Document
ARL-TR-2113
NASA/TM--1999-209459
Wave-Rotor-Enhanced Gas Turbine
Engine Demonstrator
Gerard E. Welch U.S. Army Research Laborator_ Glenn Research Center, Cleveland, Ohio Daniel E. Paxson Glenn Research Center, Cleveland, Ohio Jack Wilson Dynacs Engineering Company, Inc., Brook Park, Ohio Philip H. Snyder Rolls-Royce Allison, Indianapolis, Indiana Prepared for the Gas Turbine Operation and Technology for Land, Sea and Air Propulsion and Power Systems Symposium sponsored by The North Atlantic Treaty Organization's Research and Technology Organization Ottawa, Canada, October 18-21, 1999 National Aeronautics and Space Administration Glenn Research Center
October 1999
Trade names or manufacturers' names are used in this report for identification only. This usage does not constitute an official endorsement, either expressed or implied, by the National Aeronautics and Space Administration.
Available from National Technical Information Sen, ice NASA Center for Aerospace Information 7121 Standard Drive 5285 Port Royal Road Hanover, MD 21076 Springfield, VA 22100 Price Code: A03 Price Code: A03 Wave-Rotor-Enhanced Gas Turbine Engine Demonstrator Gerard E. Welch U.S. Army Research Laboratory Glenn Research Center Cleveland, Ohio U.S.A.
Daniel E. Paxson National Aeronautics and Space Administration Glenn Research Center Cleveland, Ohio U.S.A.
Jack Wilson Dyancs Engineering Co., Inc.
Brook Park, Ohio U.S.A.
Philip H. Snyder Rolls-Royce Allison Indianapolis, Indiana U.S.A.
Abstract The U.S. Army Research Laboratory, NASA Glenn Research Center, and Rolls-Royce Allison are working collaboratively to demonstrate the benefits and viability of a wave-rotor-topped gas turbine engine. The self-cooled wave rotor is predicted to increase the engine overall pressure ratio and peak temperature by 300% and 25 to 30%, respectively, providing substantial improvements in engine efficiency and specific power. Such performance improvements would significantly reduce engine emissions and the fuel logistics trails of armed forces. Progress towards a planned demonstration of a wave-rotor-topped Rolls-Royce Allison model 250 engine has included completion of the preliminary design and layout of the engine, the aerodynamic design of the wave rotor component and prediction of its aerodynamic performance characteristics in on- and off-design operation and during transients, and the aerodynamic design of transition ducts between the wave rotor and the high pressure turbine. The topping cycle increases the burner entry temperature and poses a design challenge to be met in the development of tile demonstrator engine.
The wave rotor is considered an enabling technology for Introduction an alternative path to meet IHPTET (Integrated High The wave rotor is a self-cooled dynamic pressure Performance Turbine Engine Technology) III goals and exchange machine that can be embedded concentrically beyond. The lower fuel consumption also translates within a gas turbine engine to increase engine overall directly into reduced emissions which is a driving factor pressure ratio by 300c_ and peak temperature by 25 to in global civil aviation (cf. NASA Strategic Plan, 1998).
30% while maintaining rotating machinery temperature levels consistent with conventional materials and cooling The Army Research Laboratory (ARL), NASA John H.
technology. The topped engine is predicted to operate Glenn Research Center at Lewis Field (GRC), and Rolls- with substantially higher fuel efficiency (i.e., lower Royce Allison are working collaboratively to demonstrate specific fuel consumption, SFC) and power to weight-flow the benefits and viability of a wave-rotor-topped gas ratio (SP, cf. Welch et al., 1997). These improvements turbine engine. A successful wave-rotor/gas-turbine- translate into potential benefits for both the military and engine demonstration is a necessary step on the path to civilian sectors. The reduced fuel bum has far reaching mature wave rotor technology. To this end, a implications in terms of the fuel economy and logistics demonstrator engine is planned that will achieve trail of Army missions and, hence, is of strategic significantly improved performance, aggressively importance to the future U.S. Army (Elber et aL, 1997).
NA S A/TM-- 1999-209459 1
incorporate wave rotor technology intothe conventional
engine are first provided. The important results from two
gas turbine engine flowpath, utilize current materials and
contracted efforts performed to date by Rolls-Royce mechanical technology, utilize existing engine hardware Allison for NASA GRC on a wave-rotor-topped Rolls-
to a large degree, andintroduce minimal mechanical
Royce Allison model 250 engine, including details of the complexity intothe engine (cf.Snyder and Fish, 1996).
preliminary design and layout of the engine and the aerodynamic design of transition ducts between the wave
The Rolls-Royce Allison model 250was selected as the
rotor and the high pressure turbine, are then provided. An baseline engine because of its primary flow path on-going preliminary mechanical design and structural configuration andits engine component modularity, analysis of the rotor is then described. Finally, future including limited interchangeability andcompatibility design challenges are discussed.
among thecompressor, turbine, andgear boxes of the
several engine derivatives in production. Thischoice
reflects the realityof economic constraints andthe
Description of Wave-Rotor-Topping Cycle
resulting philosophy forthe demonstrator engine project
for Gas Turbine Engines
toallowsharp focus ondevelopment ofthewave rotor
without development of anynew,adapted, or scaled
Aeropropulsion Engine Application
turbomachinery hardware. When examining production
The wave rotor is a machine designed to exchange energy
turboshaft engines in terms of SFC atrated power as
efficiently between gas streams of differing energy density
shown in figure1,it is evident thatincorporation ofa
and, outside of the United States, is often referred to as a
waverotortopping unit into an engine embodying
pressure-exchanger, energy-exchanger, or Comprex ®1.Its
relatively mature technology can result inSFC levels well
operating principles and rich history have been described
below those attained byengines that use the state-of-the-
in detail elsewhere (see Azoury, 1992 and Kentfield, artturbomachinery presently onthe market.
1993). Interestingly, a wave rotor was first applied in a 0 70' topping cycle for a locomotive gas turbine engine (GTE) _v_S 117 shortly after World War II (Meyer, 1947) and only later was a subject of research and development for CT 63-M-5_ _ [] [] aeropropulsion (e.g., see Goldstein et aL, 1958). General []
{ Electric (ca. 1958-1963) and Rolls-Royce (ca. 1968-1972)
ncme H1400-1 tested wave rotors with the intention of topping or .,,.r,i,.I ,_ [] i replacing turbomachinery stages of small turboshaft -650C2 /" _TS 10' E engines like the Rolls-Royce Allison model 250 (cf.
& i _Tss,. -11 proceedings of the 1985 wave rotor technology _ PW206A ,_TM 333 2B --i--I i _ IT55-'L'714 O symposium at the U.S. Naval Postgraduate School, ,,_ @5O e"l Shreeve and Mathur, 1985). More recent related work E XA T703 3S4 "d_l'Jakila 1A1_ was focused on missile applications (Taussig and C Wav_ NO_OI J _ * Hertzberg, 1984) and industrial ground power plants D_mc,n _r a:o: , / \ i _TM322.31 _ (Zauner et aL, 1993). Since 1990, the benefits derived by _L_-8_ T4_-_ 0 40 I-- topping gas turbine engines for aeropropulsion has been O GESB- ,i a subject of research at NASA GRC. The research effort "5 has included experimental investigations to understand the O.- principal loss mechanisms of the component and to establish its operating map (Wilson, 1997 and Wilson, Source:1998AHS International Directory 0 SO, ........
1998), development and validation of computational tools I00 I000 I0000 for analysis and design (Paxson, 1995, Paxson, 1996, Maximum power at sea level, hp Welch, 1997b, Larosiliere, 1995), development of design/optimization procedures (Wilson and Paxson, Figure 1.--Predicted enhancement in turboshaft engine 1996, Welch, 1997a), system studies and mission analyses performance levels afforded by wave rotor topping.
(Jones and Welch, 1996), and a contracted effort with Rolls-Royce Allison toward the demonstrator engine A summary of progress toward demonstration of a wave- described herein (Snyder, 1996).
rotor-topped gas turbine engine is reported in this paper.
Descriptions of the wave rotor component and predicted benefits from cycle studies of the wave-rotor-topped Registeredtrademarkof Brown-Boveri.
NASA/TM-- 1999-209459 2 temperature. The total pressure of the gas delivered to the Component Description The wave rotor comprises a tip-shrouded rotor as shown turbine (i.e., the low-pressure exhaust port) is typically 15 to 20c_ higher than air delivered by the compressor (the in figure 2 that is surrounded by a stationary casing as low-pressure inlet port). A detailed description of the four- shown in figure 3. The casing endwalls are penetrated by inlet and outlet ducts that port gases of different pressure port wave rotor shown in figure 3 is provided elsewhere (see Welch et al., 1997).
and temperature to and from the rotor flow-annuli. The rotor hub, tip-shroud, and blade surfaces define rotor The wave rotor component is compatible with the high passages. Gasdynamic (shock and expansion) waves are temperature, high pressure conditions of the GTE topping initiated as the rotor passages open and close to the ported flows in a timed sequence set by the rotor speed and cycle application because of several key features: azimuthal location and extent of the ports. These waves compress and expand the gas as they propagate through Self-cooling.--The rotor surfaces are alternatively washed by the relatively low temperature compressor discharge the rotor passages. In the simplest configuration, the rotor passages are straight, at constant radius, and aligned with and high temperature burner discharge at frequencies the axis of rotation; the net shaft power of the machine is much higher than the material thermal-response-time. The rotor remains substantially (e.g., 25 to 30c_) cooler than zero like any gas generator spool. The rotative speed is the burner discharge; therefore, the burner discharge set by aerodynamic design trades and the corrected tip- speeds are typically low (e.g., 100 m/s [300 ft/s]). temperature of the topped engine is significantly higher than that of the baseline engine while the rotating Although the rotor flow field is inherently unsteady, the port flows are essentially steady and the wave rotor can be component temperatures are comparable.
closely integrated within other steady flow turbo- Low correctedflow.--The component is aerodynamically machinery components.
compatible with the low corrected specific flow rates Four-Port Wave Rotor for GTE Application supplied by the core compressors of modem aeropropulsion In the GTE topping application, fresh air from an upstream engines. The discharge from the full annulus of the compressor enters the wave rotor through the low-pressure compressor diffuser is ducted at nearly constant radius to inlet port. This air is compressed by shock waves as it the partial-annular port of the wave rotor. This flow traverses the rotor and cools the passage surfaces. The concentration accommodates aerodynamically efficient compressed air is discharged at the opposite end of the rotor rotor passage geometries. Futhermore, the rotor is shrouded to an external burner at a pressure typically three times so that tip leakage losses are eliminated.
higher than the compressor discharge. The burner exhaust gas reenters the wave rotor through the high-pressure inlet Low Rotative Speed.--Typical wave rotor corrected tip- port. As it traverses the rotor, the hot gas is expanded, heats speeds are a factor of five or six lower than those of modern the passage surfaces, and is discharged to a downstream turbomachines. The simple rotor geometry, the operating turbine. The hot gas temperature is typically reduced by 25 temperature, and the need to maintain acceptable hoop to 30% during this expansion process; that is, the burner stress levels suggest that ceramic rotors may be an attractive exhaust temperature is much higher than the turbine entry design choice (cf. Zehnder et al., 1989).
From _L \F" _2 /_
urner
• //Po_ :: /__:_"'TZurbine Fr°mll_ _ _ 7=" _]_Zo Comp ressor Burner Figure 3.--Four-port wave rotor schematic diagram.
Figure 2.--Rotor of NASA GRC four-port wave rotor experiment.
NASA/TM-- 1999-209459 3 Rapid Transient Response and Stabilit3'.--The wave rotor Wave-Rotor/GTE Concept responds (gasdynarnically) to transients in adjacent A systematic diagram of the Rolls-Royce Allison model components within a couple of rotor revolutions (e.g., ten 250 is shown in figure 4. This popular helicopter engine milliseconds). The fast response is quite independent of its is configured such that the centrifugal compressor instantaneous rotative speed, in contrast to turbomachinery discharge is ducted to the aft of the engine where it is components that must spool up or down. The prompt turned ninety degrees as it enters the combustor. The response has been demonstrated in Brown-Boveri's diesel burner discharges into the two-stage high-pressure turbine engine supercharger (Comprex®); engines fitted with the (HPT) that drives the compressor. A center gearbox links wave rotor responded faster to power demand than did the the low-pressure turbine (LPT) to a power output pad.
same vehicles fitted with a conventional turbocharger (see Exhaust gas is ducted out the top center of the engine. In Berchtold and Gull, 1960). Past research has suggested that a wave-rotor-topped configuration of this engine, the a wave-rotor-topping unit can enhance the dynamic stability wave rotor and associated ducting can be installed of the gas turbine engine (e.g., Taussig and Hertzberg, between the burner and HPT as shown in figure 5. The 1984). A numerical study by Greendyke et aL (1997) wave rotor diameter and length are both approximately showed that the wave-rotor-enhanced engine is indeed less equal to the tip diameter of the HPT. In the schematic likely to surge during rapid fuel flow changes than an diagram shown, the wave rotor spins coaxially on a untopped GTE. This stabilizing feature may allow wave- separate shaft at approximately one-third the speed of the rotor-topped engine operation at significantly reduced gas generator spool through its operating range (cf.
compressor surge margin. Snyder and Fish, 1996). Addition of wave rotor topping inherently requires alteration of the design-corrected flow A formidable set of technical challenges balance ihese rates of the HPT. Conveniently, the interchangeablility enabling features: leakage flows between the rotor and among the components of the Roils-Royce Allison model casing endwalls, noise associated with gasdynarnic waves 250 derivatives accommodates rematching with existing emitted into the ports, high cycle fatigue of both the rotor components. Within the constraints of project funding blades (due to unsteady loading) and the downstream blade levels, the planned demonstrator engine will be a back- rows (due to potential interactions), ducting and associated fitted, "breadboard" engine rather than a "clean-sheet" thermal and mechanical loads, a means to spin the rotor, design and build; however, the flexibility offered by the and the need for the wave rotor to supply high-pressure- baseline Roils-Royce Allison model 250 family allows for turbine cooling air in some engine applications. The significant enhancement of efig_ne performance levels technical challenges are being identified and addressed in a while using off-the-shelf components as described below.
step-wise manner through the systematic research program at NASA GRC and are addressed to some degree in the conceptual design of the demonstrator engine. Centrifugal Exhaust compressor collector Power turbine _ l High pres_re Overview of Progress Toward Y _ \ ]" [ / turbine Demonstrator Engin_ h Evaluation of the notional demonstrator engine project began with system studies at NASA GRC and at Rolls- Royce Allison. The assumptions made for component , __mbu_tor _ction performance are continually assessed in the light of _ _/J__ Compressor lessons learned from the in-house experiments and Y_If_ discharge analysis mentioned above. Concurrently, a general layout and preliminary design study of the wave-rotor-topped demonstrator engine is underway which has included detailed analysis of the wave rotor ducting and initiation Figure 4._Rolls-Royce Allison model 250 of the rotor mechanical design and structural analysis.
turboshaft engine.
NASA/TM-- 1999-209459 4 Compressor Section Gearbox Turbine Seclion Wave Rolor Section Combustor Secliorl . - ., p - _,_ --_:-_F-,._,.__I-_--_.
• _ _ ,_..__j FLOW .,.-.r ' -.y,_ ./_.:_ Figure 5. m Demonstrator engine combines wave rotor with modules from existing engine line.
the baseline engine. It should be noted that for the purposes Cycle Analysis Cycle studies carried out by Jones and Welch (1996) and of the study, no credit has been taken for surge margin enhancement as predicted by Greendyke et al. (1997).
Snyder and Fish (1996) predicted that wave rotor topping could enhance the SFC and SP of a "clean sheet" Rolls- Relative to the design point of the baseline engine, the wave rotor acts to increase the entry total pressure of the HPT. The Royce Allison model 250 engine by approximately -15 to -22% and +18 to +20%, respectively. Recent work at resulting decrease in the inlet corrected specific mass flow rate of the turbine (with turbine inlet temperature held ONERA shows similar predicted benefits (Fatsis and Ribaud, 1997). The total temperature-entropy diagrams for constant) was accommodated by replacing the model 250- C30 turbine section of the baseline engine with the turbine untopped (baseline) and wave-rotor-topped engines shown section of the model 250-C28C engine (cf. Snyder and Fish, in figure 6 suggest how these benefits are obtained. The 1996). An additional adjustment of 5% was also allowed by compressor pressure ratio, the burner energy addition, and affecting slight modification to specific sets of turbine the HPT inlet temperature are the same for both engines; however, because heat addition occurs at higher pressures hardware. In this way, flow matching was accomplished with existing components. The topped engine overall pressure ratio and temperatures in the topped engine, and because the was 23:1 relative to the baseline engine levels of near 8:1.
expansion and compression work in the wave rotor are The demonstrator engine was predicted to operate at 547 kW equal, the total pressure into the turbine of the wave rotor (733 hp) power levels at 12.7 mg/N-s (0.45 Ib,_f_jhp-hr) SFC topped engine is 15 to 20% higher than that of the untopped as compared to the baseline engine operation at 485 kW engine. The higher availability at the HPT inlet translates directly into increased engine power and efficiency. (650 hp) and 16.7 mg/N-s (0.59 Ibm.ruCJhp-hr). While the turbine entry temperature was maintained at the baseline engine level of 1328 K (1930 °F), the burner exhaust Wave rotor performance map.--In the cycle decks used, the burner and its associated loss is replaced by the wave temperature of the topped engine is predicted to be 1702 K rotor/burner topping unit and its associated total pressure (2605 °F). Note that the temperature of the gas discharged fi'om the wave rotor to the burner, 1237 K (1767 °F), is gain. The pressure ratio (turbine-entry/compressor-discharge) is represented by a performance map shown in figure 7 that Significantly higher than the burner inlet temperature of the was computed by Paxson using his Q-I-D model (Paxson, baseline engine, 848 K (1067 °F). This increase in burner 1996). The wave rotor pressure ratio is plotted as a function inlet temperature, as well as burner pressure and outlet of corrected rotor speed, flow, and burner heat addition. The temperature, will require a more advanced burner design than rotor geometry and speed were set by the optimization that used in the model 250 production engine. An effusion • ®_ procedure proposed by Wilson and Paxson (1996). For a cooled or Larmlloy - based design along with a change in liner material are candidates for use in the demonstrator test typical optimized rotor, solidity is near 15, hub-to-tip ratio is near 0.7, rotor length to diameter is near unity, and corrected engine. Development of an appropriate combustion system is rotor speeds are near 100 m/s (300 ft/s). targeted to he an important segment of the continuing effort toward a successful demonstrator engine.
Design point operation.--The compressor surge margin in the wave-rotor-topped engine was maintained equal to that of -'Reg slered trademarkof A 1son Engine Company,Inc.
NASA/TM-- 1999-209459 5 Wave-rotor-enhanced engine ............... Baseline engine /P'r-_ P .......
E (D increased // _ j/ / /" / k.- peek / _ w. / /."
temperature ir _ ) // i'" High Pressure T ................................................................................. -_ ........................ :_:""_ .......... j_.. turbine Inlet /. _r_. / _ temperature /" / o Topped cycle _ J:..l _" "1 _ _#.p'r = '_"Com,.
....--_ ,_ ........ -, ......... x....
_..,..,,_.._ 7" EnhMlced _ : 8aeelltl. er'tglr_e C / _/_"- power L . . 1 . . _ : : : . Y . . . .
T C ...... _'ne cycle _:_:.._..Pp.__r ,nh ....... t " *com., I) Entropy Figure 6.--Temperature-entropy diagram showing thermodynamic benefit of wave-rotor-topping cycle.
2.0 .............................. 7 ............................... 3.8 i O 1.9 3.$rr q) 1.8 3,4 _) 4" Design Point _ Lk_, "_ , Take Off -_-_K '_ _tL a, ",,, 1 "O MaxCrui._j _i_ : _ q_..
3.2_" __c.,.,._:_ ......... _.__'-- .......... _. ............
1.7 Cruise B _ t_ _ : • i "'" c- 50% s.P _, 6 "_ : [ ....... T
E
3.0 i__.
'3 "3 ......... ..... * ....... i.......... ..............
#- i _, i
1.4 ............................
2,6 Ii Norrr_,,zed Dimensionl_s c:
° '
'_ 1.3 "--O_r_0Jr"P-°t_i--- 0.50 2.4 ¢- .- ..__2_L;_! ____!L _ ........... '_ _o_ • _ _._ 2.2 2.0 D:: • _ ; _ i I_Topped Engine E _= i i '='Ope,ating Points, 1.0 1,8 2 3 4 5 Ratio of Turbine Inlet and Wave Rotor Inlet Temperature Figure 7.--Wave rotor map with demonstrator engine operating points displayed.
NASA/TM--1999-209459 6 demonstrator engine could be assembled using existing Off-design operation.--The cycle deck model was exercised over six steady-state power settings ranging hardware and that the greatest challenge lies in the design of the burner due to the high burner inlet temperatures.
from idle to take-off power. The wave rotor contributes a nearly constant 3:1 overall compression ratio over the Other details of the preliminary design and general layout effort include the follow5ng: operating line. The corrected rotor speed (free-wheeling) and corrected flow of the wave rotor are essentially constant over the operating line. The predicted SP and Mechanical components.--The preliminary design and SFC of the topped and baseline engines vary as shown in layout indicated that the C30 compressor and C28C turbine units are mechanically compatible and that the figure 8. Evidently, the benefits of wave rotor topping are gearboxes of the two engine variants are interchangeable.
maintained at part power. At idle the SP is increased by 19% and the SFC is reduced by 32%; further, a sensitivity The fundamental layout of the Rolls-Royce Allison model 250 minimizes the impact of the increased thrust on the analysis showed that engine operation was found acceptable at idle for a range of power turbine speeds. component because the compressor and turbine thrusts are The temperature difference between the bumer exhaust carried on separate bearings. The performance of the free wheeling wave rotor is found to be nearly optimum; that and the HPT inlet gas is essentially a constant 330 K (600 °F) from idle to full power; unfortunately, the burner inlet is, the performance levels of a wave rotor on the passive speed schedule are nearly the same as those attained on a temperature of the topped engine operates between 330 K metered speed schedule. The wave rotor can therefore to 555 K (600 to 1000 °F) higher than the baseline engine over the operating envelope. As mentioned above, these spin on an independent shaft.
higher temperatures push the burner beyond the technology regime of the baseline engine. Cooling air requirement.--The baseline engine requires approximately 2% cooling within the turbine section.
0.9 This air is derived from compressor discharge at 625 K (665 °F) and at a supply pressure approximately 4% above the HPT entry total pressure due to the presence of the conventional burner liner pressure drop. In the wave- _0.8 Baseline engine rotor-topped engine, the compressor discharge air is Wave-rotor-topped engine approximately 20% below that of the HPT entry total E e_ pressure. If this engine utilized a highly cooled first stage nozzle or blade employing internal impingement or a o 0.7 serpentine cooling scheme, an alternate supply of cooling E air would need to be developed. However, the end-point "-,1 t- use pressures of the cooling air of the model 250 engine O o 0.6 are significantly below that of the turbine entry total pressure. Thus for the demonstrator engine, modifications to the supply circuit of the cooling air will suffice in (9 delivering required cooling air via bleed from the o. 0.5 compressor discharge stream. However, when applying (D wave rotor topping to a "clean sheet" design engine, established practices of delivery of turbine cooling will need to be reexamined.
0.4
4 o 6 o 80o
2O0 Ducting.--The wave rotor device inherently requires the Shaft Horsepower (hp) use of adaptive ducting to rout flows between the partial Figure 8.--Off-design performance of baseline and annular stations at the wave rotor and the typically full wave-rotor-topped engines.
annuli at the conventional turbomachine interfaces. These transitions--compressor to wave rotor, wave rotor to burner, burner to wave rotor, and wave rotor to Preliminary Design and General Layout HPT--must occur with minimal aerodynamic loss and yet The preliminary design work has addressed mechanical in a manner that minimizes the length and wetted- aspects of component matching, flow path and component perimeter added to the engine. The baseline engine reconfiguration, wave rotor ducting, the advanced already utilizes a non-annular compressor transfer duct combustor design, and identification of the wave rotor that is very conducive to adaptation to the wave rotor inlet component and adaptive engine parts (Snyder, 1996).
port. The wave rotor component itself sets the allowable The major conclusion of the study was that the wave rotor NASA/TM-- 1999-209459 7 pressure loss fromwave rotor high-pressure exitport to higher than the 3.25% assumed in the cycle studies. The high-pressure inlet port (burner exit station) at 8.9ck. Less addition of the non-diffusing section increases the duct than half of this will be typically used across the burner length and area and in some applications might ultimately liner, with the remainder able to be attributed to the push the design choice toward a duct with a shorter non- remaining ducting. The design of this ducting beyond the diffusing length. In addition to higher aerodynamic loss initial space claim considerations has yet to be addressed. levels, the Type 2 and Type 3 ducts preserve the The transition duct between the wave rotor and the HPT tangential nonuniformity of the wave rotor exhaust flow; was seen to be of critical importance and was studied in the residual nonuniformities in total pressure and detail as reported below. temperature are reflected in the predicted loss in HPT efficiency and large percentage increase in total engine Detailed Wave-Rotor-to-HPT Transition Duct Design weight to shaft horsepower metric as shown.
Common to all wave rotors is the need to port flow between the partial sectors of the wave rotor and the full annuli of the surrounding turbomachinery (see figure 3).
A preliminary design and analysis of the transition duct between the low pressure exhaust port of the wave rotor and the HPT was carried out by Rolls-Royce Allison (Gegg and Snyder, 1998 and Weber and Snyder, 1998).
The wave rotor has two duct sets. Each transition duct ports half (1.12 kg/s, 0.26 kg/s-corrected) of the engine mass flow rate from a 45-degree sector to a 180-degree half annulus. The turbine half-annulus flow area is nearly twice that at the wave rotor exhaust port. The transition (a) Type I---diffuser with existing nozzle ring.
must be accomplished in a short length so as to minimize engine added-length and weight, The flow exits the wave rotor at an average of Mach 0.5 and swirl angle of 19 degrees (from the axis of rotation) and enters the HPT rotor at Mach 0.7 and swirl angle of 65 degrees. To complicate matters further, the wave rotor discharge is highly nonuniform tangentially in total pressure, temperature, and axial velocity. The principal objective of the design was a low 2_o0/p0transition in which the HtrF nozzle was incorporated into the ducting to the extent possible. Surface area, length, and weight added to the engine were also importan! metrics. The initial designs established using the volute procedure of Frolov and Golubtsov (1972) were improved upon by using (b) Type 2---nozzle with integral turning vanes.
knowledge gained from the results of 3-D computations with the OVERFLOW code (Buning, 1998).
Three duct concepts were considered: diffusing duct followed by the conventional nozzle ring of the model 250-C28 turbine (Type 1); converging duct and nozzle with integral turning vanes (Type 2); and a rapid turning elbow with volute (Type 3). A number of variants were considered within each of the three concepts: co- and counter-rotating wave rotor and HPT turbine, non- diffusing and non-turning sections at the inlet of the duct, and vaned and vaneless volutes. The best performers of the three concepts are shown in figure 9 and important parameters from the study are summarized in table 1. The duct concept in which the flow is diffused before entering (c) Type 3---elbow and volute.
the nozzle ring (Type 1) is considered best based on aerodynamic loss levels. The Type 1 duct with a non- Figure 9.--Designs for each of three types of ducts for transition from wave rotor to high pressure turbine.
diffusing section has ApJp0 = 3.7% that is only slightly NAS A/TM---- 1999-209459 8 from transition duct design analysis.
........ Table L-_summaDi of key parameters
Type 3 Type 1 Type 2 Volute Diffuser with Converging duct baseline nozzle ring and nozzle ,,,,, , 1.04 1.46 0.667 1.08 0.667 Normalized duct length _ Normalized duct area b 1.92 2.77 1.51 2.23 1.65 0 3I 50 75 0 Non-diffusing length (% of total duct length) Vane count 5 8 0 12 12 -- 7.8 6.0 -- 1.1 Turbine penalty (%)c -- 27 26 -- 4.6 %-increase engine weight/SHP 5.5 4.9 7.0 4.1 3.7 _gpo (%) ....
aduct length divided by wave rotor "rotor" length.
bduct area divided by area of baseline engine nozzle ring.
Cpercentage decrease in HPT adiabatic efficiency due to residual temperature nonuniformity.
Rotor Mechanical Design and Structural Analysis mechanical design of the rotor is to be identified in the Some past efforts in wave rotor testing have listed initial stages of the mechanical design effort.
mechanical shortcomings including rotor durability as significant issues in making the concept serviceable Detailed heat transfer analysis of the rotor at the design- (cf. Shreeve and Mathur, 1985). Such rotor design issues point and during a start-up transient will be conducted to are currently under detailed examination at Rolls-Royce establish the design point rotor temperature and limiting- Allison. The goal of this effort is to develop a preliminary case thermal gradients in the rotor. Rotor stress and design of the rotating hardware capable of operating in the dynamic analysis will include rotational loads, drum wave-rotor-topped demonstrator engine previously internal and external pressure loads, thermal stress loads, described. Both the design methodology and the particular and passage way transient pressure/wall dynamic rotor design under development for the demonstrator behavior. Life predictions determined according to low- engine constitute new wave rotor technology that is and high-cycle-fatigue, stress rupture, and creep growth applicable to both the demonstrator and potential engine allowables will be compared to appropriate rotor life products. This effort is currently proceeding with criteria. A refinement to the rotor design will be identification of a candidate rotor design, preliminary formulated to conform to the rotor life requirements and rotor mechanical design, detailed rotor heat transfer life will be verified by revisiting both the heat transfer and analysis, and detailed stress and dynamics analysis.
stress/dynamics analysis. Limitations on analysis arising from any generic inadequacy of current gas turbine As stated earlier, the goals of the demonstrator engine analysis tool capabilities as applicable to wa,_e rotors will effort reach beyond that of just making an engine that will be identified.
run. The rotor fabrication techniques selected must be suitable for carry over to production hardware with regard to both producibility and ability to meet rotor life.
Summary Novel rotor and bearing mechanical designs are being The predicted benefits offered by wave rotor topping are considered to address rotor end clearance control and significant. The wave-rotor-topping unit is predicted to thermal growth/stress realities adequately. Currently, the increase the overall pressure ratio of the baseline Rolls- study is addressing the important aspects of rotor dynamics, Royce Allison model 250 engine from 8:1 to 23:1 and heat transfer, and stress concerns throughout the rotor on a burner exhaust temperature by 25 to 30%, leading to preliminary basis. Based on one-dimensional transient gas predicted SFC reduction of nearly 23% with concomitant predictions for the aero flow path, this effort will also power-to-weight-flow enhancement of 13%. The wave include a transient analysis of the rotor passage walls in rotor demonstrator engine is a necessary step toward the order to determine the potential of thermal stress and maturation of wave rotor technology for application in thermal shock of the rotor walls. A viable preliminary NASA/TM--1999-209459 9 aeropropulsion engines. Progress to-date has included Larosiliere, L.M., 1995, "Wave Rotor Charging Process: Effects of Gradual Opening and Rotation," Z of Propulsion and detailed cycle analysis, engine preliminary design and Power, II, No. 1. pp. 178-184.
general layout, detailed analysis of the critical wave-rotor- Meyer, A., 1947, "Recent Developments in Gas Turbines," to-HPT transition duct, and initiation of rotor preliminary Mechanical Engineering, 69. pp. 273-277.
design and structural and thermal analyses.
NASA Strategic Plan, 1998, NASA Policy Directive, NPD-1000.1.
The anticipated three-year project '*ill require that several Paxson, D.E., 1995, "Comparison Between Numerically technical challenges be overcome. The higher burner inlet Modeled and Experimentally, Measured Wave-Rotor Loss temperatures associated with the topping cycle pose Mechanisms," J. Propulsion and Power, 11, No. 5, September- October, pp. 908-914.
significant material challenges for the burner liner. The Paxson, D.E., 1996, "Numerical Simulation of Dynamic development of this combustion system, including Wave Rotor Performance," J. Propulsion and Power, 12, No. 5, adaptive ducting and inlet guide vanes between the burner September-October. pp. 949-957.
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Snyder, P.H. and Fish, R.E., 1996, "Assessment of a Wave Azoury, P.H., 1992, Engineering Applications of Unsteady Rotor Topped Demonstrator Gas Turbine Engine Concept," Flow, Wiley, New York.
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NASAffM--1999-209459 I 0 REPORT DOCUMENTATION PAGE Form Approved OMB No. 0704-0188 Public reporting burden for thiscollection of information is estimated to average 1 hour per response, includingthe time for re'dewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information, Send comments regarding this burden estimate or any other aspect o1this collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503 ; 3. REPORT TYPE AND DATES COVERED 1. AGENCY USE ONLY (Leave blank) 2, REPORT DATE October 1999 Technical Memorandum 4. TITLE AND SUBTITLE 5. FUNDING NUMBERS Wave-Rotor-Enhanced Gas Turbine Engine Demonstrator WU-523-26-33-00 iL162211A47A 6, AUTHOR(S) Gerard E. Welch, Daniel E. Paxson, Jack Wilson, and Philip H. Snyder 8. PERFORMING ORGANIZATION 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) REPORT NUMBER NASA Glenn Research Center Cleveland, Ohio 44135-319 I and E-11958 U.S. Army Research Laboratory Cleveland, Ohio 44135-3191 10. SPONSORING/MONITORING 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) AGENCY REPORT NUMBER National Aeronautics and Space Administration Washington, DC 20546_)001 NASA TM--1999-209459 and ARL-TR-2113 U.S. Army Research LaboraTory Adelphi, Maryland 20783-1145 11. SUPPLEMENTARY NOTES Prepared for the Gas Turbine Operation and Technology for Land, Sea and Air Propulsion and Power Systems Symposium sponsored by The North Atlantic Treaty Organization's Research and Technology Organization, Ottawa, Canada, October 18-21, 1999.
Gerard E. Welch, U.S. Army Research Laboratory, NASA Glenn Research Center, Cleveland Ohio; Daniel E. Paxson, NASA Glenn Research Center, Cleveland, Ohio; Jack Wilson, Dynacs Engineering Company, Inc., Brook Park, Ohio; and Philip H. Snyder, Rolls-Royce Allison, Indianapolis, Indiana. Responsible person, Gerard E. Welch, organization code 0300, (216) 433-8003.
12b, DISTRIBUTION CODE 12a. DISTRIBUTION/AVAILABILITY STATEMENT Unclassified - Unlimited Subject Categories: 01 and 07 Distribution: Nonstandard This publication is available from the NASA Center for AeroSpace Information, (301) 621-0390.
13. ABSTRACT (Maximum 200 words) The U.S. Army Research Laboratory, NASA Glenn Research Center, and Rolls-Royce Allison are working collaboratively to demonstrate the benefits and viability of a wave-rotor-topped gas turbine engine. The self-cooled wave rotor is predicted to increase the engine overall pressure ratio and peak temperature by 300% and 25 to 30c_, respec- tively, providing substantial improvements in engine efficiency and specific power. Such performance improvements would significantly reduce engine emissions and the fuel logistics trails of armed forces. Progress towards a planned demonstration of a wave-rotor-topped Rolls-Royce Allison model 250 engine has included completion of the preliminary design and layout of the engine, the aerodynamic design of the wave rotor component and prediction of its aerodynamic performance characteristics in on- and off-design operation and during transients, and the aerodynamic design of transition ducts between the wave rotor and the high pressure turbine. The topping cycle increases the burner entry temperature and poses a design challenge to be met in the development of the demonstrator engine.
14. SUBJECT TERMS 15. NUMBER OF PAGES Wave rotor; Gas turbine engine 16. PRICE CODE AQ3 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT 17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION OF ABSTRACT OF REPORT OF THIS PAGE Unclassified Unclassified Unclassified Standard Form 298 (Rev. 2-89) NSN 7540-01-280-5500 Prescribed by ANSI Std, Z39-18 298-102