank; % Air tank stagnation temperature in R assuming %stagnant air in tank P_tank = 2000; %Starting air tank pressure in psia %% Flow Areas and Coefficients Cd_orifice = 0.84; %Discharge coefficient of long orifices Cd_nozzle = 0.97; % Discharge coefficient for nozzle A_vitiator_orifice = 5.7424/144; %Core Orifice area in ft^2 A_core_orifice = (7*pi*(0.4375^2)/4)/144; %Core orifice area in ft^2 %Formerly 0.328 holes on core orifice A_bypass_orifice = 1.15*(20*pi*(.516^2)/4)/144; %Bypass Orifice area, ft^2 A_ASME = ((pi*5.884^2)/4)/144; % ASME Calibration Nozzle Area in ft^2 A_plug_hot = 21.5/144; %Core flow area at test article interface for plug A_plug_bypass = 41.8/144; %Bypass area at test article interface for plug A_plug_throat = 24.2/144; %Throat area for plug in ft^2 A_core_screen = pi*((6.968/2)^2 - (1.506/2)^2); %core screen area in in^2 A_bypass_screen = pi*((12.864/2)^2 - (7.839/2)^2); %bypass screen area,in^2 %% Properties of Air and Jet Fuel SL = 0.820; % Specific gravity of jet fuel

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gamma = 1.4; % Ratio of specific heats for combustion gas gc = 32.174; %in (lbm-ft)/(lbf-sec^2) R = 1716/gc; %specific gas constant in (ft-lbf)/(lbm-R) n = 1.04; % polytropic constant (somewhere between isothermal(1) and %adiabatic(1.4) blowdown --> based on empirical observations) %% Desired Test Conditions Type = 2; %Cold Flow = 1, Hot Flow = 2; NPR = 3; % Desired overall nozzle pressure ratio (from nozzle throat) BPR = 3; % Desired bypass ratio (bypass flow/core flow) T_core_F = 700; % Desired core flow temperature in degrees F T_core = T_core_F + 460; % Conversion of desired core temperature to deg R %% Calculation of Mass Flows, Pressures and Temperatures P_total = NPR*P_amb; %Total pressure at throat in psi T_throat = T_tank*(P_total/P_tank)^((n-1)/n); %Temperature in Rankine in throat assuming polytropic process if Type==1 % cold flow T_core = T_throat; elseif Type==2 %hot fire T_core = T_core; end T_mean = (1/(1+BPR))*T_core + (BPR/(1+BPR))*T_throat; % Average temperature based on weighted mass flow ratio T_mean_total = ((gamma+1)/2)*T_mean; % Assume flow is choked at throat mdot_tot = P_total*144*A_plug_throat*Cd_nozzle*((gc*gamma)/(R*T_mean_total)) ^0.5*((gamma+1)/2)^-((gamma+1)/(2*(gamma-1))); mdot_core = mdot_tot/(1+BPR); mdot_bypass = mdot_tot - mdot_core; %Guess orifice pressures and iterate to get final orifice pressures nb = 1.1; %polytropic contstant for the bypass stream c = 50; d = 400; P_bypass_orifice_guess = (c+d)/2; bypass_difference = 20; count = 1; while abs(bypass_difference) > 10 T_bypass_orifice = T_tank*(P_total/P_bypass_orifice_guess)^((nb-1)/nb); %Temperature in Rankine in throat assuming polytropic process P_bypass_orifice = (((mdot_bypass/(A_bypass_orifice*Cd_orifice)) *((R*T_bypass_orifice)/(gamma*gc))^0.5*((2/(gamma+1))^((gamma+1)/ (2*(gamma-1))))^-1))/144; % Desired pressure upstream of bypass orifice in psi assuming that Mach %number is neglibible upstream of orifice plate bypass_difference = P_bypass_orifice - P_bypass_orifice_guess; if bypass_difference > 0 c = P_bypass_orifice; else d = P_bypass_orifice; end

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P_bypass_orifice_guess = (c+d)/2; count = count + 1; end nc = 1.05; %polytropic constant for core stream; q = 100; r = 2000; P_core_orifice_guess = (q+r)/2; core_difference = 100; kount = 1; while abs(core_difference) > 10 T_core_orifice = T_tank*(P_total/P_core_orifice_guess)^((nc-1)/nc); %Temperature in Rankine in throat assuming polytropic process P_core_orifice = (((mdot_core/(A_core_orifice*Cd_orifice)) *((R*T_core_orifice)/(gamma*gc))^0.5*((gamma+1)/2)^((gamma+1)/(2*(gamma-1)))))/144; % Desired pressure upstream of bypass orifice in psi assuming that Mach %number is neglibible upstream of orifice plate core_difference = P_core_orifice - P_core_orifice_guess; if core_difference > 0 q = P_core_orifice; else r = P_core_orifice; end P_core_orifice_guess = (q+r)/2; kount = kount + 1; end P_bypass_set = P_bypass_orifice; %Pressure drop in bypass stream negligible P_core_set = 1.179*P_core_orifice; %Pressure drop through core from %emprirical observations P_vitiator_cold = (((mdot_core/(A_vitiator_orifice*Cd_orifice))* ((R*T_core_orifice)/(gamma*gc))^0.5*((gamma+1)/2)^ ((gamma+1)/(2*(gamma-1)))))/144; %% Display Values if Type == 1 fprintf('The nozzle pressure ratio is: %5.3f\n',NPR) fprintf('The bypass ratio is: %5.3f\n',BPR) fprintf('The total airflow is: %5.3f lbm/s \n',mdot_tot) fprintf('The bypass airflow is: %5.3f lbm/s \n',mdot_bypass) fprintf('The core airflow is: %5.3f lbm/s \n',mdot_core) fprintf('The vitiator pressure is: %5.1f psi \n',P_vitiator_cold) fprintf('Set the bypass stream pressure to: %5.1f psi \n',P_bypass_set) fprintf('Set the core stream pressure to: %5.1f psi \n',P_core_set) %Line Output for Type 1 Test %[mdot_tot mdot_core mdot_bypass mdot_fuel P_core_set P_bypass_set %P_vitiator_cold] elseif Type==2 %% Vitiator Calculations

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a = 30; %Low end OF guess for bisection method b = 120; %High end OF guess for bisection method OF = (a+b)/2; T_feed = T_core_orifice; % Assumed air temperature entering combustor, %colder than tank due to blowdown error = 1; k = 1; while abs(error) > 0.1; %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% % Delete archived CEA files %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% delete Detn.inp delete Detn.out delete Detn.plt %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% Pc = [100]; %% Chamber pressure [psia] --> initial value to execute CEA %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% % CEA input/output block for Jet-A and Air %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% ox1 = 'Air'; %% primary oxidizer ox2 = ' '; %% secondary oxidizer fu1 = 'Jet-A(L)'; %% primary fuel fu2 = ' '; %% secondary fuel ox1wt = 100; %% wt fraction of primary oxid in total oxid [1] ox2wt = 0; %% wt fraction of secondary oxid in total oxid [1] fu1wt = 100; %% wt fraction of primary oxid in total oxid [1] fu2wt = 0; %% wt fraction of secondary oxid in total oxid [1] ox1T = 480; %% optional input of primary oxid temperature which %enthalpy is evaluated [degR] ox2T = 0; %% optional input of secondary oxid temperature %which enthalpy is evaluated [degR] fu1T = 500; %% optional input of primary fuel temperature %which enthalpy is evaluated [degR] fu2T = 0; %% optional input of secondary fuel temperature %which enthalpy is evaluated [degR] PR = []; %% pressure ratios [Pc/Pe] subar = []; %% subsonic area ratios [A/At] supar = []; %% supersonic area ratios [A/At] CR = 0; %% chamber contraction ratio [Ac/At] flow = 'eq'; %% flow type [eq or fz] out = 'Ar t O2 H2O CO2 N2 mol gam'; %% maximum eight output parameters in one call to CEA

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for i=1:length(OF) inpgen_rocket_2(ox1,ox2,ox1wt,ox2wt,ox1T,ox2T,fu1,fu2,fu1wt,fu2wt,fu1T, fu2T,Pc,OF(i),PR,subar,supar,CR,flow,out); %% execute CEA600 system('CEA600.exe'); %% extract output from plot file (.plt) DATA = load('Detn.plt'); Tc(i) = DATA(1,2); %% adiabatic flame temperature [K] MW(i) = DATA(1,7); %% molecular weight gam(i) = DATA(1,8); %% ratio of specific heats cstar(i)=sqrt(8314.472*Tc(i)/(MW(i)*gam(i)))*((gam(i)+1)/2)^ ((gam(i)+1)/(2*(gam(i)-1))); end cstar = cstar/0.3048; %% cstar, fps Tc_F = Tc*9/5-460; %% convert adiabatic T to degF Tc = Tc_F + 460; if Tc > T_core a = OF; OF = (a+b)/2; else b = OF; OF = (a+b)/2; end error = (T_core - Tc) k = k+1; end %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% % % Chamber Pressure, Chamber Temperature, and C* % %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% FAR = 1/OF; mdot_fuel = 1.2*mdot_core*FAR; % Combustor has demonstrated approximately 80% efficiency in current configuration mdot_fuel_GPM = mdot_fuel*8.9772; % Fuel mass flowrate in gallons per minute Pc = (mdot_core+mdot_fuel)*cstar/(gc*(A_vitiator_orifice*Cd_orifice)*144); %% chamber pressure P_vitiator_cold = (((mdot_core/(A_vitiator_orifice*Cd_orifice)) *((R*T_core_orifice)/(gamma*gc))^0.5*((gamma+1)/2)^ ((gamma+1)/(2*(gamma-1)))))/144;

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%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% % % Fuel Pressures % %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% % %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% % Fuel Injector Pressure Drop Calculation P_injector = ((mdot_fuel + 0.1194)/0.0124)^2; % Injector curve fit from %factory calibration data Cv_marotta = 0.245; % Marotta MV-74 flow coefficent P_marotta = (mdot_fuel*8.9772*(SL)^0.5/Cv_marotta)^2; %Pressure drop across %Marotta valve in psi --> Experiments suggest this underestimates P_vitiator_hot = Pc; % Vitiator Pressure as calculated by CEA P_fuel_downstream = P_vitiator_hot + P_injector + P_marotta; %Required %pressure downstream of fuel control valve in psi P_fuel_upstream = P_fuel_downstream + 0.15*P_injector; % Adjustment for %pump performance %% Display of Pertinent Values fprintf('The nozzle pressure ratio is: %5.3f\n',NPR) fprintf('The bypass ratio is: %5.3f\n',BPR) fprintf('The total airflow is: %5.3f lbm/s \n',mdot_tot) fprintf('The bypass airflow is: %5.3f lbm/s \n',mdot_bypass) fprintf('The core airflow is: %5.3f lbm/s \n',mdot_core) fprintf('The fuel flow needed is: %5.3f lbm/s \n',mdot_fuel) fprintf('The vitiator pressure before ignition is: %5.1f psi \n', P_vitiator_cold) fprintf('The vitiator pressure after ignition is: %5.1f psi \n', P_vitiator_hot) fprintf('The core temperature is: %5.1f degrees F \n',Tc_F) fprintf('Set the bypass stream pressure to: %5.1f psi \n',P_bypass_set) fprintf('Set the core stream pressure to: %5.1f psi \n',P_core_set) fprintf('The required pressure upstream of the injector is: %5.1f psi \n', P_fuel_downstream) fprintf('Set the fuel backpressure regulator to: %5.1f psi \n', P_fuel_upstream) %Line Output for Type 2 Test % [mdot_tot mdot_core mdot_bypass mdot_fuel P_core_set P_bypass_set %P_vitiator_hot P_fuel_downstream] end

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Appendix B. Plume Data Reduction Code

%************************************************************ % FILENAME: BANR_Plume_Data.m % AUTHOR: Alex Sandroni % DATE CREATED: 24 March 2009 % MODIFIED BY: Alex Sandroni % DATE MODIFIED: 13 April 2009 %************************************************************ %************************************************************ %This code is used to reduce temperature and pressure data gathered by the plume rake. It reads temperature and positioning data from the main test data file. It reads pressure data from the ESP module test data files. %It uses the parameters of the plume motion control program to organize the data into arrays in order to create spatial maps of pressure and temperature. %************************************************************ clear all close all clc %% Test Setup and Configuration %************** % File Details %************** ScanRate = 100; % Set by LabVIEW program (scans/sec) StartTime = 40; % Time before test period (sec) ShiftTime = 1; % Time to position for next sweep (sec) Sweeps = 20; % Number of sweeps to perform during test SweepTime = 5; % Time to perform a sweep (sec) sweeppoints = ScanRate*SweepTime + 1; % Data points taken per scan TestPrefix = 'Plume Sweep - Exit Plane '; ZeroDataFNamePrefix = '04_09_2009_223840'; TestDataFNamePrefix = '04_09_2009_233718'; dataFilePath = [pwd filesep 'Main_Test_Data' filesep]; % set default grid on set(0,'DefaultAxesXGrid','on','DefaultAxesYGrid','on','DefaultAxesZGrid','on') %========================================================================== %========================================================================== % LOAD TEST CONFIGURATION

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%========================================================================== %========================================================================== PlotMonth=TestDataFNamePrefix(1:2); PlotDay=TestDataFNamePrefix(4:5); PlotYear=TestDataFNamePrefix(7:10); PlotHour=TestDataFNamePrefix(12:13); PlotMin=TestDataFNamePrefix(14:15); PlotSec=TestDataFNamePrefix(16:17); PlotDate=[PlotMonth, '/', PlotDay, '/', PlotYear]; PlotTime=[PlotHour, ':', PlotMin, ':', PlotSec]; TestID=[TestPrefix, ' - ', PlotDate, ' (', PlotTime, ') ']; disp(sprintf('\nData Reduction for %s\n',TestID)); %========================================================================== %========================================================================== %% CONSTANTS %========================================================================== %========================================================================== g = 32.174; % Gravity factor (lbm-ft)/(lbf-sec^2) Patm = 14.679; % Atmospheric pressure (psia) gamma = 1.4; % Ratio of specific heats for air R = 1716/32.2; % Gas constant for air (ft-lbf)/(lbm-R) %========================================================================== %========================================================================== %% LOAD TEST DATA FILE AND ASSIGN VARIABLES %========================================================================== %========================================================================== disp(sprintf('Reading Test Data and Transfer Files: %s', [TestDataFNamePrefix '.xls'])); t0 = clock; datafile = load([dataFilePath TestDataFNamePrefix '.xls']); % Calculate Length of Test scans = length(datafile(:,1)); dt = 1/ScanRate; Time = ((1:scans)*dt)'; % Length of test data file [s]

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%========================================== % Temperatures (all temperatures in deg F) %========================================== %Put all raw thermocouple data into array for conversion routine CJC_dF = datafile(:,69); % Thermistor Temperature K_F = [datafile(:,1) datafile(:,2) datafile(:,3) datafile(:,4) datafile(:,5) datafile(:,6) datafile(:,7) datafile(:,8) datafile(:,9) datafile(:,10) datafile(:,11) datafile(:,12) datafile(:,13) datafile(:,14) datafile(:,15) datafile(:,16) datafile(:,17) datafile(:,18) datafile(:,19) datafile(:,20) datafile(:,21) datafile(:,22) datafile(:,23) datafile(:,24) datafile(:,25) datafile(:,26) datafile(:,27) datafile(:,28) datafile(:,29) datafile(:,30) datafile(:,31) datafile(:,32) datafile(:,33) datafile(:,34) datafile(:,35) datafile(:,36) datafile(:,37) datafile(:,38) datafile(:,39) datafile(:,40) datafile(:,41) datafile(:,42) datafile(:,43) datafile(:,44) datafile(:,45) datafile(:,46) datafile(:,47) datafile(:,48) datafile(:,49) datafile(:,50) datafile(:,51) datafile(:,52) datafile(:,53) datafile(:,54) datafile(:,55) datafile(:,56) datafile(:,57) datafile(:,58) datafile(:,59) datafile(:,60) datafile(:,61) datafile(:,62) datafile(:,63) datafile(:,64) datafile(:,65) datafile(:,66) datafile(:,67) datafile(:,68)]; %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% %% LOAD PRESSURE DATA FROM ESP DATA REDUCTION CODE %% %%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%% PlumeTransFile = [pwd filesep 'transfer' filesep 'plume_transfer_' TestDataFNamePrefix '.dat']; % binary transfer file fin = fopen(PlumeTransFile,'r'); plume_scans = fread(fin,1,'integer*4'); plume_ScanRate = fread(fin,1,'integer*4'); % read used pressure transducers P_Plume_1 = fread(fin,plume_scans,'real*4'); P_Plume_2 = fread(fin,plume_scans,'real*4'); P_Plume_3 = fread(fin,plume_scans,'real*4'); P_Plume_4 = fread(fin,plume_scans,'real*4'); P_Plume_5 = fread(fin,plume_scans,'real*4'); P_Plume_6 = fread(fin,plume_scans,'real*4'); P_Plume_7 = fread(fin,plume_scans,'real*4'); P_Plume_8 = fread(fin,plume_scans,'real*4'); P_Plume_9 = fread(fin,plume_scans,'real*4'); P_Plume_10 = fread(fin,plume_scans,'real*4'); fclose(fin);

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%========================================================================== %========================================================================== %% DATA REDUCTION %========================================================================== %========================================================================== %========================================== % Temperatures %========================================== [CJC_mV,CJC_dF] = CJC_Volts_To_mV_For_Type_K(CJC_dF); rows = size(K_F); CJC_mV = CJC_mV*ones(1,rows(2)); [K_Output_dF] = Type_K_V_to_dF_with_Ref_Temp(K_F,CJC_mV); T_Plume_1 = K_Output_dF(:,57); % Temperature Probe 1 T_Plume_2 = K_Output_dF(:,58); % Temperature Probe 2 T_Plume_3 = K_Output_dF(:,59); % Temperature Probe 3 T_Plume_4 = K_Output_dF(:,60); % Temperature Probe 4 T_Plume_5 = K_Output_dF(:,61); % Temperature Probe 5 T_Plume_6 = K_Output_dF(:,62); % Temperature Probe 6 T_Plume_7 = K_Output_dF(:,63); % Temperature Probe 7 T_Plume_8 = K_Output_dF(:,64); % Temperature Probe 8 T_Plume_9 = K_Output_dF(:,65); % Temperature Probe 9 T_Plume_10 = K_Output_dF(:,66); % Temperature Probe 10 disp(sprintf(' Finished: %.4f sec\n',etime(clock,t0))); disp(sprintf('Performing Data Reduction, Please Stand By...')); t0 = clock; %============================================ % Filtering of Temperature and Pressure Data %============================================ [d,c] = butter(4,2/(ScanRate/2)); T_Plume_1_filtered = filter(d,c,T_Plume_1); T_Plume_2_filtered = filter(d,c,T_Plume_2); T_Plume_3_filtered = filter(d,c,T_Plume_3); T_Plume_4_filtered = filter(d,c,T_Plume_4); T_Plume_5_filtered = filter(d,c,T_Plume_5); T_Plume_6_filtered = filter(d,c,T_Plume_6); T_Plume_7_filtered = filter(d,c,T_Plume_7); T_Plume_8_filtered = filter(d,c,T_Plume_8); T_Plume_9_filtered = filter(d,c,T_Plume_9); T_Plume_10_filtered = filter(d,c,T_Plume_10); %========================================== % Conversion of Raw Positioning Data %========================================== Vsupply_X = 24; % Voltage supplying voltage divider circuit (horizontal)

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Vsupply_Y = 24; % Voltage supplying voltage divider circuit (vertical) Vread_X = datafile(:,115); % Voltage across potentiometer (horizontal) Vread_Y = datafile(:,116); % Voltage across potentiometer (vertical) Rdiv_X = 15e3; % Constant resistor in voltage divider circuit (horizontal) Rdiv_Y = 4.7e3; % Constant resistor in voltage divider circuit (vertical) R_X_initial = 5000; % Resistance of potentiometer at zero (horizontal) R_Y_initial = 110; % Resistance of potentiometer at zero (vertical) R_X = Rdiv_X./((Vsupply_X./Vread_X) - 1); % Resistance of pot. (horizontal) R_Y = Rdiv_Y./((Vsupply_Y./Vread_Y) - 1); % Resistance of pot. (vertical) X = 0.0030.*(R_X_initial - R_X); % X position from potentiometer feedback Y = 0.0013.*(R_Y - R_Y_initial); % Y position from potentiometer feedback %========================================== % Creation of Arrays %========================================== a = StartTime*ScanRate; sweep = SweepTime*ScanRate; b = a + sweep; Horizontal(1,:) = X(a:b); Vertical(1,:) = Y(a:b); Y_ave(1) = mean(Vertical(1,:)); Temp_1(1,:) = T_Plume_1_filtered(a:b); Temp_2(1,:) = T_Plume_2_filtered(a:b); Temp_3(1,:) = T_Plume_3_filtered(a:b); Temp_4(1,:) = T_Plume_4_filtered(a:b); Temp_5(1,:) = T_Plume_5_filtered(a:b); Temp_6(1,:) = T_Plume_6_filtered(a:b); Temp_7(1,:) = T_Plume_7_filtered(a:b); Temp_8(1,:) = T_Plume_8_filtered(a:b); Temp_9(1,:) = T_Plume_9_filtered(a:b); Temp_10(1,:) = T_Plume_10_filtered(a:b); Press_1(1,:) = P_Plume_1(a:b); Press_2(1,:) = P_Plume_2(a:b); Press_3(1,:) = P_Plume_3(a:b); Press_4(1,:) = P_Plume_4(a:b); Press_5(1,:) = P_Plume_5(a:b); Press_6(1,:) = P_Plume_6(a:b); Press_7(1,:) = P_Plume_7(a:b); Press_8(1,:) = P_Plume_8(a:b); Press_9(1,:) = P_Plume_9(a:b); Press_10(1,:) = P_Plume_10(a:b); shift = ShiftTime*ScanRate;

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for i = 2:Sweeps a = b + shift; b = a + sweep; Horizontal(i,:) = X(a:b); Vertical(i,:) = Y(a:b); Y_ave(i) = mean(Vertical(i,:)); Temp_1(i,:) = T_Plume_1(a:b); Temp_2(i,:) = T_Plume_2(a:b); Temp_3(i,:) = T_Plume_3(a:b); Temp_4(i,:) = T_Plume_4(a:b); Temp_5(i,:) = T_Plume_5(a:b); Temp_6(i,:) = T_Plume_6(a:b); Temp_7(i,:) = T_Plume_7(a:b); Temp_8(i,:) = T_Plume_8(a:b); Temp_9(i,:) = T_Plume_9(a:b); Temp_10(i,:) = T_Plume_10(a:b); Press_1(i,:) = P_Plume_1(a:b); Press_2(i,:) = P_Plume_2(a:b); Press_3(i,:) = P_Plume_3(a:b); Press_4(i,:) = P_Plume_4(a:b); Press_5(i,:) = P_Plume_5(a:b); Press_6(i,:) = P_Plume_6(a:b); Press_7(i,:) = P_Plume_7(a:b); Press_8(i,:) = P_Plume_8(a:b); Press_9(i,:) = P_Plume_9(a:b); Press_10(i,:) = P_Plume_10(a:b); end Y_ave = Y_ave'; Y_coord = Y_ave*ones(1,SweepTime*ScanRate+1); for u = 2:2:Sweeps Horizontal(u,:) = fliplr(Horizontal(u,:)); Temp_1(u,:) = fliplr(Temp_1(u,:)); Temp_2(u,:) = fliplr(Temp_2(u,:)); Temp_3(u,:) = fliplr(Temp_3(u,:)); Temp_4(u,:) = fliplr(Temp_4(u,:)); Temp_5(u,:) = fliplr(Temp_5(u,:)); Temp_6(u,:) = fliplr(Temp_6(u,:)); Temp_7(u,:) = fliplr(Temp_7(u,:)); Temp_8(u,:) = fliplr(Temp_8(u,:)); Temp_9(u,:) = fliplr(Temp_9(u,:)); Temp_10(u,:) = fliplr(Temp_10(u,:)); Press_1(u,:) = fliplr(Press_1(u,:)); Press_2(u,:) = fliplr(Press_2(u,:)); Press_3(u,:) = fliplr(Press_3(u,:)); Press_4(u,:) = fliplr(Press_4(u,:)); Press_5(u,:) = fliplr(Press_5(u,:)); Press_6(u,:) = fliplr(Press_6(u,:)); Press_7(u,:) = fliplr(Press_7(u,:)); Press_8(u,:) = fliplr(Press_8(u,:)); Press_9(u,:) = fliplr(Press_9(u,:)); Press_10(u,:) = fliplr(Press_10(u,:)); end

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% Create arrays for all probes d_probe = 0.125; dy = 1.26; % Probe center-to-center distance offset = 0.184; % Probe 11 is this distance above rig centerline to start Y_T_1 = Y_coord + offset + 10*dy; % Top probe (PL-1) Y_P_1 = Y_coord + offset + 9*dy; % PL-2 Y_T_2 = Y_coord + offset + 8*dy; % PL-3 Y_P_2 = Y_coord + offset + 7*dy; % PL-4 Y_T_3 = Y_coord + offset + 6*dy; % PL-5 Y_P_3 = Y_coord + offset + 5*dy; % PL-6 Y_T_4 = Y_coord + offset + 4*dy; % PL-7 Y_P_4 = Y_coord + offset + 3*dy; % PL-8 Y_T_5 = Y_coord + offset + 2*dy; % PL-9 Y_P_5 = Y_coord + offset + dy; % PL-10 Y_P_6 = Y_coord + offset; % PL-11 (closest probe to actual rig centerline) Y_T_6 = Y_coord + offset - dy; % PL-12 Y_P_7 = Y_coord + offset - 2*dy; % PL-13 Y_T_7 = Y_coord + offset - 3*dy; % PL-14 Y_P_8 = Y_coord + offset - 4*dy; % PL-15 Y_T_8 = Y_coord + offset - 5*dy; % PL-16 Y_P_9 = Y_coord + offset - 6*dy; % PL-17 Y_T_9 = Y_coord + offset - 7*dy; % PL-18 Y_P_10 = Y_coord + offset - 8*dy; % PL-19 Y_T_10 = Y_coord + offset - 9*dy; % Bottom probe (PL-20) % Create arrays to interpolate grid points between measured data [Xi Yi] = meshgrid(-2:.05:6,-10:.05:10); Temp_1i = griddata(Horizontal,Y_T_1,Temp_1,Xi,Yi); Temp_2i = griddata(Horizontal,Y_T_2,Temp_2,Xi,Yi); Temp_3i = griddata(Horizontal,Y_T_3,Temp_3,Xi,Yi); Temp_4i = griddata(Horizontal,Y_T_4,Temp_4,Xi,Yi); Temp_5i = griddata(Horizontal,Y_T_5,Temp_5,Xi,Yi); Temp_6i = griddata(Horizontal,Y_T_6,Temp_6,Xi,Yi); Temp_7i = griddata(Horizontal,Y_T_7,Temp_7,Xi,Yi); Temp_8i = griddata(Horizontal,Y_T_8,Temp_8,Xi,Yi); Temp_9i = griddata(Horizontal,Y_T_9,Temp_9,Xi,Yi); Temp_10i = griddata(Horizontal,Y_T_10,Temp_10,Xi,Yi); Press_1i = griddata(Horizontal,Y_P_1,Press_1,Xi,Yi); Press_2i = griddata(Horizontal,Y_P_2,Press_2,Xi,Yi); Press_3i = griddata(Horizontal,Y_P_3,Press_3,Xi,Yi); Press_4i = griddata(Horizontal,Y_P_4,Press_4,Xi,Yi); Press_5i = griddata(Horizontal,Y_P_5,Press_5,Xi,Yi); Press_6i = griddata(Horizontal,Y_P_6,Press_6,Xi,Yi); Press_7i = griddata(Horizontal,Y_P_7,Press_7,Xi,Yi); Press_8i = griddata(Horizontal,Y_P_8,Press_8,Xi,Yi); Press_9i = griddata(Horizontal,Y_P_9,Press_9,Xi,Yi); Press_10i = griddata(Horizontal,Y_P_10,Press_10,Xi,Yi); disp(sprintf(' Finished: %.4f sec\n',etime(clock,t0)));

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disp(sprintf('Proceeding to Plotting...')); t0 = clock; %========================================================================= %========================================================================= %% Plotting %========================================================================= %========================================================================= % Create outlines of plug throat and nozzle shroud theta = [0:pi/100:2*pi]; rplug = 5.586/2; rshroud = 9.63/2; xplug = rplug.*cos(theta); yplug = rplug.*sin(theta); xshroud = rshroud.*cos(theta); yshroud = rshroud.*sin(theta); % Create temperature and pressure maps of the plume figure(1) %title('Temperature Map') xlabel('X (in.)'),ylabel('Y (in.)') hold on pcolor(Xi,Yi,Temp_1i) pcolor(Xi,Yi,Temp_2i) pcolor(Xi,Yi,Temp_3i) pcolor(Xi,Yi,Temp_4i) pcolor(Xi,Yi,Temp_5i) pcolor(Xi,Yi,Temp_6i) pcolor(Xi,Yi,Temp_7i) pcolor(Xi,Yi,Temp_8i) pcolor(Xi,Yi,Temp_9i) pcolor(Xi,Yi,Temp_10i) % pcolor(Horizontal,Y_T_1,Temp_1) % pcolor(Horizontal,Y_T_2,Temp_2) % pcolor(Horizontal,Y_T_3,Temp_3) % pcolor(Horizontal,Y_T_4,Temp_4) % pcolor(Horizontal,Y_T_5,Temp_5) % pcolor(Horizontal,Y_T_6,Temp_6) % pcolor(Horizontal,Y_T_7,Temp_7) % pcolor(Horizontal,Y_T_8,Temp_8) % pcolor(Horizontal,Y_T_9,Temp_9) % pcolor(Horizontal,Y_T_10,Temp_10) plot(xshroud,yshroud,'k-','LineWidth',2) plot(xplug,yplug,'k--','LineWidth',2) shading interp axis([-2 6 -4 4]) caxis([-100 1000]) colorbar

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figure(2) %title('Pressure Map') xlabel('X (in.)'),ylabel('Y (in.)') hold on pcolor(Xi,Yi,Press_1i) pcolor(Xi,Yi,Press_2i) pcolor(Xi,Yi,Press_3i) pcolor(Xi,Yi,Press_4i) pcolor(Xi,Yi,Press_5i) pcolor(Xi,Yi,Press_6i) pcolor(Xi,Yi,Press_7i) pcolor(Xi,Yi,Press_8i) pcolor(Xi,Yi,Press_9i) pcolor(Xi,Yi,Press_10i) % pcolor(Horizontal,Y_P_1,Press_1) % pcolor(Horizontal,Y_P_2,Press_2) % pcolor(Horizontal,Y_P_3,Press_3) % pcolor(Horizontal,Y_P_4,Press_4) % pcolor(Horizontal,Y_P_5,Press_5) % pcolor(Horizontal,Y_P_6,Press_6) % pcolor(Horizontal,Y_P_7,Press_7) % pcolor(Horizontal,Y_P_8,Press_8) % pcolor(Horizontal,Y_P_9,Press_9) % pcolor(Horizontal,Y_P_10,Press_10) plot(xshroud,yshroud,'k-','LineWidth',2) plot(xplug,yplug,'k--','LineWidth',2) shading interp axis([-6 6 -6 6]) caxis([14.7 35]) colorbar disp(sprintf(' Finished: %.4f sec\n',etime(clock,t0)));