High Pressure Regenerative Turbine Engine: 21st Century Propulsion

Table of Contents Chapter 1 Introduction 1 1.1 Background 1 1.2 HPRTE Cycle Description 2 1.3 Objectives 1.4 Scope 4 1.5 Roles and Responsibilities of the University of F lo ri da 5 1.6 High Recirculation Combustion Program 5 Chapter 2 HPRTE Cycle Design 2.1 Preliminary Performance Studies 9 2.1.1 Turbocharger Selection 10 2.1.2 T62T32A Axial Thrust 2.1 .3 Recuperator and Intercooler 14 2.1A Duct Sizing 14 2.1.5 Combustor 15 Chapter 3 Experimental Apparatus 17 3.1 Titan HPRTE Engine Layout 3.1.1 Temperatures 3.1.2 Pressures 17 3.1.3 Nomenclature for Instrumentation 3.2 Experimental Installation 3.3 Support Equipment 3.3.1 Dynamometer System 26 3.3.2 Spray Cooler System 3.3.3 Laboratory Water Supply 3.3A Wastegate Valve and Controls 3.3.5 Oil Heat Exchanger 3 A Instrumentation 3 A.l Data Acquisition System 3A.2 Gas Analysis System Chapter 4 Operating Procedures for the Titan HPRTE 4.1 Support Systems 4.2 Engine Room Preparation 4.3 Data Acquisition and Gas Analysis 45 4A Video and Audio Recording 4.5 Operational Options Chapter 5 Experimental Data Chapter 6 Experimental Results 6.1 Final Test Series Results 53 NASNCR-200 1-21 0675 v

6.1.1 Test Run 22 Performance Data 54 6.1.2 Test Rilll 22 Emissions Data 63 6.2 Shakedown Testing Chapter 7 Data Analysis 89 Chapter 8 Conclusions and Recommendations 8.1 Experimental Conclusions 8.1.1 Emissions 97 8.1.2 Engine Performance 99 8.2 Recommendations 101 8.2.1 Experimental Program Recommendations 8.2.2 System Study Recommendations 103 Chapter 9 HPRTE Cycle Implications 105 9.1 Rolls Royce Allison Simulation 105 9.2 Helicopter Applications 106 9.3 aval Vessel Application 112 9.4 Combined Cycle Power Generation Appendix A Water Injection System Operator Instructions 121 Appendix B Data Acquisition Software Configuration 123 Appendix C Gas Analysis Setup Procedure 127 Appendix D HPRTE Engine Operation 131 D.I Kahn Hydraulic Dynamometer Set-up D.2 Spray-Cooler Set-up D.3 Crew Assignments 132 D.3.1 FirelEngine Room Overwatch 132 D.3.2 Analog Pressure Panel D.3.3 Analog Temperature Panel D.3.4 Digital Data Acquisition D.3.5 Gas Analysis and Spray Cooler D.3.6 Engine Control 133 D.3.7 Operations Supervisor 134 D.4 Titan High Pressure Regenerative Turbine Engine Operation 134 D.4.1 Pre-start Procedure 134 D.4.2 Engine Starting Procedure 134 DA.3 Standard Operational Procedure 135 DAA Shut-Down Procedure 136 References 137 NASAlCR-200 1-21 0675 vi

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Chapter 1. Introduction 1.1. Background In the early history of gas turbine development «1955) , semi-closed cycles were proposed [Anxionnaz 1945, 1948J as alternatives to conventional engines due to their potential advantages in size and weight, low fuel consumption over the power range, and significantly reduced air flow requirements. However, the more complex arrangements to achieve high efficiency together with high sulfur fuels tended to mitigate their development as a result of the risk of corrosion due to the recirculated flow in the compression flowpat h.

Two known engine development programs in this timeframe are the works of the Sulzer Brothers [Baumeister et al.] and the U.S. Navy Project Wolverine [DeWitt and Boyum 1956; Davis 1956]. A 5-MW and 20-MW power plant were in operation by the Sulzer Brothers in 1945 to 1949 that were problematic due to burning of crude oil that formed deposits and corrosi on in the combustor. Project Wolverine was a classified submarine propulsion application of a semi-closed cycle intended to provide a fallback system in the event that nuclear propulsion proved unfeasible. However, initial testing on the propulsion system demonstrated the predicted attributes of semi-closed cycle engines.

Further discussions on Project Wolverine will be provided in a later section in this introduction.

In the intervening years since Project Wolverine, the technology of gas turbine engines, component efficiency, and low-sulfur fuels, together with the rapid expansion in the market for commercial recuperated engines, has provided for an opportune re- examination of semi-closed cycles for a wide range of applications. A number of additional arrangements have been proposed [Gasparovic 1967, 1968J for semi-closed cycles that may be used in power generation or propulsion applications.

NAS A/CR- 2001-210675 The project described in this report deals with the demonstration of a type of semi-closed cycle called the High Pressure Recuperative Turbine Engine (HPR TE) , formerly known as the Regenerative Feedback Turbine Engine (RFTE). The motivation for this engine concept will be presented, followed by the descriptions of the test engine, the experimental results, an analysis of the results, and implications for applications, which could benefit from this technology.

1.2. HPRTE Cycle Description Figure 1.1 shows a schematic diagram of the High Pressure Regenerative Turbine Engine (HPRTE) cycle as implemented in this program. Air enters the low pressure compressor (LPC, also designated Cl) at state 1, exiting at state 2. The compressed air then mixes with recirculated exhaust to reach state 2.1 before being cooled. The cooling process is via an optional spray cooler and an intercooler; the exit state is state 3. The air/exhaust mixture enters the high pressure compressor (HPC, also designated C2) . The path from state 3 to state 9 is that of a conventional recuperated gas turbine engine, except that the inlet and exit pressures (P 3 and P9) are elevated (though nearly equal ). At the recuperator exit, the flow splits, with part being routed to the mixing junction by state 2 and the remainder passing through the low pressure turbine (LPT, also designated T2). In the current configuration, the low pressure turbomachinery, Cl and T2, are implemented by a turbocharger, with a wastegate valve controlling LPT bypass flow. The engine may be viewed as an intercooled, recuperated gas turbine engine that has been turbocharged and that has large exhaust gas recirculation.

As in the case of turbocharged piston engines, the low pressure spool serves to improve the power density of the HPRTE. The present demonstration project utilized a turbocharger pressure of only 2: 1, so there was no compactness benefit, especially for the modular breadboard design of the experimental rig. However, for most applications, it is expected that LPC pressure ratios of 5: 1 or more would be the optimum. The increase in gas density in the recuperator is estimated to allow a factor of 20 decrease in volume, for fixed effectiveness and relative pressure drop. Thus the compactness of a prototype I I NASAl CR - 2001-210675 2 I I

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I I HPRTE would be expected to be comparable to simple-cycle engines, with a much smaller penalt y in we ight and volume than for other recuperated engines. The improvement in efficiency, especially at part load, attributed to recuperation makes this a very attractive tradeoff for many applications, especially in light of the other advantages of this engine concept.

State Point Test Designation 1 Amb or Airin 2 LPcx 2.1 icool 3 HPci or Filtx HPcx or HPri 5 HPrx 6 HPti 7 HPtx 8 LPtx LPti Figure 1.1. Schematic diagram ofHPRTE cycle.

1.3. Objectives A multi-disciplinary program was conducted to demonstrate proof-of-concept for a novel gas turbine engine based on a semi-closed Brayton cycle. The justification for this effort lies in the unique advantage of using conventional turbomachinery components arranged with recuperation at high pressure and recirculation of exhaust flow (semi-closed part of the cycle) to achieve performance gains. A fully-developed HPRTE is projected to yield NASAlCR -2001-21 06 75 3

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I t specific power 2.5 to 3 times that of current state-of-the-art gas turbine engines, nearly l constant specific fuel consumption (SFC) over 80 % of the power range, combustion temperatures within state-of-the-art limits (no required materials development), and inherent emission reduction. These performance characteristics have been independently validated analytically by major engine companies, government organizations, and other institutions. The validation of the cycle was a result of introducing the concept to the various agencies by the team, which was led by Science Applications International Corporation with support by the U.S. Army Vehicle Propulsion Directorate (VPD) that eventually led to this NASA NRA project. The program was in part intended to contribute to NASA ' s mission goal to resolve critical environmental issues of atmospheric emissions while providing the opportunity for maintaining propulsive technology dominance and improving worldwide marketing benefits for the U.S.

economy through improvements of ongoing DoDINASA programs.

1.4. Scope The effort focused on a cost-effective, relatively low-risk program to achieve the proof- of -concept engine demonstration of the HPRTE. An auxiliary power unit CAPU) of approximately 95 shaft horsepower was selected and was modified into an HPRTE test rig. While this bread-board design does not represent an optimal engine, it provides for demonstration of improved power, part-load efficiency, and emission reduction.

Components (recuperator, combustor, turbocharger, and intercooler) were designed or modified for integration into a modular HPR TE. Design point simulations were performed for approximate component matching, a process that included minor turbo machinery modifications. Testing consisted of a high-pressure test for structural integrity of the basic APU and green testing of the HPRTE by Alturdyne, Inc. and then performance testing at the University of Florida. Control strategies and combustion experiments were evaluated before final assembly in the HPRTE rig. Post-test evaluation produced performance data that can be used for future HPRTE design and analysis. The program had further contributed value through a U.S. Army VPD SBIR Phase II effort that focused on high recirculation flow combustor investigations using a Rover Model NASAlCR- 2001-210675 4 '- - -' --' - - ' -- - ' - ~-' - - - ~-- - ,-,~- 1S60 engme developed into an HPR TE modularized test ng , as well as active participation by NASA in a supporting role.

The NASA program reported herein was directed by Science Applications International Corporation (SAIC) with subcontracts to the University of Florida, Allison Engine Company (now Rolls Royce Allison), Alturdyne, and JOW Consultants, with the U. S.

Army VPD in a supporting role.

1.5. Roles and Responsibilities of the University of Florida At the outset of the program, the role defined for the University of Florida had three major components: • Assist in design specification of test rig • Simulation of HPRTE cycle • Performance testing and data analysis In preparation for the performance testing, the Energy & Gasdynamic Systems Laboratory was developed as a complete engine test facility. As the program evolved, the hardware development difficulties experienced by Alturdyne forced a greater development role to be carried by the University after delivery of the engine. This included modifications to the HPRTE flowpath and components as well as additional infrastructure to support spray cooling and additional instrumentation. Finishing of the engine was also improved, including repair of gaspath leaks, installation of gaskets, improvement of sensors, rework of the starting system, resizing of the fuel nozzle, and fitting the engine for remote operation. Operational schemes were also evolved to overcome the considerable difficulties encountered in starting the system.

1.6. High Recirculation Combu.stion Program A synergistic program was supported by the U.S. Army Vehicle Propulsion Directorate [Crittenden, Lear, and Azzazy 1999] to investigate the design requirements for a HPRTE engine with higher recirculation ratio than that of the NASA demonstration program (recirculation ratio R is defined as the mass flow recirculated divided by the inlet air mass flow). In this program, several tasks were accomplished toward the goal of investigating design scaling laws and combustion instability limits: NASAlCR- 2001-210675 • A test facility was designed and built at the University of Florida that allows high recirculation ratio flow over a range of combustor inlet temperatures and pressures.

The High Recirculation Combustor (HiRC) facility was designed around a Rover 1S60 gas turbine engine in a manner similar to that of the Titan HPRTE, except that lower-cost heat exchanger and ducting components were used since engine performance was not a test objective. Instrumentation and control was similar to that of the Titan HPRTE, except that a bypass valve was used to allow fresh air intake without utilizing the turbocharger. This allowed a range of recirculation ratios to be achieved, within the limits imposed by the Rover maximum allowable combustor temperature.

• An experimental combustor was designed by Rolls Royce Allison for design recirculation ratio R of 2 (compared to 1 for the Titan HPRTE). Conventional preliminary design methodologies, including CFD, predicted a required increase in primary zone volume of eight times (8X), confirmed by modeling at the University of Florida.

• The 8x combustor was built, including effusion cooling and a thermal barrier coating.

Manufacturing difficulties delayed the combustor completion so that no shakedown testing was done at Allison, and only preliminary tests were performed at the University.

• In parallel with the 8x combustor development, the original combustor from the Rover engine was modified to change the air flow splits in the combustion zones.

This simple modification produced a combustor that met most of the goals of the program, with successful testing beyond R of unity without evidence of combustion instability. Limitations in the rig at that time prevented higher R tests, so the presumption is that the conventional design may have allowed the design recirculation ratio to be approached.

• Optical diagnostics were developed to allow mapping of the flow field using two non- intrusive techniques. The diagnostic system was validated and benchtop tested, but due to the time limitations, was not integrated into the Rolls Royce Allison R=2 combustor.

ASAICR -2001-210675 6 The chief conclusion from the synergistic Army program was that development of a high recirculation combustor is feasible, but that the scaling laws must be more carefully studied. The decision to build the experimental burner with an 8x volume increase was based on gas-phase reaction rate arguments, but clearly the original burner performed satisfactorily at significant recirculation ratios. One hypothesis is based on the fact that the original volume allows not only the gas-phase reactions, but also the processes of fuel droplet breakup and mixing, soot formation, and soot burnout. Previous work [Meier and Vollerin 1977; Marek and Tacina 1976] indicates that high recirculation interrupts the soot formation process, indicated by the lack of radiant soot emission in the flame zone.

The soot burnout process is likely to occur on a time scale similar to or greater than the gas-phase reaction time. If so , then the reduction of soot formation would in itself reduce the required residence time, hence the primary zone volume. This effect competes with the increase in volume necessitated by the slower gas-phase kinetics, leading to uncertainty in the volume scaling. If the hypothesis is correct, then the volume change required in comparison to conventional designs may be small.

Recommendations from the synergistic Army combustion program include shakedown development of the Allison high recirculation combustor, integration into the HiRe facility, and application of the optical diagnostics. These steps should be coupled with detailed modeling to improve the compactness of future HPRTE designs.

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Chapter 2. HPRTE Cycle Design As described in Chapter 1, the object of this program was to demonstrate a version of the HPRTE, the simplified cycle schematic of which is shown on Figure 1.l. This was accomplished by using a modified Titan T62T32A small gas turbine as the high pressure (HP) spool, supercharged with a low pressure (LP) turbocharger spool. Because of the in crease in pressure in the HP spool due to turbo charging to implement the semi-closed cycle, the Titan T62T32A engine was selected due to its stainless steel case, which would permit safe operation without modification. A recuperator and intercooler were procured to match the HPRTE cycle requirements, and the T62T32A combustor was modified to permit high inlet temperature operation with recirculated flow . Maximum power capability at SL 59 F conditions was estimated to be 150 HP.

2.1. Preliminary Performance Studies An existing performance code developed by Colin Rodgers (Alturdyne) for the T62T32A was modified to incorporate: • LP spool supercharging and intercooling • Recuperation between the HP and LP spools • Recirculation of fractional exhaust flow The existing code had embedded T62T32A normalized compressor and turbine characteristics (Figure 2.1) and had a recuperator option. In prior HPRTE demonstrator cycle studies a recirculation flow ratio R (recirculation flow divided by inlet flow) of unity had been recommended with LP pressure ratios from 1.5 to 2. 0. These ratios were used in first evaluations of the modified T62T32A code, in combination with trial intercooler, recuperator, combustor and duct loss performances. HPRTE maximum cycle temperature tends to be governed by current metallic recuperator inlet temperature limits in the 1100 to 1200 F range . Iterative computations finally led to selection of the preliminary HPRTE cycle listed on Table 2.l. With an LP pressure ratio of 2.0 and recirculation ratio of 1.0, estimated output power and thermal efficiency with a recuperator inlet temperature of 1200 F, were 150 hp and 21.8 % respectively. This NASAlCR-2001 -210675 9 estimated performance level was contingent upon a total cycle pressure loss of 13 .2% with a recuperator and intercooler effectiveness of 84 and 80% respectively .

2.1 . 1. Turbocharger Selection.

Initial turbocharger surveys were made to procure a variable nozzle unit that could be used to optimize the matching of the LP and HP modules. No variable area units capable of passing 1.3 pps with a pressure ratio of 2.0 were available. Communications with Cummins Engine Co. led to the possibility of using the relatively high efficiency HX 8038875 compressor, the performance characteristics of which are shown on Figure 2.2, capable of attaining an efficiency of 78% with a vaneless diffuser. This turbocharger was available with five turbine scroll areas, of which the smallest scroll (HX80-F32 *) best matched the HPRTE requirement. For improv ed turbine efficiency the installation of a vaned nozzle was considered but for expediency, the existing vaneless scroll was retained. Since both the LP compressor and recirculation flows mix prior to entering the intercooler, any flow discontinuity from the compressor was thought possible to precipitate backflow of the recirculated flow into the compressor. As a consequence a bypass or surge dump valve was recommended.

2.1.2. T62T32A Axial Thrust.

Since the T62T32A aerodynamic axial end thrust is proportional to the inlet pressure, supercharging the inlet with an LP pressure ratio of 2.0 essentially doubles the aerodynamic end load on the forward thrust bearing. Computed aerodynamic axial loads for the T62T32A under normal and HPRTE operating conditions increased from 41 to 87 lb. at design speed 72222 rpm., with a BIO fatigue life exceeding 10 ,000 hours. The 19. 5: 1 gearbox was designed for 6000 hours at 225hp. Increased temperature bearing buffer air supply was considered within bearing temperature limitations.

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90.-----.- 70,f-----+----c-I---,L-4 Effldency Shroud dear.nce .011 In 5·f-----r--~1---~~ Prellure R.tlo Nozzle lie. 1.65 .q In PCD/pe TIp die 6.5 In 0.9'f--- --+-----II-----+-----j "Caneeted 80 S,eed lr----PT--~~--~ Velocity Ratio 01/ Vo 1.0 1.5 2.0 2.5 D . 75 0.80 0.85 Figure 2 .1. Performance map of Titan T62T32A. The solid circles indicate data from initial testing at Alturdyne [Rodgers 1997].

NASA/CR -200 1- 21 0 675 --I I I 80~-~--~- ~ Turbine 2.1 Efficiency " 70 I-- ...,£..... ,.L-+----+~~ 0.6 0.8 1.0 Velocity ratio UNO F32- 3 F 48-~ 5.0 In 7.4 In PRESSURE RATIO HX80 Turbine 2 . 01-----+~._+__+__I_~ Expans on Ratio 1. f--_-+--I-----T4-_---I 1. 0~ __ --'- ___ ..L...- __ .-I.. ___ ....L- __ --1 1.0~_---L __ __.l..._ _ __.J o 2 3 0.5 1.0 1.5 2.0 2.5 INLET FLOW F UNCTION W I TIP INLET CORRECTED AIRFLOW PPS WRTHETAIOEl pps/Jl/p sill Figure 2.2 . Performance map of Holset HX80 turbocharger. The solid circles indicate data from initial testing at Alturdyne [Rodgers 1997] .

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Table 2.1. HPRTE design program output.

PROGRAM HPRTE DESIG POINT RUN ON 10-29-1997 RECIR. RATIO 1 RECUP I LET TEMP F 1200 LP SPOOL SPEED % 100 COMP INLET TEMP T 59 2 PSIA PRESSURE RATIO COMP I LET PRESSURE 14.7 AIRFLOW PPS 1.242 COMP EXIT TEMP F 203 COMPRESSOR EFF 0.78 COMP EXIT PRESSURE PSIA 27.8 TURBI EEFF 0.75 TURB INLET TEMP 673 F TURB FLOW FU CT 1.422 TURB EXHAUST TEMP F 547 TURB IN PRESS PSIA 29.4 HP SPOOL COMP INLET TEMP T 151 SPEED % 100 PRESSURE RATIO 4.473 COMP I LET PRESSURE PSIA 27 .8 AIRFLOW PPS 2.484 COMP EXlT TEMP F 572 COMPRESSOR EFF 0.769 COMOP EXIT PRESSURE PSIA 124 .2 7 PSIA 117.77 TURBINE EFF 0.84 TURB INLET PRESSURE TURB FLOW FUNCT 0.98 TURB INLET TEMP F 1747 RECUPERATOR BURNER 5.2 EFFECTIVE ESS % 0.84 BURNER + RECUP PRES LOSS EFFECTIVENESS CO ST 5.622 INLET TEMP F 1099 0.97 MEA TEMPERA TURE F 560 EFFICIE CY GAS SIDE PRES LOSS % 0.025 TEM PRl SE F 647 BTUIHR 2.522X10 HEA T RELEASE FUEL FLOW PPH 91.33 48.97 GAS/FUEL RATIO INTERCOOLER PARASITI CS EFFECTIVENESS 0.8 MECHANICAL LOSSES HP 10 PRESSURE LOSS % 5.5 LEAKAGE % 2 F 438 INLET TEMP EXlTTEMP F 151 HP 48.89 FAN POWER OUTPUT OUTPUT POWER HP 150.8 THERMAL EFFIC. 21.79 0. 63 SFC

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NASAlCR-200 1-210675 13 1 ____ _ 2.1.3. Recuperator and Intercooler.

Preliminary performance goals for the HPRTE recuperator and intercooler were: Recuperator. Effectiveness 82% Pressure drops, hot 3 %, cold 2% Intercooler Effectiveness 85 % Pressure drop 5.5 %, sink temp 80F Contacts with Allied signal for the use of two core modules from the GT601 tank recuperative gas turbine were responsive, and proved to match the recuperator effectiveness and pressure drop requirements . The basic core modules required installation and sealing within a recuperator tailored encasement complete with headers and structural supports. A commercial air to water intercooler built by Elanco was purchased to cool the recirculated flow down to 151 F at the inlet to the HP compressor.

The heat transfer surfaces comprise spined tubing less sensitive to the incident flow direction than conventional finned tubing.

2.1.4. Duct Sizing.

Since the HPRTE requires extensive ducting to couple the LP and HP modules and the recuperator and intercooler, it is necessary to optimize duct sizes to minimize both pressure losses and cost. The estimated effect of duct sizing on internal dynamic heads and velocities is show on Table 2.2, together with duct diameters.

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I NASA/CR-2001-2J0675 14 I I I ____ . _____ . ____________________ . _J 1 __ - __ _ ______ . ___ _ Table 2.2. HPRTE Duct Sizing and Losses DUCT DIA AREA TEMP PRESS RHO q head' LOSS LOSS INCH SQIN R PSIA Ib / ft3 PSIA COEFF PSIA RECUP COLDSIDE COMP OUT 4.0D 12.56 1015 133 0.354 0.27 1.3 0.35 RECUP MATRlX 0.36 C OMP IN 5.0D 19.63 1544 124 0.215 0. 18 1.3 0.23 Sum psi 0.94 Sum % 0.72 RECUP BOTSIDE T32A OU 28 .6 0.0465 0.32 1.3 0.42 6.5D 33.1 1660 RECUP MA TRlX RECUP EXIT 8.0D 28.6 50.2 1132 0.0465 0.092 1 0.09 Sum psi 0.72 2.6 Sum % 4.25 % 4.5 COMBUSTOR 5.5 INTERCOOLER * Bead is defined as q=15.1 /(p* A ~2) 2.1.5. Combustor.

The HPRTE demonstrator required the design of a completely new single can combustor to replace the annular reverse flow multiple injector combustor of the production T63T32A. This new single can combustor would also be conveniently sized to first allow testing in the normal gas turbine mode, followed by testing in the HPRTE mode.

Preliminary combustor performance parameters are listed on Table 2.3.

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I \ I I I Ta b le 2.3 C ombu st or Perf o rmance Par ameters. Pressure drop 4.5% Parameter Annular Can C anHPR TE Airflow pps 1.43 1.43 2.6 Fuel Flow pph III III 93 77 121 Inlet Pressure psia 77 Inlet temp F 492 492 1099 Exit temp F 1800 1800 1747 OlD inch 10 4.25 4.25 Volume cu in 270 12 8* 12 8* 6 6 6 5.29x lO - 2. 52 x l0 - HRR (dim ' Ies s) 2. 51xl0 - *excluding scroll Operating fuel/air ratios for the T62T32A at rated speed varied from .007 to .021. It was anticipated that the existing start acceleration fuel schedule and starter input torque should still provide a satisfactory start.

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Chapter 3. Experimental Apparatus In this chapter the Titan HPRTE test apparatus is described, including major components, instrumentation locations, and auxiliary equipment. Following the description of the engine, the control and data acquisition systems will be presented.

3.1. Titan HPRTE Engine Layout The original drawings of the Titan HPRTE system are presented in Figures 3.1 to 3.3.

The instrumentation locations are indicated on the system drawing that provides the clearest view for each case. The instrumentation types, specifications, and nomenclature used in the figure labels are shown below.

3.1.1. Temperatures There are total of 22 thermocouples on the Titan HPRTE and its support systems. The thermocouples are types K, T and J from Omega Engineering, Inc. as specified below in Table 3. 1.

Table 3.1. Thermocouple specifications.

Type Omega Part Number Maximum Temperature (F) JQIN-14U-12 J 700 K KQIN- 14U- 12 1600 T TQIN-14U-12 400 Thennocouple readings are indicated on an Omega DP-460 reader on the control panel as well as on the analog-to-digital computer data acquisition system.

3.1.2. Pressures Pressure data is recorded at 12 locations on the Titan HPRTE. Pressure readings are indicated on gages located on the control panel in the control room as well as on the analog-to-digital computer data acquisition system. Pressure differentials are read into the data acquisition system as well as manometers located in the control room. Pressure differentials are measured in 3 locations: (1) the high-pressure compressor, (2) the low- NAS AJ CR- 2001-210675 17 pressure compressor and (3) the burner. The first two are recorded by the data acquisition system as well as on a manometer. The burner pressure differential is recorded manually from an independent manometer. The gage designations and their specifications are shown in Table 3.2.

Table 3.2. Pressure gage specifications.

Gage Pressure Range Company (psi) Heise Gl 0-200 *G2 0-60 Heise G7 0-200 Acco Helicoid 0-200 Acco Helicoid G8 *reads a maximum of30 in-Hg vacuum The pressure transducers were all purchased from Omega, and all were calibrated with 8.0 VDC excitation. The output voltage range was 1 to 6 VDC. The Omega model numbers and measurement ranges are listed in Table 3 .3 . All transducers were housed in the overhead instrumentation patch panel in the ceiling above the test bays.

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Table 3.3. Pressure transducer specifications.

Differential pressures PX142-001DSV 0-1 PSID PX142-005D5V O-S PSID PX142 -030D5V 0-30 PSID Absolute pressures PX138-01SASV 0-IS PSIA PX138-030ASV 0-30 PSIA PX138-100ASV 0-100 PSIA Gage pressures PX242-060G5V 0-60 PSIG PX242-100GSV 0-100 PSIG 0-150 PSIG PX242-1S0GSV 3.1.3. Nomenclature for Instrumentation A code is used to designate the measurement at each location. The format is: W-XX"yz where W, XX , y, and z are chosen from: T = temperature P=pressure HP= high pressure LP = low pressure DP = differential pressure c = compressor t = turbine r = recuperator i = inlet x = exit N ASAI CR - 2001-210675 19 ___ _____ __ _ ____ ~ .. _____ . _________ .J ---- - --1 I

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I As an example, T-HPci designates Temperature-High Pressure compressor inlet. In some

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instances it was more convenient to create intuitive unique names , such as T-icool (intercooler inlet temperature), P-FILTX (filter exit pressure), or P-WGATE (wastegate \ \ valve inlet pressure).

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3.2. Experimental Installation The Titan HPRTE is shown installed in the Energy & Gasdynamic Systems Laboratory at the University of Florida in Figures 3.4 and 3.5. The rust figure is a photograph from the side of the engine, in which several key components are visible. The large vertical duct on the left, covered with foil-backed insulation, is for conditioned air to be supplied to the engine inlet if needed. It forms a tee with a horizontal duct which is open to the ambient just to the left of the picture frame, allowing air to enter the engine. Just downstream of the tee is the turbocharger, which is not easily seen in this view. Immediately to the right of the tee is the recirculation valve (with the circular handle and remote actuation cable attached). Behind it is a quick-acting valve to allow rapid depressurization in emergencies, not used for control. In parallel with it is the Fischer wastegate valve, not visible behind the recuperator hot-side exit duct. The recuperator is the large dull-colored metal box to the right of the air handler duct. The cold-side inlet duct is visible forming an arc above the recuperator. The cold-side exit duct which leads to the combustor was removed immediately after the last test to inspect the combustor liner, so it is not shown.

The combustor housing, a large short vertical cylinder, is in front of the right end of the inlet duct. Below the combustor is the rest of the Titan engine. Just to the right is a large vertical U-shaped pipe which is the HPC inlet duct, installed for this test sequence. The venturi is the dark section of this otherwise shiny duct, installed in the distal section. The proximal section of this duct contains a tee where a short horizontal section enters from the right. This allows an alternative flowpath for starting (not used yet), whereby opening a valve allows the Titan engine to operate in open-cycle mode while warming up. The large torpedo-shaped horizontal tube in the foreground is the intercooler. The exit is on the right, where a filter is installed in the large angular chamber at the base of the HPC inlet duct. The wooden box in the right foreground contains two size 8D truck batteries used for starting.

NASAICR -2001-210675 20 L __ The initial phase of testing in 1998 was hampered by high HPC inlet temperatures which caused bearing temperatures to approach dangerous limits . The intercooler was then augmented by a spray cooling system, which is visible at the left end of the intercooler, consisting of the spray nozzles, float chamber (small box in front of the frame) , solenoid and return line on the floor. Instrumentation leads are attached from the overhead patch panel, which is not visible in this view. The Rover test rig is in the background in a parallel test bay.

Figure 3.5 is a photograph taken from the right side of the engine, from the perspective of the previous photograph. The auxiliary air inlet, HPC inlet duct, and intercooler water connections are in the foreground, along with the water brake dynamometer (the black disk on the right) . The vertical flat, black panel is one of the two blast shields used to protect laboratory personnel in the event of a catastrophic rotor failure.

NASAlCR -2 001-210675 21 T- HPtx P-HPtx P-burnen T-L Pu3 ( _ ~_; . r · - r,:- -:..='- -- -- ----- --- ';~7.

1 ' I' 1 P· LPLi ;1 - .. -- - - - _ ... - --( ---- --=..... - - - -- ___ _ , ...c2 __ ---..='-- __ ---' Figure 3.1. Elevation view of Titan HPRTE with instrumentation locations.

Figure 3.2. Plan view of Titan HPRTE with instrumentation locations.

ASAlCR- 200 1-21 0675 P-HPri T-HPri Co ntrol Air F- 1 . , - LPCX \ - ~ - ~

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_ru~ ~~ . ,= - 'f 1\ .,>_ .' ~

~:::::::::l:=- _ =_ I J .. - J _ .::L ' ')

- : !. :::=>

IL

Figure 3.3 . End view of Titan HPRTE with instrumentation lo cation s.

N ASAlCR- 200 J-21 0675

____ _ ______ _ _____ J

Figure 3.4. Photograph of Titan HPRTE (side view) after Test Run 22.

NASNCR - 200 1 -2 10675 24

__ J

Figure 3.5. Photograph of Titan HPRTE (end view) after Test Run 22.

NAS NC R- 200 1- 2 10675 25

_J

3.3. Support Equipment The auxiliary equipment necessary to test the engine included the dynamometer system, the spray cooler system, laboratory water supply, wastegate valve and controls, oil heat exchanger, and data acquisition system. Each of these subsystems will be described in tum in this section.

3.3 .1. Dynamometer System The water brake dynamometer was loaned to the program by the Army Research Laboratory, Aberdeen Proving Grounds. It was manufactured by Kahn , and had specifications as li sted below for the dynamometer and supporting components. Note that the maximum power rating was 1000 hp; as wi ll be discussed later, the oversizing of the dynamometer led to considerable operational difficulties during starting.

DYNAMOMETER Model Number: 301-190-001 Serial Number: 1537 Operational Range Maximum Power 1000 hp Maximum Speed 4000 rpm Maximum Torque 2000 ft-lbf Water Supply Specific Water Flow Rate 0.067 gal/min hp Pressure 50 psig STRAIN GAGE LOAD CELL Model Number: 514-100-032 Serial Number: 1537 TORQUE AND SPEED INDICATORS Torque Reader: Omega Engineering, Inc ., DP41-S-A Speed Reader: Monarch Instruments, ACT-3 AS AlCR- 200 1-21 0675 INLET CONTROL VALVE Valtek Mark One Control Valve, Part Number: 514-100-032 Det. 20 ELECTRO-PNEUMA TIC POSITIONER Air Supply Pressure: 50-150 psi (80 psi Titan operating pressure) SHELL AND TUBE HEAT EXCHANGER (TURBINE OIL) Manufacturer: Young Radiator Company Model Number: F-301-EY-2P Serial Number: 326207 Part Number: 239951 Maximum Temperature: 350 F Maximum Pressure: 150 psig STA-RITE CONTINUOUS CENTRIFUGAL PUMP Model Number: ffiE-63HL Code Number: lE97M Power: 1 hp Speed: 3450 rpm WATER RESERVOIR TANK Capacity: 30 gallons Material : Polyethylene Figure 3.6 contains three photographs of the dynamometer with the major components labeled. A schematic diagram of the dynamometer water flow path is depicted in Figure 3.7 . A centrifugal pump is supplied from a reservoir in a closed-loop system. The pump discharge is controlled by two valves in series, one being a gate valve for course adjustment, and the other being a needle valve integral to the rotameter, used for fine adjustment. A globe valve on the return line from the dynamometer was used to create sufficient backpressure to avoid cavitation. The heat exchanger allowed the closed loop temperature to remain within allowable limits by transferring the engine output energy to the faci lity chilled water system.

NASAlCR- 2001-210675 27

___ . __ . _. _____________ J

(a) (b) I

I

I

I NASAlCR- 2001-2 10675 28 I

I I

1 __ ~ ~~~ __ ~ ~ ~ ___ ~ ____ . _____ J

(c) Figure 3. 6. Kahn water-brake dynamometer installed in the Titan HPRTE test engine.

NASA/CR- 2 001 -2 106 75 29 Dynamometer Wafer Path Kahn dynamometer R otameter Centrifugal pump o Dynamometer gate valve Control valve

X Drain valve

Return gl obe valve reservoir Chilled water in Chilled water out Figure 3 .7. Schematic diagram of dynamometer water supp ly system.

3.3 .2. Spray Cooler System The Titan HPRTE experienced bearing problems early in the program which were identified with the HPC inlet temperature, as is discussed later in Chapter 5. To allow steady-state testing of the engine, the intercooler effectiveness was augmented by a water spray system installed at the intercooler inlet. The water injection system consisted of a de-ionized water supply reservoir via one or two spray nozzles, and a drain system that allowed excess water to drain from the intercooler while maintaining positive pressure.

The drain system was a level controller built in-house for this purpose, consisting of an adjustable float switch coupled to a solenoid that opened the drain intermittently. Figure 3.8 shows a schematic diagram of the spray cooler control panel, and Figure 3 .9 is the electrical subsystem schematic. The component specifications for the complete water injection system are as listed below, and the operator instructions are presented in Appendix A.

NASNC R- 200 1- 21 0675 30 _ ~~ , _____ ,,_ .. ______ . ______ ._ . __ J Spray Cooler Control Panel Layout

o o

Sprayer 1 Sprayer 2

o

System pressure . I

'X · -- ·· ·

'. .'

N ozzle fl ow valve \/ (o n/o ff) Sprayer press ure High-range fl ow Low-range fl ow rotameter rotameter l"v .. . \I"v _ ·1 i <,,_ >.' 1 <,,_ y Rotameter bypass Flow rate control valve valve Figure 3.8. Control panel layout for the spray cooling system.

NASAlCR- 2001 -210675 31 Water Injection System Electrical Layout hot 0 Main power switch

0 Switch 1

pump Switch 2

I Float switch

v Solenoid 1 ~

Solenoid 2 *1 ' * Float switch

T "l~oid

Neutr ru __________________________________ ~ ________ ~ 1 ________ __

Figure 3.9. Electrical schematic diagram for the spray cooling system.

Water Reservoir A 100 gallon polyethylene tank (30 in. x 38 in. x 26 in.) is used for water storage for the water injection system. Three 112 inch diameter fittings were installed on the tank Two fittings are placed roughly 3 inches from the top of the tank and one placed 3 inches from the bottom of the tank. A two inch diameter vent was also installed on the top of the tank, Hoses Hoses were purchased from Fluid Power Components, Inc. The hoses are 112 diameter, synthetic rubber tube with one braid of high tensile steel wire reinforcement, with a of 2000 psi. Each hose has both ends fitted with Parkrimp 43 series working pressure fittings.

AS AlCR- 2001-210675 32 Centrifugal Pump A Procon centrifugal pump, capable of 250 psi delivery pressure, is used for water delivery to the spray nozzles. The pump is driven by a continuous 3/4 hp AC motor manufactured by GE Motors.

Pressure Switch To prevent over-pres urization of the system a pressure switch is installed downstream of the pump. If activated, water is recirculated back to the reservoir. The switch was purchased from Omega Engineering, In c. and has a maximum pressure rating of 250 psi.

Solenoid The pressure switch controls an Omega Engineering, In c. SV-501, 114 inch, normally closed solenoid. The sprayer solenoids (2) are Honeywell 0-230 psi normally closed solenoids. These solenoids are controlled via indicator switches located on the control panel and each can be activated independently.

Pressure Gages Two 0-300 psi Omega Engineering, Inc. gages are installed on the control panel for system pressure and sprayer pressure indication.

Rotameters Flow rate is measured using Omega Engineering, Inc. rotameter s. Two are installed on the control panel both with a 3000 psi maximum pressure rating. The first rotameter is used as a "high-range" rotameter, 0-4 gpm. The second is used for fme adjustments and is referred to as the " low-range" rotameter. It is this rotameter that is primarily used to control the flow rate to the spray nozzles.

3.3 . 3. Laboratory Water Supply Facility chilled water is available in the laboratory and was connected to a specialIy- constructed manifold for the supply and return connections to the intercooler and dynamometer heat exchanger. Figure 3.10 shows a photograph ofthe manifold with the key components and connection points labeled. Standard flexible hoses, specified above, are used to connect the intercooler and heat exchanger to the manifold.

NASNC R- 2001-210675 33 Figure 3.10. Chilled water supply manifold.

3.3.4. Waste gate Valve and Controls Prior to Test Run 17 , control of the turbocharger speed was available only by using the quick-acting valve shown in Figure 3.4. This proved inadequate, so a Fisher control valve was installed in parallel to allow finer control of the turbocharger. Figure 3. 11 shows three views of the wastegate control valve and the by pa ss line (wastegate flowpath). The Fisher va lve is pneumaticall y controlled, as indicated in Figure 3.11c. Visual confirmation of the percent open status of the valve is available, as shown in Figure 3 .11 c.

NAS A/ CR - 200 1- 210675 34 I

______ ~ . _._~ ___ . __ . __ J

(a) (b) NASAlCR- 200 1-2 10675 35 (c) Figure 3. 1l. Wastegate va lve, used to control turbocharger boost.

3.3.5 . Oil Heat Exchanger The Titan engine is equipped with a water-cooled heat exchanger to maintain the oil supply within safe operating limits. A photograph of the heat exchanger installation is shown in Figure 3.12.

NAS AJ CR- 200 1- 2 10675 36 l~ ~ _ _ ____ ~ __ _ _ _ _ _ _ __ __ _ _ ___ ~ -_ _ ___ --- - ~ - - - - - - ~

Figure 3.12. Titan engine oil cooling system connected to facility chilled water reservoir.

3.4. Instrumentati on The specifications for the themlocouples and pressure transducers are given above in Section 3.1. The details of the data acquisition and gas analysis systems are presented below.

3.4. 1. Data Acquisition System The data acquisition hardware is built around a generic minitower PC using an Asus motherboard, a Pentium I 90 MHz processor and a 1.2 GB IDE hard drive. The operating system is Windows 95 . All peripheral circuit boards for data acquisition were purchased from Computer Boards, Inc., and the software engine is Labtech Control, version 10.01.

A CIO-DAS 1602112 analog-to-digital daughterboard is used for converting analog input signals. It has 16 channels of 160 KH z, 12 bit conversion with 32 digital I/O channels.

Four CIO-EXP-32 multiplexers have been connected to the AI D input to convert each of eight channels into 10 's ingle-ended' common-ground analog input channel s. Each of these mu ltiplexers has been dedicated to support a separate sensor group - types J,K, and NAS NC R- 200 1- 21 067 5 T thermocouples are supported separately, and the fourth multiplexer board supports other analog inputs, such as pressure transducers and strain gages.

A CIO-CTROS counter/timer board has also been installed on the motherboard. This S- channel, 16 bit board is used for pulse and frequency measurement. A CIO-MINI37 terminal outboard is used for signal connections which has a small breadboard section for additional signal conditioning.

The operation of the data acquisition software is well-documented in the Labtech Control manual, which is a commercially available software package, so the reader is referred to that source. However, the set-up file is the user-defined set of specifications which control what data channels are scanned and how the real-time data reduction is carried out. The Labtech Control set-up file is presented in Appendix B.

3.4. 2. Gas Analysis System The gas analysis system used III the HPR TE demonstration program is shown schematically in Figure 3.13 and photographically in Figure 3.14. Continuous gas samples can be taken from either upstream or downstream of the combustor. The sampling location is controlled by a solenoid value, which in tum is actuated by a remote signal from the control room. The sample leaving the solenoid valve passes through a dryer (IMR SOOP) which removes the water content of the sample. The pressure is measured downstream of the sample valve to ensure that damage does not occur to the equipment via overpressure. The sample is fed both into a Portable C0 2 meter (Model RI-411A) and the emission analyzer (COSA 6000 H). The readings from the COSA 6000H can either be monitored manually or by a data acquisition system (MRU- GRAPHICS PROGRAM operating on the IBM Value Point). A smoke opacity measuring device was located on the engine exhaust duct.

NAS A/CR- 2001-210675 38 High Pressure Return from Recouperator

I

~ .......................... .... .....

Burn er Inlet Sample LowP ressure Exhaust to Rec ouperator -= Turbine Outlet Sample Solenoid Valve

I I Pressure Gauge

COSA

RI-411A 0

6000 HC IMR 500 P Dryer Gas Analysis Data Acquisition Figure 3.13. Gas analysis system schematic.

NASNCR-2001-210675 Figure 3.14. Gas analyzers: COSA6000 (top) and RI-411A infrared CO analyzer.

COSA6000HC The COSA 6000 HC (COSA Instrument Corporation) portable emission analyzer provides reliable measurement of engine exhaust gas emissions. Electrochemical sensors measured O CO, NO and CH content of the sample. The CO content is calculated by 2, x, x 2 the microprocessor assuming dry air.

Analyzed Gases and Measuring Ranges: Oxygen 0-20.9% Carbon Monoxide (High) 0-30,000 ppm Carbon Monoxide (Low) 0-8000 ppm Carbon Dioxide 0-25 % (calculated value) Nitric Oxide 0-4000 ppm Nitrogen Dioxide 0-1000 ppm NOx ( NO + N0 0-5000 ppm 2) Hydrocarbons 0-6.00 % NAS NC R -2 001-21 0675 40 Accuracy: O +/- Reading or 0.1 % CO (low), NO, O +/- Reading or 2 ppm CO (high) +/- Reading or 5 ppm CH +/- 10.0% Reading x Sensitivity: Sensor Type: O 0.1% Electrochemical CH 0. 01 % Catalytic Bead x CO (low), NO, N0 1.0 ppm Electrochemical CO (high) 10 ppm Electrochemical Response Time: O Approximately 15 seconds CO Approximately 30 seconds NO Approximately 30 seconds N0 Approximately 60 seconds CH Approximately 90 seconds x RKI Instruments Model RI-411A CO Analyzer The Model RJ-411A Infrared CO (RKJ Instruments) indicator was used to measure the CO content of the sample. This provides redundancy to the CO content calculated 2 2 above.

Range: CO 0-19.9% Sensitivity: CO +1 -0.1 % IMR 500 P Dryer The IMR 500 P Dryer (IMR) removes the water content from the exhaust being sampled and supplies a "dry" sample downstream for proper analysis.

NASAlCR-2001-210675 41 MRU-Graphics Program The MRU-Graphics program made it possible to obtain a graphic and textual report of the measurements which were made by the COSA 6000 H. The emissions analyzer and the IBM Value Point running the MRU-Graphics software communicated via a RS232 connection. Several options were available for the data storage. For the HPRTE demonstration program, the data was stored in a test format which allowed it to be read into a spreadsheet of choice.

Smoke Opacity Meter The Model P-6IL Smoke Opacity Meter (Robert H. Wager Co. , Inc.) provided an accurate means of measuring the opacity of smoke being emitted through in the exhaust.

The measurement of the opacity was accomplished by passing the exhaust plume between a light source and a photo sensitive receiver with the resulting smoke density appearing as a percent opacity ofthe meter.

Range: Low 0-20% High 0-100% Accuracy: High 0-20% The procedures for setup and operation of the gas analysis equipment are presented in Appendix C.

NASAlCR-2001-210675 42

J

I

Chapter 4. Operating Procedures for the Titan HPRTE The operation of the Titan HPRTE in its final configuration (Test Runs 17-22) is described in this chapter. This includes the pre-test integration of the facility with the engine, the necessary preparation of the support systems, instrumentation checkout and preparation, audio-vi ual recording system setup, automatic data acquisition, and the job descriptions of the operating crew. Following these pre-test procedures, the operating instructions for engine testing are described, including startup, bringing boost and recirculation levels to the design point, and shutdown.

4.1. Support Systems Prior to the test run, an extensive check on the engme, support systems and instrumentation was performed. Also, the test area was secured and caution signs were placed at all entrances to the laboratory and adjacent areas. Chilled water, supplied by the University of Florida facility, was used in the intercooler and the turbocharger oil cooler.

All hose connections were checked for leaks and tested to ensure flow. Prior to test runs, the main chilled water supply and return valves were turned on and air was purged from the lines, the intercooler, and oil cooler.

Wastegate control air was supplied to the wastegate control knob on the control panel in the operations room via a flexible hose. Pressurized shop air at 30 psi was supplied to the regulator at the back of the control panel. Air supply was ensured by checking that the "Shop Air Supply Pressure" gage in the front of the panel read a minimum of 20 psi. The "Fisher Control Isolation" valves were checked to ensure they were in the closed position. The turbocharger was controlled by the "Boost Control" knob on the control panel. Exhaust gases were bypassed past the low pressure turbine (LPT) and the amount of flow bypassed was controlled by a Fisher valve. Prior to the test run, verification of the operation of "Boost Control" was done to ensure that a 0 - 20 psi range could be achieved.

ASA/CR-200 1-21 0675 43 Two size 8D , 12 volt batteries were used in the Titan HPRTE for starting. Prior to the test run, the electrolyte leve ls in the batteries were checked. The electrolyte should be just covering the battery plates. Distilled water is added if leve ls are low. Voltages are measured across each battery and across the starter terminals and should read a minimum of 12.6 volts for each battery and a minimum of 25 volts across the starter terminals. The battery check can be performed a day in advance to provide sufficient time for charging the batteries if necessary. Any loose connections of the leads are a potential fire hazard and each was checked and tightened if necessary.

Fuel and oil levels were checked and refilled as necessary. A seven gallon fuel storage tank. is located under the fuel cabinet and was filled prior to start. The fuel transfer pump was checked for proper operation. Any and all fuel leaks were recorded and fixed.

During test runs it may be necessary to refill the seven gallon tank. , so a full five gallon fuel can is standing by in the fuel storage cabinet. Oil used for the Titan HPRTE is Exxon 2380. The oil level for main engine was checked by pulling the dip-stick and ensuring that the level read "ful l. " The turbocharger oil sump tank level was checked in the same way.

4.2. Engine Room Preparation To ensure good ventilation of the engine room, the bay door was opened half-way and the laboratory ventilation fan was turned on. There was an option to use conditioned air at the engine inlet; however prior to the run the air handler switch was checked to be in the " off position.

to ensure that no foreign objects were All debris was cleared from the surrounding area ingested by the engine. The test area was secured by attaching a safety chain in the hallway leading to the laboratory entrance and posting caution signs at all entrances. The safety policy is that all visitors are checked in and supplied with ear and eye protective gear. Any late or unannounced visitors are directed to return after the test run is complete.

Fire extinguishers were placed in locations easily accessible to personnel and scatter shields were positioned on both sides of the core (Titan) engine.

NASAlCR-2001-210675 44 I

L

4.3. Data Acquisition and Gas Analysis Both the analog and digital data acquisition systems were checked to ensure proper operation. All thermocouples were checked for proper readings (no "off -scale").

Temperature values were read from an Omega DP-460 reader and measured near ambient temperature prior to the run. All pressure lines were pressurized to check for leaks and proper readings at the appropriate gages. Also, labels for the various data sets were checked for readability and to make sure they coincided with the correct locations on the engine. These checks were performed in advance of the run to allow time for correcting any problems which may have been found. Both gas analysis and the Analog-to-Digital system have separate set-up procedures and will be explained later.

4.4. Video and Audio Recording The video camera was set up on a tripod in the laboratory and connected to the video/audio cables which run to a monitor in the operations room. A new video tape was used for each run and was placed in the camera. A system check was performed to ensure proper visual and audio operation.

4.5. Operational Options In future tests it may be desirable to obtain data at other boost pressures, engine speeds, or recirculation ratios. Testing at lower boost pressures than the 2: 1 ratio specified for this program is achievable simply by adjusting to wastegate valve to a more fully-open position. The wastegate setting and the dynamometer setting are the only two independent controls available to the experimenter, since the turbocharger geometry is fixed. The recirculation ratio is set by the size of the turbocharger and by the speeds of the Titan engine and the turbocharger. Opening the wastegate alone would not only increase the turbine inlet temperature (to hold the same load), but would also increase the recirculation ratio simultaneously. The recirculation valve should not be considered as a means of controlling recirculation independently because it does so at the expense of power and efficiency. That is analogous to throttling the inlet of a conventional turbine in order to vary power, rather than controlling the fuel flow .

NASAlCR-200 1-21 0675 45 , ___ J In order to test other recirculation ratios, there are two options. One is to operate the Titan at lower speed by changing its governor setpoint. This has the disadvantages that the vibrational characteristics of the engine at lower speeds are unknown, and the lower speed would produce lower HPC pressure ratios, decreasing power and efficiency. The second option is to replace the turbocharger by one of a different size. This appears to be the only viable option for altering the recirculation ratio of the engine (because of the fixed flow area turbocharger), consistent with maintaining high efficiency.

NASNCR - 2001-210675 46

___ ~. __ J

Chapter 5. ExperimentaJ Data

In all test runs of the Titan HPRTE, data were acquired using the automatic data

acquisition system dri v en by Labtech Control software. The majority of test runs resulted

in premature shutdown of the engine as the startup procedures were developed and

hardware and instrumentation issues were resolved. After a number of trial runs to

acquire operational experience and to assure data acquisition fidelity, three test runs

resulted in significant data: Test Run 16 (April 1998), Test Run 20 (April 1999), and Test

Run 22 (May 1999). Prior to Test Run 20 the engine was configured, for reasons

described below, to include spray cooling to augment the intercooler effectiveness,

improved wastegate control, inlet temperature control, and the low-restriction HPC inlet

duct with venturi . The data collected by the Labtech software was exported into a

spreadsheet format and reduced further in Microsoft Excel. Appendix E contains the

spreadsheet output for each of the three tests, including all raw data such as temperatures, pressures, speeds, and torque. Reduced data is also included such as mass flow rates, power, recirculation ratio, pressure ratios, temperature ratios, and component efficiencies.

The data is presented in the next chapter in graphical form both as a function of time and

as cross plots, where appropriate.

Test Run 16 was the culmination of a series of tests in which steady-state operation was

prevented because of overheating of the Titan spool roller bearing. The bearing was

instrumented and the operator shut the engine down when the safe bearing temperature

was reached, 270 O F. The approach to the maximum bearing temperature was

accompanied by HPC inlet temperatures close to the design maximum as well, as seen in

Figure 5.1. In Test Run 16 , the maximum EOT was nearly achieved at the same time.

Clearly, there was not a possibility of operating for a longer duration in order to complete the recuperator thermal transient, indicated by the upper curve in Figure 5.1 , and achieve

steady-state. This was the motivation for modifying the engine prior to Test Run 17 to

include augmented HPC inlet cooling, a reduced-restriction crossover duct (HPC inlet

duct), and conditioned air augmentation at the engine inlet. Control difficulties also

NASAlCR-200 1-21 0675 47

_ __ . _____ I

_

dictated that an improved turbocharger wastegate subsystem be designed and

implemented.

Dynamometer load was gradually increased during Test Run 16 up to a maximum of

49 HP, achieved just before the shutdown based on high bearing temperature. It appears

that the maximum fuel flow capability may have been reached during this test, seen in

Figure 5. 2. The figure shows the engine speed decreasing at the same time that the fuel

flow rate reached a plateau. A more extreme case of this effect was observed in Test

Run 22, described in detail in Chapters 6 and 7. It is believed that the maximum power

output was limited by the fuel system rather than by the exhaust gas temperature (EGT).

Component efficiencies for the HPC and HPT (i.e. , the Titan turbomachinery

components) were calculated for Test Run 16 from measurements of pressure ratio and

inlet and exit temperatures, and are presented in Figure 5.3. After most of the initial

transient was finished, both component efficiencies stabilized to values of approximately

67 % and 84% for the HPC and HPT, respectively. The calculations were perfoIDled

under the assumption that total leakage was 5%, and that the average specific heat ratios

for the compressor and turbine were 1.38 and 1.35, respectively. It is expected that the

calculated HPC compressor is somewhat low because the heat gain to the HPC discharge

flow from the combustor exit duct is ignored.

For Test Run 22, a similar trend was observed at the outset, as shown in Figure 5.4.

However, in this case a water spray cooler was installed at the inlet of the intercooler, and when it was activated, the calculated component efficiencies changed. It is believed that a small fraction of the injected water escaped the de-misting screen/filter and was carried to

the inlet of the HPC. Although only a few percent of the total flow by mass, the liquid

water evaporated during compression, so the process was non-adiabatic.

Emission measuring equipment was in place for all three test runs. The most complete

data set was acquired during Test Run 22 , which will be presented in Chapter 6.

NAS Al CR-200 1-21 0675 48 I

L~ __ ~~ ~ ___ _

500 - 1200

j 1000

350 T

+ 800

300 -'- LL -- T-HP ci Cl Q) - Hp ci Limi t :3- 600 -- T-Bear ~ :J - Bear i ng limit ~ -- T- HP rx Q) a.

E Q) f-

~ -------+ -I ~--------------------------------------~~------ t 400

100 -1- + 200 50 + --t-~ 0 0 -1-- 0:00 :00 0:02 :53 0:05:46 0:08:38 0: 11 :3 1 0:14:24 El a p se d Ti me Figure 5.1. Variation ofHPC inlet temperature, bearing temperature, and combustor inlet temperature for Test Run 16 .

NASAlCR- 200 1- 21 0675 49 ---------- - ----- - --- 4000 140 ~ P M2 rpm ~ Cl.

- FUEL Ib l hr a: 1000 -- o o -j!-- ----- --- --- 0:0 :00 0: 01:26 0:02 :53 0:04 :19 0:05:46 0:07 :12 0:08 : 38 0: 10 : 05 0:11:31 0:1 :58 -500 --'----- ____________________________ ---1 -20 Elapsed Time Figure 5.2. Fuel flow and output shaft speed for Test Run 16 . The engine speed is higher by the gear ratio of 19 .5:1.

NASAlCR- 20 01 -2 10 675 50 HPC and HPT Efficiencies , Test Run 16 I- HPCeft - H PT eft I 12 0 \

\

i .,..

1. ~

\ ..r r

"" u

'"'

c: Q)

/ ~ ~ II

u ~\ Jl! "

w

f

'" ,t " ,§f MJj; :ri 0 :2 3:38 0: 2 5 : 05 0 : 26:3 1 0 : 27 : 58 0:15 : 00 0 : 16 : 26 0 : 17 : 53 0: 19: 19 0 : 20 : 46 0 : 22 : 12 T i me Figure 5 .3 . Calculated component efficiencies for Test Run 16.

NASAlCR- 200 1- 21 0675 51 HPC and HPT Efficiencies, Run 22 I- HPCeff - HPT eft I >. 80 t) c

'" t)

ill 60 o ~",,-- --,--:;;. :te,--=- ---.-"'-' 0:00 : 00 0 : 02:53 0 : 05 : 46 0 : 08 : 38 0: 11 :31 0: 14 : 24 0 : 17 : 17 0: 20: 1 0 0 : 23:02 0 : 25 : 55 Ti me (5)

Figure 5.4. Calculated component efficiencies of the high pressure compressor and

turbine for Test Run 22.

NASA/CR-2001-210675 52

Chapter 6. Experimental Results

In this chapter, the reduced data are presented for the final test series, which culminated in Test Run 22 , as well as for the preliminary test series. The data are shown in two types of presentation: temporal data, in which the time history of the measured or calculated data are shown, and crossp lots, which show the influence of two or more variables on each other. The final test series resulted in steady-state data, complete with power output and gas analysis information, so those data are presented first. Preliminary data and a brief chronology of testing will be presented in Section 6.2.

6.1 Final Test Series Results Following modification of the HPRTE to include spray cooling, the larger crossover duct, direct mass flow measurement of the HPC flow, the capabihty of cooled inlet air, and finer wastegate control, a series of shakedown tests was performed. Test Runs 17-21 were necessary to work out operational procedures and to correct instrumentation and flowpath deficiencies. In addition, the procedure necessary to reach steady-state operation was worked out. Prior to Test Run 22 the tests were prematurely terminated either by the option of the operator or by automatic shutdown due to the control system sensing high exhaust gas temperature, EGT (State 7 in Figure l.1). Steady-state data were obtained during Test Run 22, so the emphasis in this report is on that data set.

Two categories of data will be presented for this test: performance-oriented data, and emissions data. The performance data will be shown first and includes temporal variations of state point variables, engine parameters, and reduced parameters. Reduced data cross-plots will then be shown which indicate key performance characteristics such as specific fuel consumption (SFC) as a function of percent power. The emissions data will then be presented in Section 6.1.2.

NASAlCR-200 I -2 I 0675 53 6.1 . 1. Test Run 22 Performance Data The approach to steady-state operation can be seen in the temporal graphs of the bearing temperatures and component inlet/exit temperatures, shown in Figures 6.1 to 6.6. Figure 6.1 shows the bearing temperature in the uppennost curve, with the initiation of spray cooling indicated. It is clear that the increased cooling effectiveness had a dramatic effect on the bearing temperature, which is cooled primarily by HPC discharge air bleed. The augmented cooling was initiated when the bearing temperature reached approximately 230F, forty degrees below the safe limit. As shown in the figure, the bearing temperature dropped within a few seconds to a stable range well below the limit. Simultaneously, the HPC inlet temperature stabilized, fluctuating thereafter only in response to changes in engine load or turbocharger boost. In contrast to the previous test series (prior to Test Run 17), the stable bearing temperature allowed sufficient operating time for the large thennal transient associated with the recuperator mass to pass. The recuperator reached operating temperature prior to engine loading, so that all of the subsequent data were essentially steady-state values.

The HPC discharge temperature (recuperator inlet temperature) is shown in Figure 6.2, along with the HPC inlet temperature for reference. Upon initiation of the spray cooling, the discharge temperature substantially decreased. While a small increase in the compressor specific speed may partly explain this phenomenon, carryover of fine water droplets is the most likely explanation. Less than 10% liquid by mass would be sufficient to cause the decreased HPC discharge temperature observed, which indirectly enhanced the bearing cooling as noted earlier. It had been expected that the wire mesh filter installed downstream of the intercooler would serve as a de-misting screen, minimizing liquid carryover. However, during the shakedown trials, Test Runs 17 and 18 , there were leaks observed in the newly-installed crossover line (which were then repaired) which allowed mist-laden gas to escape from the region downstream of the filter, so visible quantities of liquid water are known to have been present at that point. Even though the spray cooler water flow rate was halved for subsequent tests, including Test Run 22, it is NASAlCR-200 1-21 0675 expected that fine liquid droplets still escaped the filter, leading to the decrease in HPC exit temperature.

There were two dips in the HPC discharge temperature that occurred after initiation of spray cooling, evident in the top curves in Figure 6.2 (elapsed time intervals 12 :45 to 13:08 and 15:08 to 15 :32). These correlate with small dips in the HPC inlet temperature, and occurred when the turbocharger wastegate valve was temporarily opened due to incipient LPC surge. The decrease in boost pressure lowered the HPC mass flow rate by almost a factor of two because of the decreased gas density. Since the spray cooler water flow rate was constant, it was more effective in lowering the HPC inlet temperature, and allowed more water mist carryover into the HPC. The latent heat of the additional liquid water fraction caused a decrease in the HPC discharge temperature. It should be noted that this effect is essentially the same as water fog injection used in some terrestrial power plants for power and efficiency augmentation. In several HPRTE applications, the design recirculation ratio would be sufficiently high that water for this purpose could be extracted from the engine if desired, rather than requiring a separate water supply. This option will be discussed further in Chapter 7.

The LPC inlet and exit temperatures are presented in Figure 6.3 and show the expected trend. The lower curve represents the ambient laboratory air temperature which rose slowly during the experiment due to heating from the engine. The speed of the LPC spool was slowly increased after startup by controlling the wastegate valve setting. The slow increase in LPC discharge temperature in the curve labeled "T -LPcx" is due to the increased LPC pressure ratio which accompanied the speed increase. Thereafter, the temperature ratio remained nearly constant, except for the two events noted earlier in which the speed was temporarily decreased due to incipient surge.

Downstream of the LPC, the recirculated flow is mixed with the LPC discharge air. The temperatures entering and leaving the mixing junction are shown in Figure 6.4. The upper curve is the temperature-time history of the recirculation flow (which is essentially the same state as the LPT inlet and recuperator exit). The initial recuperator warmup transient NASAl CR - 200 1-21 0675 is apparent, followed by small temperature changes that depended on the ambient air temperature, then the engine load. The second curve is the temperature of the mixed flow as it exited the mixing junction. The third curve is the LPC exit, as shown earlier in Figure 6.2. The exit temperature is nearly the mass-weighted average of the two inlet temperatures. The two spikes in mixed flow temperature correspond to the dips in the LPC exit temperature, which initially seems counter-intuitive. However, the cooler LPC flow rate decreased significantly during those excursions, so that the mass-averaged temperature rose, as it should.

A comparison of the second and fourth curves in Figure 6.4 indicate the effectiveness of the intercooler. The HPC inlet temperature (intercooler exit) remained nearly constant after the water spray augmentor was initiated, in spite of the intercooler inlet temperature variations caused by engine control changes.

The HPT exhaust temperature is shown in Figure 6.5. The control normally controls the maximum EGT to be 11S0F , at which point the engine automatically shuts down. The maximum temperature limit is over-ridden during start, as seen in the early part of the curve. The second peak occurred before the low pressure spool was allowed to spin up, so the engine load of about 25 HP was sustained via high EGT. Afterward, the 2: I boost allowed the same load to be sustained with an EGT approximately 200 F lower. Also varied during this initial period was the recirculation ratio, since the recirculation valve was slowly opened after the engine came to full speed. These two effects together allowed the EGT to decrease to the normal range. The small spikes midway through the test were the result of decreased boost alone, as the recirculation valve remained open during those events. Finally, the slow rise in EGT near the end of the test was due to increased dynamometer loading. As will be shown later, limitations in the engine fuel system prevented testing at maximum EGT, but a useful extrapolation to that point was performed.

Figure 6.6 shows the LPT inlet and exit temperatures, as well as the temperature ratio.

The early engine-temperature transient is again evident, as are the two spikes caused by NASAl CR-2001-210675 56 opening the turbocharger wastegate. Otherwise, the temperature ratio in the latter part of the test was nearly constant , which was consistent with the intent in controlling the pressure ratio to be 2: 1.

The next series of figures shows the variation of pressure at the various state points in the engine. Figure 6.7 displays the inlet and exit pressure of the HPC, along with the pressure ratio . The lowest curve shows the HPC inlet pressure. The inlet pressure was slowly increased toward the goal of 2 atm absolute by gradually closing the wastegate valve, which caused the turbocharger to rotate faster. The two events where the wastegate was re-opened partially are clearly visible as dips in this curve. The latter third of the test, in which the power data were obtained, had a steady inlet pressure near the design value.

After the initial thermal transient, the pressure ratio was nearly constant, except for the two transients associated with the wastegate adjustment. The HPC inlet temperature dropped during those events, so that the corrected speed increased. Increased mist loading also caused an increased "continuous intercooling" effect within the HPC , resulting in momentarily greater compression ratio. The final quarter of the test showed a decrease in pressure ratio due to decreasing speed of the HP spoo l, as will be shown later.

The LPC pressures are shown in Figure 6.8. Independent control of the LPC pressure ratio was achieved during the experiment by varying the wastegate valve setting. Figure 6.8 shows that the pressure ratio was less than one during startup, because the wastegate was fully open. In that state, the turbocharger speed was so low that the LPC acted as a throttling valve rather than as a compressor. The pressure ratio increased slowly as the wastegate was gradually closed until a vibration believed to be incipient LPC stall began.

By decreasing boost twice, the operator was able to work carefully past the stall regime to a stable operating point with a 2: 1 pressure ratio. The final portion of the test shows a slight overpressure of about 5% due to the decrease in HP spool speed and the subsequent LPC flow rate dec r ease and LPT inlet temperature increase, which caused the turbocharger speed to increase.

NASA/CR- 2001-210675 57 The LPT inlet pressure varied according to the engine operating point as shown in Figure 6.9. Except for a deviation during the early temperature transient, the turbocharger turbine pressure ratio mirrored that of the compressor shown in Figure 6.8 .

The recirculation ratio is a key parameter for semi-closed cycle engines since it affects the specific power, the relative size of the LP spool compared to the HP spool, and the combustion environment. The Titan HPRTE was equipped with venturi flowmeters at the two key locations : the inlet of the LPC, and before the inlet of the HPC. The recirculation ratio R was inferred from the data as the ratio of flow rates (wHP c/ wLP c) minus one. The lower graph in Figure 6.10 shows the variation of R with time for Test Run 22. The upper graph utilizes the same data, but represents the variation of R where 20% leakage occurs before the HPC venturi. This is presented as an extreme case; the likely leakage was less than 1 % after the shakedown tests. However , it is clear that the sensitivity to leakage is not extreme, and that the lower curve can therefore be taken as accurately representing the recirculation ratio. The equilibrium value of R was less than 0.6, which was less than the design value of 1.0. The matching of the low pressure spool to the Titan engine produced the design pressure ratio, but was somewhat large, resulting in a lower R. Note that the design recirculation ratio was achieved during the first wastegate transient (elapsed time 12:45 to l3:08), so that the stability of the combustion system was demonstrated as desired.

The net engine power delivered is presented in Figure 6.11 , along with the LPC air flow rate (lbm/s). After startup, the power was approximately 25 HP, the minimum setting of the dynamometer due to cooling requirements. The power drifted slowly lower as the dynamometer water supply heated, lowering its viscosity. The water flow was fixed during the slow transient, but the lower viscosity caused less torque in the water brake dynamometer. After approximately 17 minutes of run time, the dynamometer control valve was gradually opened, increasing the power absorbed. As this occurred, the engine speed began to decrease, as shown in Figure 6.12. The maximum power output was 53 HP, at which point the engine speed had dropped to 92% of design. The mechanical ASNCR-2001-210675 - - -_ ._ - -- - - - -- governor under-speed setpoint was reached, and an automatic shutdown of the engine occurred.

The explanation for the observed engine shutdown relies on the fuel flow curve in Figure 6.12. Normal operation of the governor would have allowed increased fuel flow in response to the decreased engine speed. However, when the fuel flow rate reached approximately 100 lbm/hr, no further increase was observed. Post-test evaluation suggests that a second fuel solenoid is required, and that if such a system had been installed, then the maximum power design. point would have been achieved at normal operating speed. The details of this analysis will be presented in Chapter 7.

One anticipated feature of the HPRTE cycle is good part-load efficiency, characterized by a relatively flat curve of SFC versus percent power. Figure 6.13 presents the experimental variation of SFC for Test Run 22 over the range of power output recorded.

The estimated full power point for this test is estimated to be 153 HP, as will be presented below, so the percent power data are normalized by that value. For reference, the full power and 50 % power points of an unmodified Titan T62T32A APU are shown (open squares). Several observations can be made regarding this figure. First, the curve is not entirely flat, i.e. , the part-load SFC is higher than the full-power design point. This is to be expected based on the experimental method of control employed in this test. The ideal means to reduce power in the HPRTE system is to decrease the LPC pressure ratio, which in turn decreases the gas density in the high pressure portion of the cycle. In the present implementation of the HPRTE cycle, this means utilizing the wastegate valve as a throttle, allowing the temperatures in the high pressure components to remain essentially constant. If that were the case, the constant temperature ratios would specify constant pressure ratios, so the high pressure engine would operate at constant dimensionless operating point, hence constant efficiency. Only at very low power, after the LPC pressure ratio approached one, would the SFC increase as the burner temperature dropped. The experimental data presented in Figure 6.13, in contrast, are for the case of nearly constant LPC pressure ratio, so that the burner temperature decreased with decreasing power. Therefore, the experimental SFC curve is not flat, but instead NASAlCR-2001-210675 resembles that of a conventional recuperated cycle. Further experimentation would allow testing of the HPRTE at reduced LPC pressure ratio, so that the optimum SFC curve could be detennined and the control requirements specified for specific applications.

A simulation was performed for the optimal control case, using the simulation code described in Chapter 2, and is plotted in Figure 6.13 for comparison with the data. Even with the relatively low LPC pressure ratio of this demonstration engine, the predicted flatness of the SFC curve is significantly better than the experimental data indicate. A second simulation was performed in which the engine speed was held constant at the design value, rather than drooping as it did in the experiment (discussed below). While the flatness of this SFC curve is not as favorable as for the optimal control case, it is nevertheless considerably better than the experimental result.

The second observation concerning Figure 6.13 is that only the low power portion of the curve was obtained experimentally. This is due to the limitations in the fuel system discussed with Figure 6. 12 , which did not allow full power testing. A third observation is that the engine speed did not remain constant (see Figure 6.12), but monotonically decreased as power increased. The lower engine speed translated to lower HPC compression ratio, hence to a penalty in SFC. If the engine speed had been constant, therefore, the SFC curve would have dropped more rapidly with increasing power than is shown experimentally in Figure 6. 13 . Thus the part-load SFC values shown are somewhat pessimistic, and the flatness of the SFC curve, even at constant LPC pressure ratio, would have improved as shown in the constant speed simulation.

The same design-point model was used to simulate the performance of the Titan HPRTE at selected points representing high, medium, and low power. The results are included on Figure 6. 13 , indicated by triangles, and show excellent agreement with the experimental data. The simulation was performed by matching the experimentally-determined values of compressor pressure ratios, HPT exit temperature, HPC inlet temperature, recirculation ratio, air flowrate, pressure losses, and heat exchanger effectivenesses. Design values of the turbomachinery efficiencies were used, a limitation of the simulation software.

NASA/CR-2001-21 06 75 60 Leakage was included as a flow from the HPT exit to ambient, a coarse model imposed by the software. All other state points and flowrates were calculated in the simulation, including HPT inlet temperature. The results allowed calculation of power output, SFC, efficiency, fuel flowrate, combustion stoichiometry, and specific power.

The predicted value of power output agreed with the experimental values in each of the three cases within 5%. As may be seen in the figure, the SFC calculations agreed very well with the experiment for the medium and high power cases, and agreed acceptably for the low power case. Poor off-design turbo machinery efficiency in the low power case is the most likely explanation for the discrepancy in SFC. Another important result from the simulation was that the HPT inlet temperature increased only 70% as much as the HPT exit temperature as the power increased. This is due primarily to the HPT expansion ratio decrease with increasing load, which resulted in less temperature drop across the turbine. This conclusion supports the maximum power extrapolation to be shown later in this section.

As noted, the fuel system limited the power output to a value significantly lower than the maximum. The maximum power is normally determined by operation at the maximum EGT. In the case of the HPRTE, it is determined by operation both at maximum EGT and at maximum LPC pressure ratio. During the engine loading portion of Test Run 22, the LPC pressure ratio was nearly constant (Figure 6.8), but the EGT varied as shown in Figure 6.5. Figure 6.14 is a cross plot of the engine power output versus the EGT. As is usual for gas turbine engines, the power output was observed to be a strong function of the EGT over the test range. A regression analysis of this data was performed in order to extrapolate to the full-power design point. The analysis was performed using the regression tools built into Excel 97. Both linear and quadratic regressions were performed. In spite of the apparent curvature of the curve in Figure 6.14, the quadratic fit yielded unrealistically high power output values at the design EGT of 1150 F. Therefore, only linear fits are presented, which is consistent with the nearly-linear trend usually observed in conventional gas turbines.

ASAlCR -2 001-210675 Linear least-squares regression of the full set of data was first performed with the complete data set during the engine loading sequence, the same data set graphed in Figure 6.14. Both standard regression coefficients and upper 95 % statistics were obtained; the results are shown below in Table 6.1. The predicted power output at the design EGT rating of 1150 F is shown, along with the power output at 1215 F. The low- power part of the curve in Figure 6. 14 , showing a nearly constant power output over a range of EGT, is considered less reliable because of concerns about the thermal transient and the apparent low HPC efficiency at this off-design operating point. Since the higher- power portion of the is very likely to be m9re reliable, the regression statistics for the upper 95% of the data are included. The projected power output at 1150 F for this case is 153 HP, a figure that has been used in this report as representative of the full power output. A regression was also performed on all data above 30 HP in the data set, an alternative means of eliminating the scatter in the low-power data. The projected power was about III HP for an EGT of 1150 F or 151 HP at 1215 F.

Table 6.1. Results of regression analysis of power versus EGT.

Slope Intercept Case HP @ 1l50F HP @ 1215F Full data set 0.3892 -36l.1 86 .5 11l.8 upper 95% 0.4197 -329.7 153.0 180.2 Data> 30HP 0. 6181 -600.1 110.7 150.9 The rationale for choosing an intermediate value of 153 HP for the estimated design power output is based on several arguments.

1. The upper portion of the HP versus EGT curve is free from ambiguity in the thermal delay. The Titan engine response to increased load was to increase EGT almost

instantaneously « 1s response time). Therefore the regression of the upper data

appears the most reliable.

2. The projected output is conservative from the standpoint that the decreasing engine speed penalized the power output at higher power levels (the speed was only 92% of design at the maximum recorded power). This must have had the effect of decreasing ASAlCR-200 1-21 0675 62

- - "- - , .. _- --

the slope of the HP versus EGT curve, artificially lowering the extrapolated power estimate at the design point.

3 . The EGT control limit for conventional gas turbines is an indirect means of controlling the HPT inlet temperature, used because of the unreliability in measuring temperature at the combustor exit. The full power point corresponds to a certain gas composition , specific heat ratio , and HPT pressure ratio , so there is a direct correlation between the HPT inlet and exit temperatures. However, the HPRTE combustor inlet st ream contains significant concentrations of CO and H 0 due to recirculation, so the combustor exit stream has much higher exhaust product concentrations than conventional gas turbines. The specific heat ratio is decreased significantly, by about 0.03 for the Titan HPRTE. For a given pressure ratio, HPT efficiency, and EGT , a decrease in specific heat ratio corresponds to a decrease in HPT inlet temperature. For the conditions of the Titan HPRTE, this decrease is estimated to be 65 F. This effect is in addition to the effect of reduced pressure ratio, discussed earlier, which led to the conclusion that the HPT inlet temperature increase was about 70% of the HPT exit temperature increase. Therefore, controlling the HPRTE to the same EGT as the original Titan engine, 1150 F, is tantamount to controlling the HPR TE turbine inlet temperature to a value at least 65 F lower than the original engine. A fairer comparison would be to extrapolate the power to 1215 F, as shown in the final column of Table 6. 1. The value of 180 HP predicted by the upper 95% analysis may be optimistic, but the value of 151 HP based on the higher power points falls in line with the earlier projection of 153 HP. Therefore, that value was adopted as being the most representative projected power output at the design point.

6.1.2. Test Run 22 Emissions Data A reduction in exhaust emissions is an expected characteristic of the HPRTE cycle due to the inherent exhaust gas recirculation entering the combustor. The dilution of the oxidizer by combustion products has the same qualitative effect as steam injection, a well- established method of controlling emissions in terrestrial gas turbine power plants. The NASAlCR -2 001-210675 exhaust concentrations of NO CO, and O were continuously monitored during Test x, 2 Run 22, and the results are presented in Figure 6. 15 . The CO concentration peaked early in the test due to the start sequence, in which the fuel/air ratio is large in order to accelerate the engine. After startup, the fuel/air ratio remained fairly high in order to sustain the minimum dynamometer load before the turbocharger speed increased and before the recirculation valve was opened. After the recirculation valve was opened, the CO concentration dropped precipitously at about the 8:38 mark. This was due to the recirculated exhaust products reaching the combustor, as well as the increase in HPC inlet pressure shown earlier in Figure 6.7. The turbocharger boost (LPC pressure ratio) was gradually increased, accompanied be a decrease in CO concentration until reaching a stable value of approximately 5 ppm.

The variation of the NO concentration was more complex in the early portion of the test x due to the competing factors of burner stoichiometry and recirculation ratio changes, along with burner pressure and inlet temperature changes. After the initial transient, the NO concentration stabilized near a value of 30 ppm except during the two brief dips x associated with opening the waste gate valve (elapsed time intervals 12:45 to l3:08 and 15:08 to 15:32). Those two events show that the increased recirculation ratio caused a decrease in NO although the simultaneous change in burner pressure and temperature x, masked the effect partially. The temperature increase alone would have caused an increase, not a decrease, so the variation of NO concentration with recirculation ratio R x is shown to be quite sensitive. This is especially important since future prototype engines based on the HPRTE concept would be expected to operate with recirculation ratios four or five times as high as that of the present demonstration.

When the engine load was gradually increased, the NO concentration increased as x expected, reaching a plateau near 50 ppm until the under-speed automatic shutdown occurred. The recirculation ratio was low during the loading period because of the decreased engine speed, so the NO level was higher than it normally would have been x for this engine. Even so, the NO concentration was slightly lower than that of a x comparable APU operated under similar loading. Tests performed independently by VOC NASAlCR-200 1- 210675 64

L

Testing Inc. of San Bernardino, CA showed NO concentrations of 53 and 65 ppm at 30 x kW and 60 kW output, respectively. The lower figure of 53 ppm is shown as a horizontal line in Figure 6.14 for reference, along with the reference for CO concentration, 124 ppm.

Note that the HPRTE CO concentration measured was a factor of 25 lower than that of the reference engine under similar loading conditions.

It is important to note that the pollutant concentrations alone give a misleading picture of the impact of HPRTE design choices on traditional emission indices, based on emissions per time or distance depending on the application. On that basis, the emi sion index is proportional to the product of pollutant concentration and exhaust flow rate, that is, proportional to the total flow rate of the pollutant in the exhaust. The HPRTE cycle allows the inlet and exhaust flow rate to decrease by a factor of (1+R) for a given power output. For the HPRTE Titan demonstration engine in Test Run 22, the recirculation ratio R was near 0. 35 during loading, which corresponded to a 25% decrease in exhaust flow relative to the HPC flow. Therefore, even if the pollutant concentrations of the HPRTE and reference engines had been identical, the emission rate of th e HPR TE would have been 25% lower. Accounting for the observed reduction in emission concentrations in the test, the NO emission rate was about 30% lower than the reference, and the CO emission x rate was a factor of 33 lower (one and a half orders of magnitude). Future prototypes designed for high recirculation (R> 2) would doubly benefit due to the reduced pollutant concentrations and the reduced exhaust flow rates . Assuming the observed trends can be slightly extrapolated, it appears reasonable to expect NO emission rates to be reduced by x more than an order of magnitude and CO emission rates to be reduced by more than two orders of magnitude. These conclusions will be dependent upon the combustor design, but regardless of the design tradeoffs made, the results are highly encouraging.

NASNCR-2001-210675 250 ~-------------------------------------------------------------, Initiation of Spray

/

LL 150 O'J Q) E- - T-HPci deg F Q) - T-Bear deg F "§ - T-Filtr de~ F Q) a.

~ 100 -- I- 50 -- o ~------------------------------------------------------------~ 0: 00:00 0: 02 :53 0:05 : 46 0:08 :38 0: 11 : 31 0 :14 :24 0:17:17 0:20 :10 0:23 :02 0:25:55 Elapsed Time Figure 6.1. Bearing temperature and HPC inlet temperature versus elapsed time for Test Run 22.

NASAlCR- 20 0 1 -2 10 675 66 600 r------------------------------------------------------------. 2 1.8 C2 T Ratio 1.6 1.4 ~ 400 u. .Q Cl OJ Q> c: T-H P ri (C2 Exit) 1.2 :3- ~ ~ '" ~ - T- HP ci deg F '" OJ Q> ~ 300 ~ I- T- HP ri deg F E ~ - C2 T Ratio ~ 0.8 0 .6 --...

0 .4 0 .2 T- Hp ci (C2 Inlet) o 0 0:00:00 0:02:53 0:05 :4 6 0:08 :38 0:11 :31 0:14: 24 0:17 :17 0:20 :10 0:23:02 0: 25 :55 Elapsed Time (sec) Figure 6.2. HPC inlet temperatur e, exit temperatur e, and temperature ratio versus elapsed time for Test Run 22.

NASAlCR- 200 1-2 10675 67 300 .-------------------------------------------------------------------------, 1.4 1.2 250 1- -- o L- ______________________________________________________________________ ~ 0 0 :00 :00 0:02 :53 0: 05 : 46 0:08:38 0: 11 : 31 0:14:24 0:17 :17 0 :20 :10 0:23 :02 0:25: 55 Elapsed Time (sec) Figure 6.3 . LPC inlet temperature, exit temperature, and temperature ratio versus elapsed time for Test Run 22 .

NASNCR- 200 1- 21 06 75 68 -''''- T -l Pti -

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0:23 :02 0: 25:55 0:00:00 0:02:53 0:05 : 46 0:08:38 0: 11 : 31 0:14:24 0:17:17 0:20:10 Elapsed Time Figure 6.4. Mixing junction inlet and outlet temperatures versus el apsed time for Test Run 22.

NASAlCR-200 1-21 0675 69 ____ - .. _---_ ..

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NASAlCR-2001-210675 70 800 ~------------------------------ --------------------------------------~ 1 . 2 7oo +-----------------------------~ ~ ------------------ ==~~~~ ------~ 600 +---------------~~-- ~ ----------_P~--~A_----------------~i--------~ 0.8

~

500 +------------------, r- --~~--------~----~~-------------- ~ ------~ LL ~ Ol Q) a: T-LPti (T2 Inlet) ~ t':' Q) :::> 2 400 +----------------- ~~-------------------------------------- ~ ------~ 0.6 ~ Q) ~ 0.

2i E Q) E f- OJ r 300 +--------------- -+~------------------------------------------------~ 0.4 200 +------------- ~YL--------------------------------------------------~ 0.2 100 +-----------~~------------------ ----------------------------------~ O +-------~------~----~-- ----~------~------~------~------~------ + O 0:00 :00 0: 02 : 53 0:05: 46 0:08:38 0: 11 :31 0:14 : 24 0:17:17 0:20:10 0:23:02 0:25:55 Elapsed Time Figure 6.6. LPT inlet temperature, exit temperature, and temperature ratio versus elapsed time for Test Run 22.

NASA/CR-2001 -210675 71 120 6 ~ t? 80 a: 'iii ~ S :::J <Il ~ <Il :::J <Il ~ <Il Cl.

60 3 ~ N Cl.

() 20 ~------ r-~~~~~~~~J------\t-------------------~------+

o+--------- __ -------- __ --------~--------~--------~--------~o

o 100 200 300 400 500 60 0 T ime Figure 6.7. HPC inlet pressure, exit pressure, and pressure ratio versus elapsed time (s) for Test Run 22.

NASNCR-2001-210675

------- .. __ ._--------_._------ ------

35 ------ ..- ---.-.-. -. - ..- ... -.- -.-- - -- ...... . -- ... . ....... .-... .. ---- . -- - .. -- --. --- .. - .. - -- -. --- .. - -- . ... -.----- . _-- 2.5 ./'I.~~ ..........

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0.5 0+-----~------~---~----- __ ----- __ ----~-----~---- __ ----__40 0 :00 : 00 0:02 :53 0 : 05 :46 0 :08 : 38 0: 11 : 31 0:14:24 0: 17:17 0:20 :10 0:23 :02 0:25 : 55 Elapsed Time (sec) Figure 6. 8. LPC inl et pressure, exit pressure, and pressure ratio versus elapsed time for Test Run 22.

NASAlCR-2001 -210675 73

__

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+------------ __ ------ __ -- ---- __ ---- __ ------~------ __ ---- __ ------~o 0:00 : 00 0:02 : 53 0:05 : 46 0:08 : 38 0 : 11 : 31 0:14 :24 0:17:17 0:20:10 0:23 :02 0:25:55 Elapsed Time Figure 6.9. LPT inlet pressure, exit pressure, and pressure ratio versus elapsed time for Test Run 22.

AS N CR- 2001-210675 1_- I -R di rect -- R di rect leak = 0 .21 1 . 6 r--------------------------------------------------------------------~ 1.4 t---------------------------------+_----~--------------------------~ 1 .2+---------------------------------H-----~--------------------------~ ~ 0 . 8 +-------------+---+-~~----------H-----~----~~------------------~ 0 . 6 +-------------~~----------~--~~----+4----------~--~~--------~ 0.4 +-----------~+_--~~~------------------------~--------4_--------_4 0 . 2 +-----------~+_------------------------------------------4_--------_4

O+-----~----~~------~------~------~------ __ ---- ~----J- __ ------ ~

0: 00 : 00 0: 02 : 53 0: 05 : 46 0: 08 : 38 0: 11 : 31 0: 14 : 24 0: 17 : 17 0:20: 10 0: 23 : 02 0: 25 :55 Elapsed Time Figure 6.10. Recirculation ratio bounds versus elapsed time for Test Run 22.

NASNCR-200 1-21 0675 60 ,- ------------ -- -------- ------ ---- ------ ---- -- ---- -- ---------- -, 1.8 1.6 50 +-------------------------------~c_--~~------ ~ ~+_~4 ~------~ 1.4 Inl et Air Mass Flow 40 +----------- ~ --_+--------~--_++_--~+_------ ~ ---- ~ ~------__+ 1.2 0:: (J) (J) '" ~ ~ 30 +-----------~--~-------- ~ --~r_----r_---- ~ -------- ~------~ :;;: a; 0.8 E 20 +----------- r.~-------------- ~ ; T ~I~~I~~ ~---------- ~------__+ 0.6 HP 0.4 10 +---------~ rl- ---------------------------------------- ~------~ 0.2

0+-----~--~~~----~------~----~------~------~--~ __ ----_40

0 :00 :00 0 :02:53 0:05 :46 0 : 08 :38 0: 11 : 31 0 : 14:24 0 : 17: 17 0:20 : 10 0:23 :02 0:25 :55 Elapsed Time . 1 I Figure 6. 11 . Power output and inlet air flow rate versus elapsed time for Test Run 2 2.

I

I NASA/CR -20 01 -2 10 67 5 76 r----------------------------------------- ----------------------~ 120 ~ E ~ Fuel Flow g 0..

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3300+----- __ ----~ __ ------ __ ----~------ __ ----_,-- ---- __ --~-- __ ----_+·20 0:00 :00 0:02: 53 0:05:46 0: 08 :38 0: 11 : 31 0: 14 : 24 0:17:17 0:20: 10 0:23:02 0: 25 :55 Elapsed Time Figure 6.12. Output shaft speed and fuel flow rate versus elapsed time for Test Run 22.

The engine speed is higher by the gear ratio of 19.5: 1.

NASA/CR-200 1-21 0675 77 EXDerimental data Simulation with optimal control Simulation with

/

const. speed & boost ~ 0.

Simulation using / ~

=

exp 'l parameters U LL CfJ Reference APU

o

o 0 .1 0 .2 0 .3 0.4 0 .5 0 .6 0 .7 0 .8 0 .9 % Power Figure 6.13. Specific fuel consumption (SFC) versus percent power for Test Run 22.

NASN CR-2001 -210675 78 60 y----- ------------ ---------------------- -- ------------ -- --------------- iL ~ Q; "$ a.

Q) ~ I 20 - 10 +- ------------------------------------------------------------~------ ~ O+------- __ ----------------------~---- ---- ----------------------~----~ 990 1000 1010 1020 1030 1040 1050 1060 1070 E GT (T- HP bc , d ey F) Figure 6_ 14 . Power output versus exhaust gas temperature (EGT) for Test Run 22.

NASAlCR -- 200 1- 2 106 75 79 Tes t Run #22 Gas Analysis 350 r---------~--------------------------~------------------------~ 25 :2 CL CL 200 - x 'if'.

N --- m ~ 2- o<l --- NOx 0' - CO ret ~ - NOx re f --- 02 o L-------L--------------- ~======~~ ~ -- ~o 00:00.0 02:52 .8 05 : 45 .6 08:38 . 4 11 :31.2 14 :24 .0 17 : 16.8 20:09 . 6 23 : 02.4 25 : 55.2 Elapsed Time Figure 6.15. Exhaust pollutant concentrations versus time for Test Run 22.

NASAlCR- 200 1- 2 10675 80 · -- .- _. -"'--"-' -- -~ "- -- 6.2. Shakedown Testing Five shakedown tests were performed prior to achieving steady-state operation in Test Run 22 . The first test in the final engine configuration occurred 5 March 1999. In this test, Test Run 17 , less than four minutes of operation was possible due to high EGT shutdown before the recirculation valve could be fully opened and the wastegate valve closed. Instrumentation deficiencies were sorted out, and flowpath leaks between the intercooler and HPC were corrected. Test Run 18 was conducted on 9 March 1999. High EGT was still observed, so modifications to the dynamometer control valve system were made in order to increase the control sensitivity. Test Run 19 was on 11 March 1999, still very short due to high EGT. In order to diagnose the source of the high EGT, a decision was made to operate with the dynamometer de-coupled from the engine to determine whether the minimum dynamometer load was greater than the engine could sustain. Test Run 20 was conducted on 7 April 1999 with no load, and resulted in successful engine operation at idle conditions. Thus the hypothesis that the startup engine load had been too great was validated. This test also resulted in a sufficiently long run time that useful data were recorded; these are presented in this section. The final shakedown test, Test Run 21 , was conducted 27 May 1999 and was hampered by a defective battery. No engine start was achieved, so the battery was replaced, leading to the successful Test Run 22 on 28 May 1999.

Representative zero-load temporal graphs are presented in Figures 6.16 to 6.20. The HPC inlet temperature and bearing temperature are shown in Figure 6.16. The upper curve indicates the gradual increase in bearing temperature that occurred without spray cooling, interrupted when the turbocharger boost was initiated. The spray cooler was highly effective in reducing the bearing temperature, and simultaneously stabilized the HPC inlet temperature. The temperature behavior was very similar to that observed in Test Run 22.

The HPT exit temperature (EGT) is shown in Figure 6.17. Four regions are discernable.

First is the startup period in which rapid EGT changes occurred due to large changes in AS AlCR- 200 1-21 0675 81 fuel flow. Second is a gradual increase in EGT, due to the long thermal transient of the recuperator. Note that steady-state operation was achieved after approximately 300 s.

Third is a slight dip to under 1000 F when the turbocharger boost was applied. Fourth is a more significant dip corresponding to the start of spray cooling, which lowered the HPC inlet temperature. The observed temperatures were thus rather well-behaved.

Figure 6.18 shows the variation of the high pressure recuperator inlet and outlet temperatures . The temperatures did not plateau because the boost was continually adjusted and the laboratory air continued to warm from engine waste heat, causing the HPC inlet temperature to rise slowly. It appears that the recuperator thermal transient was essentially finished after 300 s, after which the inlet air heating was responsible for the further temperature rise.

The hot-side recuperator temperatures are presented in Figure 6.19. Note that the inlet temperature is the same as the EGT presented above. The increasing exit temperature reflects the increase in temperature noted on the recuperator cold side in Figure 6.18.

Finally, the temperature drop across the LPT is shown in Figure 6.20. As expected, the temperature ratio increased late in the test when the wastegate valve was fully closed.

NASAlCR- 2001-210675 82 30 0 r-----------------------------------------------~ Spray cooling initiated 250 1- ---- u:- 0> Q) 2- ~ - T-Bear - T-HPci ~ ~ E Q) I- o ~----------------------------------------------~ 100 200 300 400 500 600 o Time (sec) Figure 6.16. HPC inlet and bearing temperatures versus elapsed time for Test Run 20.

NASNC R- 200 1-2106 75 83 j -- ~--~ 1400 - LL en Q) E.

- T-H Pt x1 Q) - T-HPtx2 :0 co -T -HPtx3 Q; - T-HP tx avg c.

E Q) r- a a 100 200 300 400 500 600 700 Time (sec) Figure 6-17 _HPT exit temperatures versus elapsed time for Test Run 20.

NAS N CR- 200 1 -2 10675 84 1000 r-.---- .. .. . --.------- .. ----.-- .. ----- ...... --- .. -.--

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NASAlCR-2001-210675 85 1400 .--------------------------, u:- 0) Q) ~ ~ '" 1§ Q) Cl.

E Q) I- O~--~----_-------~---_----~--~ o 100 200 300 400 500 600 700 Time (sec) Figure 6.19. Recuperator hot-side inlet and exit temperatures versus elapsed time for Test Run 20 .

NAS Af CR- 200 1- 2 10675 86 ___ _ _________________________ ..J 70 0 u:- Cl Q) :s.

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NASAJCR -2 00 1- 210675 87 I

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Chapter 7. Data Analysis One important operational question from the final testing series involves the physical limitation that prevented the engine from being fully loaded, in spite of the moderate EGT and bearing temperatures. The performance of the governor is in question, due to the under-speed shut downs which have occurred during norn1al engine operation. It was first hypothesized that the governor was not responding to the load applied to the engine by the dynamometer due to mechanical friction or other failure. To investigate this possibility, two test runs were compared, one during April 1998 (Test Run 16) and one during May 1999 (Test Run 22).

First consider Test Run 16 which included the use of the dynamometer to provide an engine load, yet did not use spray cooling. The test run was terminated due to excessive bearing temperatures encountered eleven minutes into the run. Figure 7.1 below shows the engine speed and the fuel flow as a function of time during the test run.

As can be seen, the engine speed remained nearly constant at 3750 RPM (dynamometer speed) for the first seven minutes. The response of the fuel flow tended to mirror the engine speed; that is, for every peak in the engine speed there was a corresponding valley in the fuel flow curve. This trend is as expected, and shows that the governor was properly responding to the demands of the engine.

To better understand the performance from seven minutes elapsed time to the end of the run, it is helpful to see the load being applied to the engine. Figure 7.2 shows the horsepower generated by the engine as a function of elapsed time.

ASAlCR-200 1-21 0675 89 Titan Run 16: Fuel Flow and RPM 3800 140 :0 0- - RPM2 rpm 0:: 3550 60 - FUEL Ib l hr o 3350 +-~----------------------------------------------------~------~ 33oo+- L----.------~------._------------------_.------~----~._---- _+-20 0: 00 :00 0: 01 :26 0:02:53 0:04 :19 0:05: 46 0:07:12 0:08:38 0 :1 0:05 0: 11 : 31 0:12:58 Elapsed Time Figure 7.1. Test Run 16 output shaft speed and fuel flow versus time. The engine speed is higher by the gear ratio of 19.5 : 1.

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ASNC R -2 00 1- 21 0675 90 ote that in the first seven minutes of the run that the engine was only producing 30 to 35 horsepower. Following the seven-minute mark the load was increased to a maximum of 48 HP. It is in this region that the fuel control system began to fail. Referring to Figure 7.1, the response of the fuel flow to the load increased significantly during the 7-11 minute elapsed time period, but this increase was still insufficient to maintain the engine speed. Consequently, the engine speed dropped from 3750 to 3720 rpm during the last four minutes of the run. This response raised the question of whether the fuel metering system is capable of supplying enough fuel when the engine demand exceeds approximately 50 HP. This hypothesis is exarpined in a later section.

Next, Test Run 22 is considered, in which spray cooling was incorporated to eliminate the bearing temperature problem. The run was terminated prematurely due to an engine under-speed problem. Figure 7.3 shows fuel flow and engine speed as a function of elapsed time.

As can be seen, the engine speed remained nearly constant (3740 rpm) for the first twelve minutes of the run. During this time the fuel flow response mirrored the engine speed with corresponding peaks and valleys, again indicating proper control by the governor.

Following the twelve-minute mark, the engine speed drastically dropped off to 3400 rpm, while the fuel flow remained essentially constant at 100 Ibm/hr. For reference, Figure 7.4 shows the horsepower generated as a function of elapsed time, where the increased load is evident, correlating with the decrease in engine speed.

N AS N CR- 200 1-21 0675 3800 120 r-----------------------------------------------------------.

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o 3300~ ----~--~~------ __ ----~-- --~------ __ ----~--~~----__+·20 0:00:00 0:02 :53 0:05:46 0:08: 38 0:11: 31 0: 14 : 24 0: 17 : 17 0:20 :10 0:23:02 0:25:55 Elapsed Time Figure 7.3. Output shaft speed and fuel flow rate versus elapsed time for Test Run 22 (a duplicate of Figure 6.12 shown for convenience).

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NASAlCR- 200 1-21 0675 The trend is very similar to that of Test Run 16 ; that is, when the horsepower was low (20-25hp), the fuel metering response was normal. When the load was increased, the fuel flow rose to 100 lbs/hr (which seems to be its limit) and the engine speed rapidly decrease d.

The second hypothesis, generated during post-test analysis of Test Run 22 is that the fuel system has a maximum flow rate near 100 lb /hr as it is configured. Testing thi s possibility required a detailed look at the components involved in the fuel metering system. Figure 7.5 below is a schematic of a complete (stock) fuel system.

The main components of the fuel system are the governor assembly, fuel control housing, bellows cover assembly and the fuel solenoid valves. The components of interest in this analysis are the fuel solenoid valves and the supporting fuel circuitry.

The start, main, and maximum fuel solenoid valves are normally-closed valves which are activated by an electrical input of 14 to 30 volts DC. The start fuel solenoid valve is energized when the engine is at five percent of its rated speed. At 60 % rated speed the valve is de-energized and closes. The main fuel solenoid valve is energized at 15 % rated speed and remains open for the remainder of the engine operation. The maximum fuel solenoid valve is energized at 90 % rated speed (allowing two seconds of delay) and is intended to be in operation when the engine is approaching full load.

The fuel metering system on the Titan engine is a modified version of this stock system.

The photo below , Figure 7.6, shows the primary components and their respective locations. As can be seen, the maximum fuel solenoid is not incorporated in this metering system. Instead, the stock system was configured to work without the maximum fuel flow circuitry. In the photo there is an empty connection where this solenoid would have been attached.

NASNCR-2001 - 21 0675 93 I I I I

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NASN CR- 2001-210675 94

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Figure 7.6. Photograph of fuel control system for the Titan HPRTE.

In the original development of the engine, the governor was oversized because of the uncertainty in the required maximum fuel flow, so the maximum fuel solenoid was not included. The minimum flow orifice, shown in Figure 7. 3, was adjusted to a smaller passage size to avoid overheating the engine during start. The fuel passes through this orifice during nonnal engine operation as well, except for the flow through the ma ximum fue l solenoid, if installed. The apparent limitation in maximum fuel flow seems to be the restriction caused by the minimum flow orifice, unrelieved by the parallel path through the maximum fuel solenoid. Thus the power limitation appears to have nothing to do with the HPRTE cycle, but rather an easily-remedied flaw in the fuel control system.

NAS AJ CR- 200 1-21 0675 95

Chapter 8. Conclusions and Recommendations The philosophy behind this demonstration program has been to experimentally validate the thennodynamic analysis tools as applied to the HPRTE cycle, allowing confidence in the application of those tools to a broad range of application-oriented designs . The experimental program substantially validated the simulation methods, so that several applications could be the focus of system studies based on this concept. Four examples of such studies completed by the team are presented in Chapter 9, dealing with naval vessels, helicopters, and stationary power generation. Improvements in range are predicted for ships and helicopters of up to 24% and 47%, respectively. Combined-cycle powerplant efficiencies of greater than 60% are also predicted for an HPRTE combined with a steam bottoming cycle, with significantly decreased plant size and emission levels.

Thus further effort appears warranted in developing a prototype based on the HPRTE cycle.

8.1. Experimental conclusions A proof-of-concept testing program has been completed to demonstrate the attributes of a version of the HPRTE cycle. The program goals have been met, although with a diminished data set due to developmental difficulties umelated to the focus of the study.

The program objectives included demonstration of increased specific power relative to the baseline engine, increased part-load fuel efficiency, and decreased emissions. In addition, one purpose of this program was to gain operational experience so as to identify design issues at an early phase prior to construction of future prototypes or larger-scale demonstrations of this techno lo gy.

8.1.1. Emissions The projected decrease in emissions was realized in the test. The emission rate of carbon monoxide was almost two orders of magnitude lower than the reference engine (an umnodified Titan T62T32A), while the emission rate of NO x was approximately half that NAS N CR- 200 1-210675 ---_ .... - _.

~--- - - -- ---- ------ --- - Relative Emission Rate C Ref ere n ce APU • HPRTE I o~----~ ---- ---- Nox m Figure 8. 1. Improvement in the NO and CO emissions of the HPRTE relative to the x baseline (unmodified) engine.

of the reference. The extreme improvement in CO emission is attributed to the substantial increase in water vapor in the combustor due to the recirculation of exhaust products and, to a lesser degree, carryover from the spray cooler. The hypothesis is that the extra water vapor provides a source of radicals such as hydrogen atoms and hydroxyl radicals, which efficiently attack soot during the incipient polymerization process. By preventing much of the soot formation, less soot escapes the primary flame zone, so that the reaction has the chance to approach completion in the primary zone. Therefore, the production of CO and other incomplete combustion products was dramatically suppressed. Apparently, this effect outweighed the reduction in the gas-phase reaction rate, which accompanied the reduced oxygen concentrations, as least in the design regime tested.

NASAJCR- 200 1-21 067 5 98 The reduction in NO production was almost entirely attributable to the decrease in x exhaust flow rate due to the semi-closed cycle. The NO concentration, as opposed to the x total flow rate of NO produced, was comparable to the reference engine, a somewhat x unexpected result. The anticipated effect was that the increased diluent in the burner would reduce the peak flame temperature in the primary zone, exactly the same mechanism by which steam injection suppresses Ox. However, to realize the benefit of th e inherent steam injection in the HPRTE, the uniformity of the combustion process must still be comparable to that of the reference engine. The flowpath of the HPRTE unit was distorted significantly by the insertion and extraction of flow to accommodate the added recuperator. Si nce the production of NO is known to be highly sensitive to the x detailed temperature field and its uniformity, it is reasonable that the expected Ox concentration reduction was not observed. ote also that future HPRTE engines are expected to operate with significantly more recirculation, which would lead to further reductions in NO concentrations and emission rates. The observed reduction in NO x x concentration was obtained with a recirculation ratio less than 0.6. High recirculation (R>2) is projected to reduce NO emission via the dual effect of exhaust flow reduction x and concentration reduction, leading to an order of magnitude decrease.

8.1.2. Engine performance The configuration of the fuel system limited the power output of the HPRTE engine to approximately 35 % of its maximum, a limitation easily remedied in future designs. An extrapolation to the full power point was therefore required; a common practice in qualifying large-scale commercial engines [Nemec, 1999]. The maximum power for the HPRTE was projected to be 153 HP, which is in good agreement with the design models.

The baseline engine (Titan APU) power was approximately 90 HP when corrected to the ambient conditions of the HPRTE test, so a significant power increase was confirmed.

The increase in power was accompanied by only a small change in inlet airflow. Further, the density increase in the heat exchangers is projected to allow weight and volume penalties in prototypes to remain small, less than 40%, as opposed to conventional recuperated engine penalties of 200% or more.

NAS N CR - 200 1-21 0675 99 The agreement with models from several sources provides a high degree of confidence that the modeling is accurate. This was expected, since the processes in the HPRTE are well-understood : compression, expansion, heat exchange, and adiabatic mixing.

umerous engine analysis codes are in use which implement the same component models, so it is no surprise that the thermodynamic predictions for the complete cycle agree with the experimental results within the experimental uncertainty. evertheless, it is a gratifying result, since the system studies for HPRTE applications, also conducted at the University of Florida with support from this program, are now validated to some degree.

Another important feature of the HPRTE is its projected flat SFC curve versus percent power. As shown earlier in Figure 6.13 , the experimental result obtained was marginal at part load, rather than exhibiting the high part-load efficiency (flat SFC curve) claimed for the HPRTE cycle. The experimental result is highly misleading due to two main factors.

First, the Titan engine speed was considerably lower than design in the higher-power portion of the data. The low engine speed resulted in decreased HPC pressure ratio in that regime, with a corresponding rise in SFC. Therefore, if the engine speed had remained constant, the SFC would have dropped more steeply than the data show, resulting in a considerably flatter SFC curve as it approached the full power point. A simulation at constant engine speed, included in Figure 6.13 , bears out this assessment.

The second point is that the calculated flatness of the SFC curve for the HPRTE cycle stems in part from the capability of varying the LPC pressure ratio as a means of throttling the engine without affecting the temperatures in the high pressure region, instead of operating at fixed LPC pressure ratio. The principle is that changing the LPC pressure ratio changes the density of the working fluid in the core engine. The intercooler allows only slight variation in the HPC inlet temperature, so the core engine dimensionless operating point (pressure ratios, component efficiencies, etc.) can remain essentially fixed over the range of LPC pressure ratios achievable. In that case, the engine efficiency and SFC would also remain fixed. The experiment instead held LPC pressure NASAlCR-2001-210675 100 ratio nearly constant, so that power was decreased by decreasing turbine inlet temperature, as in conventional engines. It is not surprising that the experimental SFC curve flatness was similar to that of conventional recuperated engines. It should be noted that the original test plan provided for operation of the Titan HPRTE over a range of LPC pressure ratios 0 that the optimum SFC flatness could be determined, but that portion of the testing was curtailed due to time constraints. However, a simulation of the engine under optimal control did indeed show the expected flatness of the SFC curve (see Figure 6.13).

Future HPR TE designs would benefit from higher LPC pressure ratios than in the present study in order to provide a wider throttling range at constant SFC. For example, a LPC pressure ratio of 5 would in principle allow operation at 20% power by just decreasing the LPC pressure ratio to 1. That implies independent control of the low pressure spool via a wastegate or variable geometry which may impose a tighter constraint, but clearly, higher pressure LPC components will improve the range over which the SFC remains nearly constant.

8.2. Recommendations The testing program has provided sufficient confidence in the HPRTE technology that two classes of future work are recommended. An orderly progression toward a large- scale prototype is suggested for the experimental program, while a simultaneous focus on more detailed system analysis is conducted.

8.2.1. Experimental program recommendations To obtain further benefit from the current program, the existing Titan HPRTE could be used in order to improve the data set and provide additional design and operation experience. Modifications to the engine should include: • incorporate a dynamometer which has lower capacity, so that the minimum engine load is reduced.

• include a maximum power fuel solenoid in the fuel control system to allow high power testing (or install a more sophisticated electronic fuel control system) NAS N CR-2001-21067 5 101 • operate the engine through the original test matrix, including variations in the LPC pressure ratio as well as dynamometer load • replace the intercooler with a higher-effectiveness unit in order to allow operation with spray cooling optional • replace the engine power head to avoid bearing failure concerns The first three recommendations in the list above would be required, whereas the latter two are optional.

A larger-scale demonstration/early prototype program is also recommended. In such a program, more aggressive choices of the HPRTE design parameters would actually decrease the program risk. Specifically, a considerably larger value of the LPC pressure ratio should be chosen, and a larger design recirculation ratio should be incorporated. Both effects depend on the choice of the low pressure turbomachinery components. The higher pressure ratio would not only allow the engine power density to increase, but would also significantly improve the part-load SFC. Variable geometry or a wastegate arrangement would be required in order to provide independent control over the LPC pressure ratio.

The recommended increase in recirculation (R=1.5 or more) appears feasible with convention combustor technology, provided that the recuperator technology is state- of-the-art. This level of R would decrease the emission levels significantly while increasing specific power. The original R=2 combustor development plan proposed by Rolls-Royce Allison should also be followed in an effort to resolve the design issues in high recirculation burners. That plan called for cold flow and lightoff testing at Rolls -Royce Allison followed by parametric testing in the High Recirculation Combustion (HiRC) facility at the University of Florida. The original combustion program was not completed due to scheduling and funding constraints; however, the HiRC facility is in place and operational, as are the optical diagnostic tools developed for that program. Improvement of the design of high recirculation combustors for future HPR TE development does not appear to present a technological roadblock.

NASA/CR-200 1-21 0675 102 Given that a larger engine (say 1500 HP or greater) would have considerably better HP component efficiencies than the Titan, all of the attributes of the HPR TE cycle could be directly demonstrated . It is therefore recommended that the logical progression to a larger-scale demonstration or early prototype be undertaken.

8.2. 2. System study recommendations In parallel with the progression in experimentation, system design studies should be conducted to prepare for prototyping in several applications, and to assess the impact of HPRTE options. Those options include the possibility of extracting water from the HPRTE cycle [MacFarlane, 1997], generating water fog in situ at the HPC inlet, the turbofan combined cycle discussed in Appendix E Section 4, and distributed co generation configurations.

NASN CR-200 1-21 0675 103 , I

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Chapter 9. HPRTE Cycle Implications The Titan HPRTE test program established, with a limited data range, the validity of the predictions made for the modified engine. These included: • Increased power - maximum projected to be 153 HP versus 90 HP for the ba eline • Reduced emissions - Two orders of magnitude reduction in CO emission rate, factor of two in NO (this would improve significantly at higher recirculation ratios) x • Constant efficiency curve - non-optimal control strategy resulted in equivalent SFC vs. % power to that of recuperated engines; simple scaling shows improvement for HPRTE with improved control The lack of surprises in the test performance of the Titan HPRTE provides increased confidence in the ability of cycle simulation codes to predict performance and develop preliminary design parameters for a wide range of potential applications of this technology. Several such system studies have been performed by member organizations of the present development team as well as engine companies and government organizations outside of the team. Four representative system simulation results from members of the team are presented in this section in order to provide insight into the potential of the HPRTE cycle for several applications.

9.1. Rolls-Royce Allison Simulation Rolls Royce Allison performed an in-house simulation of the HPRTE prior to the initiation of this contract using their proprietary cycle simulation tools [Allison 1993].

Engine parameters appropriate for a modern engine in the 1500 HP class were chosen.

Both design-point and off-design studies were performed, and the results were compared to existing recuperated engines. The key results of this study are summarized in Figure 9.1, which shows the calculated SFC curve for an HPRTE design normalized to the SFC at the design point for fairness in comparing to the other cycles. The normalization allows a direct comparison, since the component technology level, along with the cycle, determines the design-point SFC.

NASAlCR-2001-210675 105 ____ J The recuperated AGT 1500 engine exhibits a much worse penalty in relative SFC at part- power conditions th an either of the other technologies; the intercooledlrecuperated SC- 21 engine shows an improvement relative to the AGT 1500, as expected for a more modem engine. The SFC curve for the HPRTE was calculated to be essentially flat down to about 30% of the maximum power, which has important implications in fuel economy for an y application in which part power is required for a significant portion of the mission.

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en RFTE-7 1.0 0.9 ~----1--_+_--_+_--_-__+--_+_--_+_-___I--_-~ o 10 20 30 40 50 70 100 60 80 90 Percent Power Figure 9.1. Comparison of SFC curves for the HPRTE, SC2l I C/R naval engine (under development) , and AGT 1500 M-l tank engine.

9 .2 . He lico pt er ap pl ications The University of Florida performed mission analyses using a version of the ROCETS system simulation code which was originally developed by NASA Marshall Space Flight Center with participation by Pratt and Whitney [Danias 1997]. This code was modified for improved input/output clarity and some components were re-written so as to be fully NASNC R- 2001 -210675 106 in the public domain . The code was thoroughly benchmarked against test cases and existing industrial cycle simulation codes, such as the one used in the Rolls-Ro yce Allison study.

The helicopter study was focused on generating a realistic comparison of range and SFC to the state-of-the-art for a mission profile provided by the U.S. Anny Vehicle Propulsi on Directorate. Component maps were generated for use within t he ROCETS code by utilizing a suite of design codes provided by NASA Glenn Research Center. The se included QUIK and CCODP for the compressor modules, AXOD for axial turbine s, and RTD and RTOD for centrifugal turbines. The T700-701C engine was chosen as the baseline, and, since its component maps are not public information, the design codes were exercised to develop maps which provided very similar performance to the T700 (pressure ratios, SFC, a nd specific power). The same level of component technology was then applied in optimizing two versions of the HPR TE cycle, again using the design code suite. The two HPRTE configurations analyzed are shown in Figures 9.2 and 9.3. The shafting arrangement is the only difference between the two configurations, with the cycle in Figure 9.2 being a 3-shaft machine, and the cycle in Figure 9.3 being a 2-shaft configuration. Both implement the HPRTE thermodynamic process, but the off -design behavior is distinct. The cycle state point designations shown on the figures were used internal to that study only, and differ from those shown in Figure 1.1.

NASAlCR-2001-210675 107 ------------ --- 5 .0 2 .0 High Pressure CompressHigh Pressure Turbin . Low

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Low Pressure Compressor 4.5 Figure 9.2. HPRTE efficiency (HPRTE E) mode flow path.

High Pressure ComprcssHigh Pressure Turbine 5 .0 2 .0 4.0 Low Pressure Compressor Medium 4.2 Pressure Turbine 4.5 3.4 2.5 3 .3 3.5 C b om ustor Recuperator 4.35 3 .0 Precooler 4.4 2.2 2.4 2.3 2.1 Figure 9.3 . HPRTE efficiency mode version 2 flow path (HPRTE 2).

NASNCR-2001-210675 108 The two HPRTE configurations and the baseline engine (s imilar to the T700) were analyzed in ROCETS and the range equation applied using the SFC result s. The outcome is summarized below in Table 9.1. As can be seen, the predicted range for both HPR TE configurations exceeds that of the baseline by at least 40%. This provides a very attractive incentive to consider the HPRTE for this application in spite of the somewhat increased complexity of the engine. It should be mentioned that the HPRTE also incurs a weight penalty due primarily to the heat exchangers. However, that penalty is estimated by several organizations to be 30 to 40% compared to simple-cycle engines, in contrast to the 200 to 400% penalty of conventional ICIR engine s. For missions in which the fuel weight is a significant fraction of the total propulsion system weight, the enhanced fuel economy provides a strong net benefit, as in the example case analyzed here.

Table 9.1. Nondimensional range for helicopter engine configuration.

M* Range increase baseline T700-701C 0.962 HPRTEE 1.413 47 % HPRTE2 1.374 43 % . .

Note: M* is the range for a speCIfIed mISSIOn profile, normalized by the range for the baseline engine operating at its design SFC.

The analysis involved first optimizing the design-point parameters for the HPRTE configurations. Multiple trade studies were performed; a typical result is shown in Figure 9.4, in which the sensitivity of SFC and specific power to the LPC pressure ratio is presented for the HPRTE E configuration. Note that the SFC levels are reasonably low for this engine size class and that the specific power is about three times the state-of-the- art due to high internal recirculation. An analogous chart is presented in Figure 9.5 for the HPRTE 2 configuration.

NASAlCR-2001-210675 109 RFTE Efficiency Mode

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NASNCR-2001-210675 110

RFTE Efficiency 2 Mode

..-...

....----------------,- 650 .2

0.5

..-...

a..

-

o :r: CD (J)

* 0.45

600*

a..

:r:

--------------------------

-

"- CD

550~

a..

-----

~0.35

~ (f) '0 CD 0..

+--+---1--+--+--I---+---+--+---+--+ 500 (f)

0.3

3.4 4 4.2

3.2 3.6 3.8

TIC Compressor Pressure Ratio

- LPCPR vs. Specif. P

- LPCPR vs. SFC

Figure 9.5. Design point study for the HPRTE 2 (fonnerly called RFTE 2) configuration by varying the Low Pressure Compression Ratio. TIC denotes Turbocharger (LP spool).

An off-design study of the optimized HPRTE E and HPRTE 2 configurations was conducted and compared to the perfonnance of the approximated T700 engine calculated under identical technology assumptions. The result is shown in Figure 9.6. As expected, the HPRTE configurations both provided an efficiency improvement at the design point, and maintained reasonably high efficiency at part loa d. Note that no variable geometry was assumed in the low pressure components which would have allowed the HPRTE engines to maintain an SFC curve that was even more nearly constant.

NASAlCR-2001-210675 111

:::=:: 0.9

0...

;:- 0.8

~

6-0.7

-

E

..0 0 .6

~ 0 .5

if)

0.4

0 .3

90 100

40 50 60 70 80

0/0 Output Power

- T700 - RFTE - E RFTE 2

Figure 9.6. Off-design comparison of SFC versus % power for the HPRTE (formerly called RFTE) configurations to the generic T700 (performance ofT700 and HPRTE estimated using the same synthesized component maps).

9.3. Naval Vessel Ap p lication A similar mission analysis was performed for a naval vessel application, in which a

mission profile was specified that is typical of u.s. warships. The baseline engine was

taken as the LM2500, an approximately 30,000 HP engine. The same philosophy was followed as that described in Section 9.2, that is, the NASA design codes were used to generate approximate component maps such that the global performance of the LM2500 was duplicated. The same component technology was then applied in optimizing two versions of the HPRTE, and comparisons in range or mileage were made [Landon 1997].

NASAlCR-2001 -210675 112 The design-point study once again required variation of several HPRTE parameters to find the optimum; a typical example i depicted in Figure 9.7. The figure of merit was taken to be SFC (or efficiency), ince all of the HPR TE configuration were expected to be sufficiently compact. A second example output i shown in Figure 9.8, where the effect of HPC pres ure ratio, at the design point, on the efficiency and specific power is plotted .

..--..

E 750

RFTEP ::::::::: ------------- . - ---- ---------------- _ .. _-- -----

~ 740

CJ) it

0... 730

I -- ~ --- ---------- ---- ----- --- -- ~~ :: '---'

Q) 720

RFTER

S

~ 710

-- ----- ~

~ ~ - ~ --- - --- - -=-=--:: - ~

~ ' - , u - - ::;::- ~ -; - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - ~ ~ -:. :- .:- .:- ~ - '"-" , - --::==:: - - -

;- 700

u

~690

(f)

8 9 10 11 12 13 14 15 16 17 18

High Pressure Compressor Ratio

- LPR = 5.5, RFTE R - -LPR = 7.5, RFTE R

- -- LPR = 9.5, RFTE R LPR = 3, RFTE P

LPR = 4, RFTE P - LPR = 5, RFTE P

Figure 9.7. Specific power versus compression ratio of the HPC for a naval vessel.

LPR designates the pressure ratio of the LPC. RFTE R refers to an HPRTE 2-shaft configuration; RFTE P refers to a 2-shaft configuration of a similar cycle, but where only the recirculated flow enters the recuperator hot side.

NASNCR-2001-210675 113 ..--..

'#-

--- 50

>- u HP C PR= 13 --- - --- - -------- -- -- ----- - ------ -- ------------- - -----Jl------ -

~ 49

HP C PR=1 8 \l

u

- - ------- - ------- --- ---- ---- ----- - -- ~ ---- --- ~ --- ------

~ 48

RFTER .~ - - ------- - -------- -- ---- -- ~ - ---------- - ----~--- --- -----

~ 47

HP C PR-1 8 HP C PR-1 3 HP C PR= 8 L-

- If RFTEP

ID --- -- ~- - ------- ~ ---- ---- ---- -- ---------- - - - - -- ---- - --------

..r=. 46

r-

\j ~ ~ HP C PR= 8 phi= .9

~ 45

>-

o 690 700 710 720 730 740 750

Specif ic Power (HP*sec/lbm)

- LPR = 5.5, RFTE R - - LPR = 7. 5, RFTE R

--- LPR = 9.5, RFTE R LPR = 3, RFTE P

LPR = 4, RF TE P - LPR = 5, RFTE P

Figure 9.8. Cycle thermal efficiency versus specific power. LPR de ignates the pressure ratio of the LPC. HPCPR is the pressure ratio of the HPc. RFTE R refers to an HPRTE 2-shaft configuration; RFTE P refers to a 2-shaft configuration of a similar cycle, but where only the recirculated flow enters the recuperator hot side.

Calculation of the average fuel con umption (mileage) was calculated on a normalized basis so that two HPRTE configuration could be compared to the baseline LM2500. The results are shown in Table. 9.2. Note that the fuel consumed by the HPRTE cycle engines can be approximately 24% less than the baseline engine. This would tran slate to 24% greater range, or, for a clean-sheet ship design, maIler fuel tanks. It is noted that naval vessels operate below 40% power over 50% of the mission profile, providing a premium on the off-design fuel efficiency of the propulsion ystem. Additional space savings would be expected becau e of the high specific power, which translates to air and exhaust flow reductions of a factor of three. The air intake and exhaust duct in multi-deck ships ASAlCR- 2001 -2 10675 114 consume much valuable real estate, so reductions of this order would have a significant beneficial effect on the ship architecture.

Table 9.2. Nondimensional mileage co m pariso n for a samp le marine vessel.

M* Range increase LM2500 1.34 baseline HPRTEP 1.66 24 % HPRTER 1.61 20% M* is defined as the range nonnalized to the range of the baseline system operating at it s design point. The baseline value of M* exceeds 1 because of the mission profile allows si gnificant cruise time at low speed.

9.4. Combined cycle power generation The final HPRTE application to be presented is that of a combined-cycle plant for baseload power generation. In this case, the off -design capability of the HPRTE is not utilized, since the plant would operate at full power continuously. The high part-load efficiency would allow effective load following, a very important attribute in distributed power applications. However, several other attributes of the cycle make this cycle an attractive alternative for baseload power as well. First, the intercooler in the HPRTE is already present, and now becomes the boiler for a bottoming cycle fluid with little or no additional cost. The boiler is compact by virtue of the high gas-side density discussed earlier, so its volume may be an order of magnitude smaller than a typical waste heat recovery boiler. The cost should be decreased in comparison with a waste heat recovery boiler. The emission levels should be low without recourse to steam injection or other ancillary remediation, again saving capital cost and complexity . Finally, the efficiency is calculated to be equivalent to the best combined cycle plants available, as shown below.

The ROCETS code was used for the system calculations, this time including a simple steam-based bottoming plant in addition to the HPRTE [Nemec 1995] . No component maps were necessar y, since the on-design performance was the focus of the study; instead, component efficiencies were specified as typical for large engines (e.g. 90% efficient turbines). Minimum pinch point temperature differences were specified to be NA SAJCR- 2001-210675 115 15 Fl. Numerous system parameters were iterated to find the optimum efficiency. Table 9.3 lists the optimal performance and design parameters; Figures 9.9 and 9.10 show example trade study results. Note in the figures that the sensitivity of the efficiency to recirculation ratio and to LPC pressure ratio is not great. This is a beneficial result, since those parameters may then be chosen so as to optimize system size (LPC pressure ratio) and combustion performance (recirculation ratio).

As a final note, it should be mentioned that the combined-cycle concept also has a potential aerospace application in a turbofan engine, although the concept has not been studied in detail. The HPRTE would be the core engine, rejecting heat to the bypass stream downstream of the fan . In that way, the bypass stream would undergo a Brayton cycle, rather than serving only as a propulsion medium, so the bypass stream would be a bottoming cycle, increasing the thermodynamic efficiency of the engine. It is expected that the overall efficiency would also increase, given proper design, so that range may be enhanced.

I A pinch point is the closest temperature difference in a heat exchanger. When one of the fluids is changing phase, the other fluid temperature approaches that of the first. Specification of the minimum temperature difference avoids large surface area requirements.

NASNCR-2001-210675 116 , Table 9.3. Cycle Results for Combined-Cycle Optimal Configuration Value Cycle Parameter Combined cycle thermal efficiency (Gross) 60.35 %" Gas turbine thermal efficiency 39.4% .- .... - .. - .. -.- . -~-- .. -.--- . ----- ...... - . .. . --- .... .. -.- .. . __ ...... -._ .. _ ..... I Bottoming cycle thermal efficiency 39.1 % -=: . __ .._ ._ ._ ... -_._ .. ._ .. _ .. __ .. _ ._ .. _---_ ._ .. .. .._ ..... -.

I

Topping cycle (HPRTE) - % of total power 65.3 %

I

Gas turbine specific power (Total HP / LPC 930.

_ inlet f!~w ), ._ ~..P! (lb~~ . ~c) .. _ .. _-_ •. ..

. _. . -- - .~ . ... -.-- .• - -. - LPC pressure ratio 4.5

... .. __ . _-_ .. __ .. _ .... . _----_ .. __ ... _._- _ .. . _ . • _ .. _ ..•. --- .. -

-"- ' HPC pressure ratio 8.3 Combustor exit equivalence ratio 0.9 Recirculation ratio (recirculated flow / LPC 1.92 inlet flow) Inter-cycle heat exchanger effectivene~ --'-' 0. 95 (intercoolers) Gas turbine recuperator effectiveness 0.80 • Under review. Moisture-handling loss in steam turbine may result in a decrease of up to one efficiency point, depending upon bottoming cycle design.

NASA/CR-2001-21067S 60 .5 ~

./'

V

/

~ 59.5 >.

u

/'

c: a> 'u

L

it: w c;; §

/

a> ~ .s= 58.5 I- a> U >.

U

/'

'C 58 a> c ~ :0 E u

/

57.5

/

~

/

56.5 0.60 0.80 1.00 1.20 1. 40 1.60 1.80 2.00 Rec ircu lation Ratio (rec irculati on flow I inlet flow) Figure 9.9. Combined Cycle Thermal Efficiency versus Recirculation Ratio.

NASNCR- 2001-210675 ll 8 60 .5 ~ ~ 60 .3 ........

<f!.

>- c CI>

/'

~

'0 lE

/

w 60 .1 iij E Q; ,t; ~ CI>

/

U ) >- u 59.9 CI> " c :0 E u

/

59.7 ~ 59.5 2.5 3 3.5 4 4. 5 5 5.5 6 Low Pressure Compressor Pressure Ratio Figure 9.10. Combined Cy cle Thermal Efficiency versus Low Pressure Compressor Pressure Ratio.

NA SNCR- 2001-210675 119

APPENDIX A. Water Injection System Operator Instructions The Titan HPRTE is equipped with a spray cooler system to inject a fine water mist at the inlet of the intercooler, thereby increasing its effectiveness. The following steps are required in order to operate the spray cooler: 1. Ensure that the electric power to the spray-cooler (SC) subsystem is OFF. The MAIN PWR switch should be OFF (down).

2. Examine all electric conduit and connectors for exposed wiring. If there is exposed wiring repair before proceeding with this procedure.

3. Make sure there is no water puddled underneath electric conduit lines.

4. Fill spray-cooler water supply tank to the 80 gallon level. The scale is visible on the north side of the tank. Do not allow the tank water level to rise above the return port from the pump. If water is allowed to rise above the pump return then the water level must be drained below the return before the system will operate properly.

5. Turn handle HI on the SC control panel to the fully closed position (full right).

6. Ensure that the Turbo Oil Pump is not plugged in, nor that any other electric equipment is connected to the outlets controlled by the MAIN PWR switch.

7. Turn on DC PWR.

8. Turn on MAIN PWR (switch up).

9. Turn on SC water pump by turning OIL PUMP 2 switch to ON.

10. PI pressure should read approximately 120 psig.

11. Open handle HI to the left approx. 50 degrees until the handle is in a horizontal position.

12. Depress the spray nozzle buttons to open the solenoid valves. PI pressure should drop below 90 psig. If not , adjust the H2 handle until P2 reads about 30 psig. Then PI should also be correct. The flowrate reading should be between 1.4 and 1.8 gpm.

13. Visually inspect the return water stream to the tank. The stream should indicate strong flow.

14. Inspect the pump area for water leaks.

NASNCR-2001-210675 15. Adjust handle H2 to decrease flow to l.2 gpm. H2 should be horizontal. PI should read about 50 psig and P2 should be about 15 psig.

CAUTION: Do not allow the water tank to run dry! Damage to the pump may result.

16 . Allow the SC to operate for approximately 25 minutes.

17. Tum the spray nozzles OFF. Pressure on PI and P2 should read about 90 psig.

18. Tum the pump off by switching OIL PUMP 2 to OFF. The water lines between the pump and the nozzle solenoid valves should still be pressurized.

19. Check for water leaks. If there are no leaks the system checkout is complete. Proceed to prepare for run operation.

20. Relieve the pressure on the water lines by opening a nozzle solenoid valve. The pressure on PI and P2 should drop to zero. Close the nozzle solenoid valve.

21. Prepare for run operation by leaving HI and H2 in their existing positions. Ensure that the nozzle solenoids are closed. The buttons should be fully extended.

22. Refill the SC supply water tank to the 80 gallon level. Do not overfill above the pump return.

23. Before engine startup tum the pump on (OIL PUMP 2 to ON).

24. During engine operation, begin spray cooling by pressing the SC nozzle buttons. PI should read about 60 psig and P2 about 90 psig. The flowrate should read about 0.8 gpm.

25. To adjust the flowrate in step-changes, the nozzles can be turned on and off as needed. For finer control manipulate H2 to give the flowrate desired.

26. To shut down, tum OIL PUMP 2 to OFF, tum off the nozzle buttons, and tum handle HI fully shut (all the way to the right).

27. For emergency stop of the water flow to the engine, tum handle HI fully to the right.

Tum off pump as needed.

28 . Secure MAIN PWR and DC PWR to OFF as required.

Note: All pressures and flowrates mentioned are approximate. Even with limited experience with the spray cooler subsystem, the system has shown some tendency for hysteresis. The values should be checked and updated at regular intervals.

NASAlCR-2001-210675 122

Appendix B. Data Acquisition Software Configuration

The Lab Tech Control set -u p file used in Test Runs 17 to 22 is shown Table B .l below.

Note that the data input names in the fi le fo llow the naming conven ti on presented in

Section 3.4. 1.

Table B. 1. Set-up fi le fo r Labtech Control for Test Runs 17 to 22.

Start FileName Scale Bl ock Block Block (first) Factor Offset Iter Stg duration Rate State Deveh Rack Slot Function No. Name D: \TITAN\T3 @. TIT 7200 0.5 ON Time I TIME D: \TITAN\T3 @. TIT 7200 0.5 ON 2 ET Time

o Thermocouple 0 7200 ION

3 J2 49 0 0 7200 ION

21 0 o Thermocouple

4 T6 7200 ION

o Thermocouple 0

5 T4 I 19 0 ION

o Thermocouple 0 7200

48 0 6J1 ION

o Them1ocouple 0 7200

80 0 7 KI ION 0 7200

8K2 81 0 o Thermocouple

ION

o Thermocouple 0 7200

9 K3 82 0 7200 ION

o Thermocouple 0

JO K4 83 0 7200 ION

o Them1ocouple 0

lIKS 84 0 7200 ION

o Thermocouple 0

12 K6 85 0 7200 I ON

o Thermocouple 0

13 K7 86 0 7200 ION

o Thermocouple 0

14 K8 87 0 7200 ION

0 o Thermocouple 0

IS K9 88 7200 ION

o Thermocouple 0

16 KLO 89 0 7200 ION

o Thermocouple 0

17 KlI 90 0 7200 ION

o Them10couple 0

18 KI2 91 0 7200 ION

o Thermocouple 0

19 KI3 92 0 7200 ION

o Thermocouple 0

20 KI4 93 0 7200 ION

94 0 o Thermocouple 0

21 KI5 7200 I ON

o Thermocouple 0

22 KI8 97 0 7200 ION

o Analog Input 30.98 -29.781

23 G7 2 148 0 7200 ION

o Analog Input 31.58 -30.381

2 149 0 24 G8 7200 ION

o Analog Input 31.27 -30.064

25 G-I-l 2 150 0 0 7200 I ON

0 o Analog Input 1

26 DP2 2 IS2 -0.1825 7200 ION

0 o Analog Input 0.186

27 DP4 2 154 -20.489 7200 I ON

0 o Analog Input 20.22

28 G 1-2 2 ISS -21.391 7200 ION

o Analog Input 20.27

29 G 1-3 2 156 0 7200 ION

o Analog Input 20.24 -20.895

30 G 1-4 2 157 0 7200 ION 20.25 -21.057

2 158 0 o Analog Input

31 GI-5 7200 ION 20.25 -21.057

2 160 0 o Analog Input

32 GI-6 7200 ION

o Analog Input 20. 18 -20.228

2 161 0 33 GI-7 7200 ION 20.34 -20.985

2 162 0 o Analog Input

34 GI-8 7200 ION 20.37 -21.115

2 163 0 o Analog Input

35 GI -9 7200 ION

o Analog Input 6.0 19 -6.0727

2 165 0 36 G2 -1 7200 I ON

o Analog Input 0

2 174 0 37 Tq-inpt 7200 I ON

o Analog Input 0

2 175 0 38 SPEED2 7200 ION

3 0 o Frequency

39 FUEL

NASAlCR-2001-210675

o Frequency 7200 ION

40 SPEEDI 3 2 0 7200 0.5 ON D: \TITAN\T3 @. TIT 41 T-DynTnk Block Av(3) 0 D:\TIT AN\T3@.TIT 42 T-CW2i Block Av(4) 0 7200 0.5 ON Block Av(5) 0 7200 0.5 ON D:\TITAN\T3 @.TIT 43 T-CW2x 44 T-xgas Block Av(6) 0 7200 0.5 ON D:\TITAN\T3 @.TIT D: \TITAN\T3 @. TIT 45 T-HPtxl Block Av(7) 0 7200 0.5 ON D: \TITAN\T3 @. TIT 46 T-HPtx2 Block Av(8) 0 7200 0.5 ON D:\TITAN\T3 @. TIT 47 T-HPtx3 Block Av(9) 0 7200 0.5 ON D: \TITAN\T3 @. TIT 48 T-turbin Block Av(lO) 0 7200 0.5 ON 49 T-amb Block Av(l!) 0 7200 0.5 ON D: \TITAN\T3 @. TIT Block Av(12) 0 7200 0 .5 ON D: \TITAN\T3 @. TIT 50 T-HPci Block Av(l3) 0 7200 0.5 ON D: \TITAN\T3 @.TIT 51 T-LPtil 52 T-Bear Block Av(l4) 0 7200 0.5 ON D: \TITAN\T3 @. TIT Block Av(l5) 0 7200 0.5 ON D:\TITAN\T3 @. TIT 53 T-LPti3 Block Av(J6) 0 7200 0.5 ON D: \TITAN\T3 @. TIT 54 T-LPcx Block Av(l7) 0 7200 0.5 ON D: \TITAN\T3 @.TIT 55 T-HPrx 56 T-HPri Block Av(l8) 0 7200 0.5 ON D:\TITAN\T3 @. TIT Block Av(19) 0 7200 0.5 ON D: \TITAN\T3 @. TIT 57 T-Airin Block Av(20) 0 7200 0.5 ON D: \TITAN\T3 @.TIT 58 T-Filtr Block Av(21) 0 7200 0.5 ON D:\TITAN\T3 @.TIT 59 T-icool Block Av(22) 0 7200 0.5 ON D: \TITAN\T3 @. TIT 60 T-LPtx Block Av(23) 0 7200 0.5 ON D: \TITAN\T3 @. TIT 61 P-HPcx 62 P-HPri Block Av(24) 0 7200 0.5 ON D: \TITAN\T3 @. TIT Block Av(25) I 0 7200 0.5 ON D:\TITAN\T3 @.TIT 63 P-HPrx D: \TITAN\T3@.TIT 64 DP-HPci Block Av (26) 6 -5.89 7200 0.5 ON 65 DP2vdc Block Av(26) 0 7200 0.5 ON D: \TITAN\T3 @. TIT 66 DP-LPci Block Av (27) 0 7200 0.5 ON D:\TITAN\T3 @. TIT Block Av(28) D:\TITAN\T3 @. TIT 67 P-HPci 0 7200 0.5 ON Block Av(29) 0 7200 0.5 ON D:\TITAN\T3@.TIT 68 P-HPtx Block Av(30) 0 7200 0.5 ON D: \TITAN\T3 @. TIT 69 P-LPti Block Av(31) 0 7200 0.5 ON D: \TITAN\T3 @.TIT 70 P-LPcx Block Av(32) 0 7200 0.5 ON 71 open2 Block Av(33) 0 7200 0.5 ON 72 open3 Block Av(34) 7200 0.5 ON D:\TITAN\T3 @. TIT 73 P-Wgate 0 74 P-Filtx Block Av(35) 0 7200 0.5 ON D:\TITAN\T3 @. TIT Block A v(36) 0.5 ON D:\TITAN\T3@.TIT 75 P-amb 0 7200 Block Av(37) 0 7200 0.5 ON D:\TITAN\T3 @. TIT 76 Torque 450 Block Av(38) 3300 7200 0.5 ON 77 S2avg 100 ulimit(77) I 0 7200 0.5 ON 78 ULimit D:\TITAN\T3 @. TIT 79 PPH Block Av(39) 0.702 0 7200 0.5 ON 0.5 ON D:\TITAN\T3 @. TIT

80 HPI (76) * (8 2) ##### 0 7200

0.5 ON D:\TITAN\T3@.TIT

81 HP2 (76) * (83) ##### 0 7200

D:\TITAN\T3@ .TIT 82 RPMI Block Av(40) 60 0 7200 0.5 ON D:\TITAN\T3@.TIT

83 RPM2 (77) * (78) 0 7200 0.5 ON

84 LPtiALM 850 0 7200 0.5 ON 85 BearALM 270 0 7200 0.5 ON 0.5 ON 86 HPciALM l75 0 7200 7200 0.5 ON 87 HPriALM 515 0 7200 0.5 ON 88 EGTALM 1200 0

NASNCR-200 1-210675 124

ABS(66) IE+05 7200 0.5 ON 89 calc l 0 I 7200 0.5 ON 90 Vel-LPci SQRT (89) 0 D:\TITAN\T3 @. TIT 7200 0.5 ON 91 MFLOair -90 0.007 0 7200 0.5 ON 92 A (5 1) + (53) 0 .5 0 (59) - (54) 0.889 7200 0 .5 ON 93 B 0 (92) - (59) 0 7200 0.5 ON 94 C (92) - (57) 0.002 0 7200 0.5 ON 95 D -94 0 .27 0 7200 0.5 ON 96 EI

(9 1) * (96) 0 7200 0.5 ON

97 E 7200 0 .5 ON 98 F (54) - (57) 0.002 0 460 .67 7200 0.5 ON 99 GI -92 460.67 7200 0.5 ON 100 G2 -57 7200 0.5 ON 101 Gla (99)**r 0 ( lOO )**r 7200 0.5 ON 102 G2a 0 (1 01) - (102) 0 7200 0.5 ON ##### 10 3 G -54 460.67 7200 0.5 ON 104 HI (l 04 )**r 0 7200 0.5 ON 105 Hla (105) - ( 102) 0 7200 0.5 ON 10 6 H ##### (93) / ( 94 ) 0 7200 0.5 ON 107 !

7200 0.5 ON (95) + ( 98 ) 0 108 Ja 0.5 ON (1 03) + (106) 0 7200 109 Jb (l08) + (109) 0 7200 0.5 ON 110le ( 110) / (97) 0 7200 0.5 ON III J 0.5 ON D:\TITAN\T3 @.TIT (107) + (Ill) 0 7200 112 MFLOrat 0.5 ON D:\TITAN\T3 @. TIT (91) * (112) 0 7200 113 MFLOrec 0.5 ON D:\TITAN\T3 @. TIT ( 91 ) + (113) 0 7200 114 MFLOtot 0.5 ON (64) / (67) 0 7200 liS MFHPbc 0.5 ON I -0.004 0 7200 116 MFHPa 0.5 ON ( 115)**r -4.387 0 7200 11 7 MFHPb 0 .5 ON (l15)**r 7.439 0 7200 118 MFHPc 7200 0.5 ON (116) + ( 117) 0 119 MFHPab 7200 O.S ON (118) + (119) 0 120 MFHPabc 7200 O.S ON -50 460 121 MFHPd 0.5 ON SQRT(l21 ) 0 7200 122 MFHPe 0.5 ON (67) / (122) 0 7200 123 MFHPf 0.5 ON (120) * (123) 0 7200 124 MF-HPci 7200 0.5 ON (75) - (66) 144 0 125 MFLPa 7200 0.5 ON -57 I 460 126 MFLPb 7200 0.5 ON -126 53.3 0 127 MFLPc 7200 0 .5 ON ( 12S ) / ( 127) 0 128 MFLPdens 0.5 ON -66 9274 0 7200 129 MFLPd 0.5 ON (129) / (128) 0 7200 130 MFLPvel 0.5 ON 0.082 0 7200

(128) * (130)

131 MF-LPci 0.5 ON ( 124) - (131) 0 7200 132 Ra 7200 0.5 ON (132) / (131) 0 133R 0.5 ON 0 7200 (5 6) - ( 50 ) 134 DT -HPi 0.5 ON (8 0) + ( 81 ) 0.5 0 7200 135 HPavg 7200 0.5 ON

(124) * (134) 0. 24 0

136 Tlpwr-a (135) + ( 136) 1 0 7200 0.5 ON 137 TJpwr

NASNCR-2001-210675 125

I

I

I

I

I

I

I

L

Appendix C. Gas Analysis Setup Procedure The following steps are necessary to operate the gas analysis equipment in the Energy & Gasdynamic Systems Laboratory at the University of Florida: 1) Connect the heated sample hose to the sample probe. Connect the other end of the heated sample hose to the IMR500P Flu Gas Drier.

2) Connect the heating wire plug, which is attached to the heated sample hose, to the jack on the IMR500P.

appropriate 3) Connect the 5-pin thermocouple plug, which is attached to the heated sample hose, to the IMR500P jack labeled "THERMOFUHLER HEIZUNGEN."

4) Connect one end of the long side of the silicon sample line to the IMR500P jack labeled "MESSGASAUSGANG."

5) Connect the other long side of the silicon sample line to the 0 to 5 psi pressure gauge located in the vicinity of 3-way ball valve.

6) Connect one end of the short side of the silicon sample line to the COSA6000 Gas Analyzer jack label "smoke gas."

7) Connect the other short end of the silicon sample line to the RI-411A CO Analyzer jack, located on the side of the unit, labeled "INLET."

8) Verify that the 5-pin thermocouple plug which is attached to the sampling probe is plugged to the COSA6000 jack labeled "Temperature Smoke Gas."

9) Verify that the 5-pin thermocouple plug is plugged into the COSA6000 jack labeled "Temperature Room."

Smoke Opacity Meter Setup Procedure 1) Release the clamps holding the light source and sensor to the exhaust duct. Inspect the glass and to verify that it is clean. Reattach the light source and sensor to the exhaust duct.

2) Connect the gray cables from the smoke opacity meter to the appropriate locations on the light source and sensor.

3) Verify that purge is attached.

4) Verify that inlet cooling water is attached.

NASAlCR-2001-210675 127 MRU-Graphics Startup Procedure 1) Verify that the RS 232 cable is connected to the COSA 6000 connection labeled " RS 232" 2) Power on the IBM Value Point 3) When you encounter a " 164 Memory Size Error", pre ss "Esc."

4) At the prompt, enter "cd MRUGRAPH" 5) At the next prompt, enter "MRUGRAPH" 6) Press enter at the firs t MRUGRAPH screen 7) Select "DATA" 8) Select "DIRECT INDIC."

9) Select "File for Storage" 10) Enter an appropriate me name.

11 ) Select OK 12) Select "INDICATE" 13 ) Select OK 14 ) MRUGRAPH should now be ready to receive data.

Gas Analysis Pre-Test Startup Procedure Note: The COSA6000 AND RI-411A SHOULD NOT BE POWERED ON WHEN THE SAMPLING PROBE IS EXPOSED TO EXHAUST GASES. In a similar fashion, do not power down either unit while the sampling probe is exposed to exhaust gases.

1) Switch the IMR500P power switch "AUS" to "EIN."

2) Switch the COSA6000 power switch from "0" to " I."

3) Switch the RI-411A knob from " OFF ' to "CONT."

4) Tum on the heating wire control box, if it is being use d.

I I I NASNCR-2001-210675 128 Note: The COSA6000, RI-411A, and IMR500P will require several minutes to warm up.

When the COSA600 has wanned up, it will prompt the user for information about the fuel type being used. Use the up/down arrow buttons and "enter" button to answer the questions. It will then bring up the normal gas analysis screen. When the RI-411A is ready for use, the display will read "CONT x.x%", where x.x is the CO concentration entering the analyzer at the time. When the IMR500P is ready for use, the red light will stay lit, while the green light will blink.

Smoke Opacity Meter Pre -T est Startup 1) Supply 5 to 10 psi of purge air.

2) Verify that the cooling water exit hose is running out the bay door.

3) Turn on cooling water.

4) Power on opacity meter.

5) Adjust the zero re ading with the "0 % ADJ" knob. If the reading cannot be brought to zero, refer to operators manual.

MRUGRAPH Shutdown Procedure 1) Select "Esc."

2) Select "FILES" 3) Select the appropriate file and enter OK 4) Select "EXPORT" 5) Select "ASCII-TAB -CALC" 6) Name it appropriately and select OK 7) Select "END" 8) Select "END" Select "Yes" 9) 10) At the prompt, verify that the file has been written. It will be stored in the MRUGRAPH directory.

11 ) Download the file to the disk used for the Main Data Acquisition data.

NA SNCR- 2001-210675 129 Gas Analysis Shutdown Procedure 1) When finish testing, disconnect both of the short ends of the silicon sampling lines from the hose barb tee and let the COSA6000 and RI-411A run for a few minutes to purge engine exhaust gas from the units.

Note: DO NOT STOP EITHER UNIT WlllLE EXHAUST GASSES ARE IN THEM.

This can affect the calibration and shorten the life of the sensors.

2) Power off the IMR500P, at any time.

3) Power off the heating wire control box.

4) After the COSA600 and RI-411A have been purged with ambient air, they may be powered off.

NASAlCR-200 1-210675 130

Appendix D. HPRTE Engine Operation D.l. Kahn Hydraulic Dynamometer Set-up A Kahn water brake dynamometer was loaned to the program by the Amly Research Laboratory, Aberdeen Proving Grounds. This section describes operation of the unit and its supporting equipment as installed at the University of Florida.

Air at 80 psig is supplied to the dynamometer control valve. The control box is turned on and the control valve potentiometer is set to zero (no load on the dynamometer) . All water connections are checked for proper installation and for any leaks. A 30 gallon surge tank is filled to about 25 gallons of water. The level should be about 2 inches below the return connection.

Operational testing starts by fully closing the return line globe valve and then opening it two full turns. This provides a small amount of back pressure to the dynamometer. The supply gate valve, located on the high pressure discharge side of the centrifugal pump, is checked closed. The pump is powered and the pressure gage is checked to read 40 psig.

Once up to pressure, the gate valve is opened s lo wly to allow flow to the dynamometer.

There should be a small amount of flow back to the surge tank (note that the control valve potentiometer is in the zero position). The dynamometer is checked for leaks around the casing split. Leaks indicated over-pressurization and can be corrected by adjusting the return line globe valve.

With the discharge gate valve open, operation of the dynamometer is performed by cycling the control valve potentiometer. The control valve should operate smoothly and flow back to the surge tank should increase. After verification of operation, the potentiometer is set back to zero, the gate valve is shut and the pump is turned off. The dynamometer is to be drained by opening the 1/8 inch needle valve. This is to be done to ensure that there is a little load as possible on the engine during start-up. Once drained, the needle valve is closed.

NASAlCR-2001-210675 D.3.2. Analog Pressure Panel The person assigned to this station has the following duties: 1. Monitoring boost and maximum main loop pressure during transients 2. Being familiar with expected pressures to report any anomalies 3. Reporting fuel flow during transients 4. Recording fuel flow, manometer readings and pre ss ures 5. Refilling fuel reservoir tank 6. First to assist engine overwatch D.3 . 3. Analog Temperature Panel The person assigned to this station has the following duties: 1. Monitoring Titan engine roller bearing temperature during transients 2. Monitoring recuperator low pressure exit temperature to determine steady- state operation 3. Being familiar with expected temperatures to report anomalies 4. Monitoring dynamometer return water temperature 5. Recording temperatures 6. Second to assist engine overwatch D.3.4. Digital Data Acquisition The person assigned to this station has the following duties: 1. Obtaining atmospheric pressure before and after the test run 2. Monitoring the TEMPERATURES panel 3. Providing time checks to the audio recorder 4. Making occasional observations on the exhaust gas stack gas 5. Taking a continuous data set D.3.5. Gas Analysis and Spray Cooler The person assigned to this station has the following duties: 1. Providing continuous exhaust carbon monoxide levels during transients 2. Recording gas analysis data set 3. Recording opacity data set 4. Monitoring spray cooler pressure and flow rates 5. Recording spray cooler system pressure, nozzle pressure and flow rates 6. Starting and stopping the dynamometer pump D.3.6. Engine Control The person assigned to this station has the following duties: 1. Starting and shutting down the Titan engine 2. Monitoring all critical engine parameters during transients 3. Being familiar with all engine parameters 4. Controlling boost 5. Controlling dynamometer load 6. Opening and closing recirculation valve 7. Opening wastegate on shutdown 8. Recording inlet air manometer readings NASA/CR-2001-210675 133 D.3.7. Operations Supervisor The person assigned to this station has the following duties: 1. Supervising test runs 2. Updating run log 3. Directing the crew when to take data 4. Providing support to engine control 5. Monitoring critical temperatures and pressures DA. Titan High Pressure Regenerative Turbine Engine Operation D.4.l. Pre-start Procedure Prior to starting the engine the operations supervisor verifies the following: 1. Monitoring equipment is on and recording 2. Communjcations are functional (test by calling each crew member) 3. No load on the dynamometer (zero reading on the potentiometer) 4. The Fisher valve is fully shut (verify visually) 5. The air conditioning switch is in the "off ' position.

6. Spray cooler pump is operational and runillng with the solenoid valves in the closed position D.4.2. Engine Starting Procedure 1 Power to the control console is provided by a DC power source located behlnd the instrumentation panel.

2 Verification of power supply to the control panel is done by cycling the DIC breakers in the on/off position. Prior to start Die breakers should be in the " off ' position.

3 The turbocharger lube oil pump is turned on and pressure is verified at 80 psig locally and at the control panel.

4 The turbocharger wastegate valve is cycled shut then open. This is done by supplying pressure to the valve to close it and removing pressure to open it. The valve is to be in the "open" position prior to start (0 psi on the wastegate pressure gage).

5 The bypass valve is cycled full open then shut. A reading of 0 psi on the gage indicated full open and visual verification on the valve is done to confirm. The valve is then fully closed and remains closed prior to the run.

6 The recirculation valve is checked fully shut by turning the control wheel clockwise.

NASA/CR-2001-210675 134 7 Prior to starting the engine ("hot light" of the HPRTE) , it is first purged. The purgelstart switch is set to the "purge" position and the Die breakers are switched to the " on " position. The engine is then started by turning the start/stop switch to the "start" position. During the "purge," oil pressure is verified at 20 psig and fuel pressure is verified at 450 psig. Engine speed is verified at 40 % speed. This procedure is to be repeated until the above conditions are achieved.

8 After verification of engine speed and oil and fuel pressures the start/stop switch is put in the "stop" position. Visual verification of a complete engine stop is performed prior to a "hot light" of the HPRTE.

9 Once the engine comes to a complete stop the Die breakers are cycled. With the Die breakers in the "on" position the purgelrun switch is placed in the "run" position. The start/stop switch is then switched to the "start" position. Verification of a "hot light" is indicated with an increase in exhaust gas temperatures and an engine speed reading of about 102%.

D.4.3. Standard Operational Procedure After verification of a "hot light" and full speed is achieved, the recirculation valve is immediately opened two turns. The dynamometer inlet gate valve is opened two turns and the rotameter control valve is opened until a flow rate of 0.8 gpm is achieved.

The turbocharger waste gate valve is closed by slowly supplying 20 psig of shop air. The valve is fully shut when the gage reads 8 psi. Turbocharger adjustment is governed by the operator who observes the low pressure compressor exit pressure (P-LPcx) and the high pressure compressor exit pressure (P-HPcx). When the pressure in the low pressure 15 psig or the pressure in the high pressure compressor exit compressor exit reaches reaches 120 psi, shop air is slowly supplied to the Fisher turbo speed trim valve. This reduces boost pressure. The Fisher valve begins to open at 5 psig and is fully open at 15 psig.

The recirculation valve is opened slowly after the low pressure turbine inlet temperature reaches 400 F or greater. Spray cooling is initiated after the recirculation valve is fully NASNCR-2001-210675 135 opened or the bearing temperature reaches 260 F. The flow rate for the spray cooler should be 1.8 gpm at full power and initially only one sprayers is used.

After providing boost to the engine and spray cooling has been initiated, the engine is loaded to the 1150 F EGT limit and at no less than 99 % speed (engine will automatically shut-down if EGT exceeds 1150 F or engine speeds drops below 92 % speed or exceeds 104 % speed). The dynamometer control potentiometer is used to specify the load on the engine. The load is displayed on an Omega DP- 41 reader and is set to directly read torque in foot-pounds (ft-Ib£). While loading the engine it is necessary to increase the water flow to the dynamometer to ensure proper cooling of the unit (4 gallhr*hp as specified) and adjust the turbocharger trim valve to maintain the desired boost pressure.

D.4.4. Shut-Down Procedure Prior to shut-down of the HPRTE, the dynamometer control potentiometer is returned to the zero position and the dynamometer gate valve is fully shut; this effectively unloads the engine. The start/stop switch is then put in the "stop" position. The spray cooler nozzle(s) are then shut off and the dynamometer pump is shut off. After verification of complete engine stop, the recirculation valve is closed fully then opened to one turn. The run/purge switch is then put in the "purge" position. D/e breakers are cycled and the start/stop switch is put in the "start" position. The engine is run for 30 seconds. This procedure is repeated until the exhaust gas temperature is below 500 o f.

NASAlCR-2001-210675 136 REFERENCES Allison Gas Turbine Company, "Evaluation of the RFTE," Internal Report, 1993 .

Anxionnaz, R. (Societe Rautear, SA): Installation a turbines a gaz a circuit semi-ouvert, Franz, Pat. 999-133 (14.11.1945).

Anxionnaz, R. (Societe Rautear, SA): Improvements in or relating to gas turbine plant with semi-open circui t. Brit. Pat. 651-166 (20.7.1948).

Baumeister, T. , A vallone, E.A. , and Baumeister III, T., Marks Standard Handbook for th Mechanical Engineers, 5 edition.

Crittenden, J.F., Jr., Dilute, Kinetically Controlled Combustion Efficiency Prediction /or Recirculating Semi-closed Gas Turbine Cycles: a Non-dimensional Approach Using the First Damkohler Number as a Parameter, Masters Thesis, University of Florida, Gainesville, Dept. of Mechanical Engineering, 1999.

Crittenden, J.F., Jr. , Lear, W.E., and Azzazy, M., "Exploratory Design of a Depleted Oxygen Gas Turbine Combustor," Final Report NASA Contract NAS3-27759, Oct. 1999.

Danias, G. A. , Design and Off-D es ign Point Study o/Two Regenerative Feedback Turbine Engines for Helicopter Applications, Masters Thesis, University of Florida, Gainesville, Dept. of Mechanical Engineering, 1997.

Davis, R.C., "Final Test Report of Wolverine Gas Turbine Plant", United States Naval Engineering Experimental Station Research and Development Report 030076 (NS-622- 210), August 1956.

DeWitt, S.H. and Boyum, W.B., "Internally Fired Semi-Closed Cycle Gas Turbine Plant for Naval Propulsion", AS ME Gas Turbine Power Conference, Paper No. 56-GTP- 16 (1956).

Gasparovic, N. , "Semi-Closed Cycle Gas Turbine Plants", Combustion, Nov. 1967, pp.14-25.

NASAlCR-2001-210675 137 Gasparovic, ., "The Advantage of Semi-Closed Cycle Gas Turbines for aval Ship Propulsion," aval Engineers Journal, April 1968, pp. 275-29 1 (p. 333 ).

Landon, lC ., D es ign and Off-D es ign Point Study of Two Regenerative Feedback Turbine Engines f or Marine Applications, Masters Thesis, University of Florida, Gainesville, Dept. of Mechanical Engineering, 1996.

MacFarlane, R.S. , A Study of th e Impact of Water Extraction on th e Regenerative Feedback Turbine Engine Cy cl e, Masters Thesis, University of Florida, Gainesville, Dept. of Mechanical Engineering, 1997.

Marek, C.J. and Tacina, R.R. , "Effect of Exhaust Gas Recirculation on Emissions from a Flame-Tube Combustor Using Liquid Jet-A Fuel," ASA TMX-3464, Dec. 1976.

Meier, J.G. and Vollerin, B.L., "The Design of an Integrated Burner-Boiler System Using Flue-Gas Recirculation", Research Conducted at Battelle Research Center, Geneva, Switzerland, Proc. of 1 t il International Meeting of th e Combustion Institute, 1977.

Nemec, T. S. , Th ermodynamic Design Point Study of a Semi-Closed Recuperated Intercooled Gas Turbine Combined with a Rankine Bottoming Cycl e, Masters Thesis, University of Florida, Gainesville, Dept. of Mechanical Engineering, 1995.

ernec, T.S ., Private Communication, 1999.

Rodgers, c. , "RFTE Testing Summary," Alturdyne internal report, Feb. 1997.

NASNCR-2001-210675 138 I I L _______________________________________ -.J Form Approved REPORT DOCUMENTATION PAGE OMB No. 0704-0188 Public reporting burden for this col/ection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of informat ion . Send co mm en ts regardin g this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 12 15 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Redu ction Project (0704-0 18 8) , Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2 REPORT DATE 3 REPORT TYPE AND DATES COVERED 1 . 1 .

March 2001 Intelim Contractor Report--4/22/95 to 9/30/99 4. TITLE AND SUBTITLE 5. FUNDING NUMBERS High Pr ess ur e Regenerative Turbine Engine: 21 st Century Propulsion VVU-708-28-13-00 6. AUTHOR(S) NAS3-27396 w.E. Lear and A.L. Laganelli PERFORMING ORGANIZATION 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8.

REPORT NUMBER Science Applications International Corporation E-12602 1100 First Avenue, Suite 300 King of Prussia, Pennsylvania 19406 SPONSORINGIMONITORING 9. SPONSORINGIMONITORING AG ENCY NAME(S) AND ADDRES S (ES) 10 .

AGENCY REPORT NUMBER National Aeronautics and Space Administration Washington, DC 20546-0001 NASA CR-2001-210675 GDL99-3 11. SUPPLEMENTARY NOTES W.E. Lear, University of Florida, Department of Mechanical Engineering, PO . Box 116300, Gainesville, Florida 32611-6300 and A.L. Laganelli, Science Applica ti ons International Corporation, 1100 First Avenue, Suite 300, King of Prussia, Pennsylvania 19406. Project Manager, Paul Senick, Aeronautics Directorate, NASA Glenn Research Center, organization code 2200,216-977-7 024.

12b . DISTRIBUTION CODE 12a . DISTRIBUTION/AVAILABILITY STATEMENT Unclassified - Unlimited Distribution: Standard Subject Categories: 07 and 34 Available electronically at htro:llgltrs.grc.nasa.gov/GLTRS This publication is available from the NASA Center for AeroSpace Wonnation , 30 1-621-0390.

13. ABSTRAC T (Maximum 200 words) A novel semi-closed cycle gas turbine engine was demonstrated and was found to meet the program goals. The proof-of- principle test of the High Pressure Regenerative Turbine Engine produced data that agreed well with model s, enabling more confidence in designing future prototypes based on this concept. Emission levels were significantly reduced as predicted as a natural attribute of this power cycle. Engine testing over a portion of the operating range allowed verification of predicted power increases compared to the baseline.

15. NUMBER OF PAGES 14 . SUBJECT TERMS Gas turbine engines; Brayton cycle; Semi-closed cycle; PRICE CODE 16 .

Combustion; Gas turbine combustion A07 17. SECURIT Y CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT OFTHIS PAGE OF ABSTRACT OF REPORT Unclassified Unclassified Unclassified Standard Form 298 (Rev . 2-89) NSN 7540-01-280-5500 Prescribed by ANSI Std. 239-18 298-102