† PhD Student, Power and Propulsion, Cranfield University 1 Copyright © 2013 by ASME
HELICOPTER MISSION ANALYSIS FOR A REGENERATED TURBOSHAFT
Ali Fakhre†, Vassilios Pachidis, Ioannis Goulos, Mahmood Tashfeen, Pericles Pilidis
School of Engineering, Department of Power and Propulsion, Cranfield University, Cranfield, Bedford, MK43 0AL, UK
Email: [email protected]
ABSTRACT The aim of the study presented in this paper, is to compare helicopters employing simple cycle turboshaft engines, with helicopters employing novel regenerated turboshafts. Two existing helicopter configurations, a Twin Engine Light and a Twin Engine Medium are compared against regenerated configurations. The reference installed engines of both helicopters are notionally optimized by incorporating a heat exchanger, which enables heat transfer between the exhaust gas and the compressor delivery air to the combustion chamber. This process leads to a lower fuel input requirement as well as higher overall thermal efficiency compared to the reference simple cycle engine.
The benefits arising from the adoption of the heat exchanger for both configurations are firstly presented by conducting part-load performance analysis for each optimized engine against its reference simple cycle engine. The obtained results suggest substantial reduction in specific fuel consumption for a major part of the operating power range with respect to both helicopter configurations. The results also demonstrate that the heat exchanger effectiveness is a critical parameter in achieving further reductions in specific fuel consumption.
The study is further extended to investigate mission fuel burn saving limits for both helicopter configurations under the simulated part-load performance conditions by conducting a heat exchanger tradeoff study. The weight estimation correlation for the heat exchanger is adopted from the previously reported studies of similar fashion and is simulated accordingly for both helicopter configurations. A multi-disciplinary simulation tool with an integrated range of capabilities applicable to helicopter performance evaluation and mission analysis is adopted to simulate various types of missions, targeting wide range of helicopter operations. The results of the mission analysis suggest that the regenerated counterpart configurations are capable of achieving significant reductions in mission fuel burn. However, the level of gain from mission fuel burn savings is dependent on the selected helicopter mission profile, the recuperator design effectiveness as well as the overall evaluation criteria. The results also conclude that while the amount of benefit is dependent on various parameters, there is always an optimum “saving” region for each mission that justifies the need for regeneration.
1. INTRODUCTION The aero-industry has had many significant challenges since the beginning of the 21
st century, according to Colin
F. McDonald [3], the most salient ones being the reduction in emissions, improvement in Specific Fuel Consumption (SFC), reduction in noise levels and achievement of efficient and most economical life cycle costs. Currently the prominent increase in environmental concerns has demanded the aviation industry to enhance the operational life of the powerplants (aero-engines) and has also strongly positioned the aviation industry to innovate and produce more sustainable and “low carbon” environmental friendly solutions.
In response to the demand and to extend and maintain effective rapid transition to the sustainable and greener aviation industry (specifically in Europe) the ACARE (Advisory Council for Aviation Research and Innovation in Europe) has set some ambitious and complex targets for Vision 2020, under the strategic research and innovation agenda [1]. Vigorous programs of Aeronautics and air transport research are already underway, delivering important initiatives and benefits for the aviation industry [1]. This has led to a renewed interest in heat exchanged engine concepts an effective alternative to make future air transportation more environmental friendly. Benefits of the regenerated engines and a complete evaluation of regenerative powerplants is reported by Colin F. McDonald in Recuperated Gas Turbine Aero-engines, Part I [2], Part II [3] and Part III [4].
Regenerated engines are recognized as one of the most promising alternative (Aero-engine) powerplant configurations when targeting significant reductions in fuel burn and lower emissions. The consideration of replacing a simple cycle engine with a regenerative engine in the current industry climate (increased environmental concerns and high fuel costs) can significantly enhance the integrated economical and operational performance of the existing Simple Cycle (SC) engine helicopter. However, the level of benefits offered by a rotorcraft regenerative powerplant solely depends on the type of operation and range the rotorcraft is designed to serve [5]. One of the major disadvantages of today’s rotorcraft
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powerplant is the increased specific fuel consumption during part power. A typical helicopter cruises between 55% to 65% of installed power and more importantly this particular flight regime represents around 80% to 90% of the helicopter’s mission [4]. One of the most effective technique to overcome this challenge is by adopting the regenerative powerplant, as this offers a clear benefit during part-power operation.
2. REGENERATED TURBOSHAFT Both helicopter configurations investigated for the purpose of this study are notionally optimized by adding a Heat Exchanger (HE) in their existing simple cycle turboshaft, demonstrating a regenerated turboshaft engine.
The regenerated turboshaft incorporates a HE (Fig. 1); hot side is placed downstream of the Free Power Turbine (FPT) and cold side upstream to combustion chamber. This arrangement enables heat transfer between the exhaust gas and the compressor delivery air of the combustion chamber. Depending on the heat transfer rate defined by the HE design effectiveness, an increase in (working fluid) compressor delivery air temperature is achieved. This process of preheating upstream to combustion chamber leads to lower fuel input requirements compared to the reference simple cycle engine. However, one of the side-effects resulting from the incorporation of the HE is the additional pressure losses introduced by the heat transfer process and by the installation arrangement of the HE.
Figure 1: Schematic Layout of a Single-spool Regenerated
Turboshaft
3. HECTOR SIMULATOIN FRAMEWORK A comprehensive and simultaneously cost-efficient simulation framework targeting the performance of integrated rotorcraft – engine systems has been developed at Cranfield University under the name of HECTOR [6]. HECTOR “HeliCopTer Omni-disciplinary Research-platform” is an integrated tool consisting of a rotorcraft flight mechanics code and an engine performance code (TURBOMATCH). HECTOR has been extensively used in the past for rotorcraft engine cycle design optimization
studies in [10]. A brief description of the individual codes is provided below.
3.1 HECTOR HECTOR is designed to calculate the helicopter’s steady state (trim) and dynamic (manoeuvre) performances. For the purpose of this study the helicopter is assumed to be operating in trim during each segment of each mission task element. No dynamic performance is taken into account.
3.2 TURBOMATCH TURBOMATCH is a gas turbine performance code developed by Cranfield University, TURBOMATCH is capable of calculating both steady state and transient gas turbine performance. The simulation is purely based on zero-dimensional modeling of the various thermodynamic processes occurring within the various engine components. For the purpose of this study the engine is assumed to be operating strictly at steady-state design point and off-design conditions.
4. SIMULATION METHODOLOGY The schematic representation of the integrated tool is illustrated in Figure 2. Each defined mission profile is translated into discrete segments based on user defined input values. The initial All Up Mass (AUM) is equal to the sum of the Operational Empty Weight (OEW), the useful payload, and the on-board fuel supplies. The required amount of fuel for a given mission has to be initially assumed; therefore an initial guess is made for the weight of the on-board fuel supply which is then refined through an iterative process.
For each flight segment HECTOR calculates the engine power requirement, intake inlet conditions, and updates the new space-wise position of the rotorcraft. TURBOMATCH subsequently establishes the engine’s operating point to meet the power demand required. Thus, the engine fuel flow can be established for the given flight condition. During this process, the total aggregate of fuel flow with respect to time is calculated implicitly from zero up to the current mission flight segment.
5. HELICOPTER ENGINE CONFIGURATIONS Two helicopter categories are optimized as follows:
1. A Twin Engine Light (TEL), multipurpose helicopter, based on the configuration of MBB BO105 (now under Eurocopter).
2. A Twin Engine Medium (TEM), Utility and Rescue helicopter based on the configuration of PUMA SA330.
A Twin Engine Light helicopter model based on the configuration of the MBB Bo105 was implemented in HECTOR. This rotorcraft is equipped with two Rolls Royce Allison 250-C20B turboshaft engines. Thus, the corresponding, Reference Allison 250-C20B (Engine A1)
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0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
SFC
µg/
J
Relative Power (P/PDesign)
SC REC e40% REC e50%REC e60% REC e70% REC e80%
SC=Reference Simple Cycle REG = Regenerated e = HE effectiveness
Figure 2: HECTOR Integrated Tool Framework and a Regenerated Allison 250-C20B (Engine B1) engine models were introduced and implemented in TURBOMATCH. The design point comparison for various engine parameters for Engine A1 and its notional counterpart, Regenerated Engine B1 are presented in Table 1.
The Twin Engine Medium helicopter model based on the configuration of the PUMA SA330 has also been implemented in HECTOR. This specific rotorcraft is equipped with two Turbomeca Turmo IV-C turboshaft engines. Therefore, the corresponding Reference Turbomeca Turmo IV-C (Engine A2) and its notional Regenerated Turbomeca Turmo IV-C (Engine B2) engine models were also introduced and implemented in TURBOMATCH. The design point comparison for various engine parameters for Engine A2 and Engine B2 is presented in Table 2. For both Reference engines, Engine A1 and Engine A2, design point is defined at take-off conditions and is validated against public domain [7].
Table 1: Design point, engine parameters for
Reference Engine A1 and Regenerated Engine B1 turboshafts
Engine Design Parameters
Reference Engine A1
Regenerated Engine B1
Pressure Ratio 7:1 7:1 Mass Flow (Kg) 1.56 1.56 TET (K) 1470 1470 TO Power (KW) 313 313 SFC @ TO (µg/J) 119.8 95.55 Dry Weight (Kg) 75 75 HE Weight (kg) - 9.75 HE ε % - 40 Total Weight (Kg) 75 84.75
Table 2: Design point, parameters for Reference
Engine A2 and Regenerated Engine B2 turboshafts
Engine Design Parameters
Reference Engine A2
Regenerated Engine B2
Pressure Ratio 5.8:1 5.8:1
Mass Flow (Kg) 5.9 6.1
TET (K) 1335 1350
TO Power (KW) 1163 1163
SFC @ TO (µg/J) 106.54 89.75
Dry Weight (Kg) 225 225
HE Weight - 38.35
HE ε % - 40
Total Weight (Kg) 225 263.35
Figure 3 presents the variation in SFC at design point and at part-load (P/PDesign) for Reference Engine A1 and Regenerated Engine B1 turboshaft. The heat exchanger effectiveness was varied between 40% and 80%. A significant improvement in SFC for regenerated Engine B1 relative to Reference Engine A1 at part power is apparent. For an assumed HE effectiveness range from 40% to 80% the SFC improvement at medium part power lies between approximately, 19% at 40% HE effectiveness up to 43% at 80% HE effectiveness. The Overall Pressure Ratio (OPR) for both Engines A1 and B1, at design point is 7:1, TET is 1470K and mass flow is 1.56 kg.
The design point and part-load SFC comparison for Reference Engine A2 and Regenerated Engine B2 turboshaft is shown in Figure 4. The SFC reduction at medium part-power (P/PDesign) is approximately 18% at 40% HE effectiveness and increases up to 40% at 80% HE effectiveness for regenerated engine B2 respectively.
Figure 3: SFC VS (P/P.Design) for Reference Engine A1 and Regenerate Engine B1 turboshaft
HECTOR “HeliCopTer Omni-
disciplinary
Research-platform”
TURBOMATCH
ROTORCRAFT
INPUT FILE
MISSION INPUT FILE
MISSION FUEL BURN
REQUIRED POWER, BASED ON FLIGHT
CONDITION
ENGINE MODEL
INPUT
ITERATIONS
Fu
el F
low
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SC REC e40% REC e50%REC e60% REC e70% REC e80%
SC = Reference Simple Cycle REC = Regenerated e = HE effectiveness
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Heat Exchanger Specific Weight LOWER LIMITHeat Exchanger Specific Weight UPPER LIMIT
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s B
1 (k
g)
Heat Exchanger Effectiveness ε %
Gross Weight Lower Limit
Gross Weight Upper Llimit
Engine Mass Flow =1.56kg
Figure 4: SFC VS (P/P.Design) Reference Engine A2 and Regenerated Engine B2 turboshaft
One of the drawbacks of incorporating a HE is that
it introduces additional pressure losses, causing reduction in engine shaft power. The amount of increase in pressure loss introduced by the HE is mainly dependent on how well the heat transfer process is arranged between the exhaust gas and the compressor delivery air to the combustion chamber as well as the amount of total engine mass flow through the HE. For Reference Engine A1 “Bolted on” type regenerators are assumed similar to the one reported in [5] and for Reference Engine A2 a fixed geometry tubular type heat exchanger is assumed as reported in [9].
The additional pressure losses introduced by the heat exchanger are assumed and are modeled accordingly for both regenerated engines. Similar values of pressure losses are used as reported previously, e.g. by the Allison T63 regenerative program [5] and by Grieb, H and Klussman. W, (May 1981).Table 3 and Table 4 present the pressure losses assumed as a function of HE effectiveness. (Note: previous studies only report pressure loss at 60% HE effectiveness, the pressure losses below and above 60% HE effectiveness are based on best engineering judgments).
6. WEIGHT ESTIMATION The HE weight correlation used for the purpose of this study is adopted from the previously reported study by Nicolas C. Kailos [8].The correlation is presented in Figure 5. It needs to be emphasized that the HE weight correlation adopted for the purpose of this study is for the fixed surface HE concepts.
The weight added by the HE for Regenerated Engines, B1 and B2 is derived using the correlation presented in Figure 5. Both helicopter configurations investigated in this study represent Twin engine configuration, therefore the gross heat exchanger weight for each helicopter is extended for “two” engines. The extended correlation for Regenerated Engine B1 is presented in Figure 6 and for Regenerated Engine B2 is presented in Figure 7.
Figure 5: Fixed geometry tubular type heat exchanger specific weight correlation adopted from [7]
The correlation shown in Figure 6 is used as an input to define the weight of Engine B1 and Figure 7 is used for Engine B2. For the purpose of this study the upper limit of the heat exchanger weight is defined as installed weight with respect to both helicopter configurations. It is evident from both correlations (Fig. 6) and (Fig. 7) that the gross weight of the HE is sensitive to the total engine mass flow and it significantly increases above 60% HE effectiveness. This is one of the reasons that all previously reported studies so far on Regenerated helicopters, utilize heat exchanger design effectiveness of around 60%. Purely due to the fact that the HE matrix weight and volume increase by a factor of approximately two, from a HE effectiveness of 60% to 75% [2].
Figure 6: Regenerated Engine B1 turboshaft, heat
exchanger gross weight correlation for two engines
7. CASE STUDIES In the context of this study several mission analyses (case studies) were conducted targeting various helicopter operations. Four generic reference missions were defined between two classes of helicopter configurations as detailed in Table 5.
Each mission scenario and flight segment throughout the mission profile is defined so that it reflects the realistic capability of both helicopters investigated. The power requirements and mission time and range are kept the same for the both helicopter types, the Reference Simple Cycle and the Regenerated helicopter. The mission profiles for velocity and altitude for reference mission A, C and D are presented in (Figure. 8, 9, 10).
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or T
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ine
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Engine Mass Flow 5.9kg
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ise
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Figure 7: Regenerated Engine B2 turboshaft, heat
exchanger gross weight correlation for two engines
Table 3: Regenerated Engine B1 turboshaft, heat exchanger pressure loss, added weight, and SFC@TO
Heat
Exchan
ger ε %
Total
Pressur
e Loss
%
SFC
@
TO
Weight Added by HE (kg)
40 5 95.56 9.75
50 6 89.37 12.75
60 6 83.16 20.25
70 8 76.9
3
33
80 9 70.68 58.5
Table 4: Regenerated Engine B2 turboshaft, heat exchanger pressure loss, added weight and SFC@TO
Heat
Exchanger
ε %
Total
Pressure
Loss %
SFC @
TO Weight
Added by HE (kg)
40 5 89.75 38.35
50 6 84.27 50.15
60 6 78.28 79.65
70 8 72.51 129.8
80 9 66.73 230.1
Table 5: Reference Missions and Representative
Rotorcraft Configuration
ID Reference
Mission RC MTOW
A Passenger
Transport /Air Taxi
TEL
2500kg
B EMS TEL 2500kg
C Oil and Gas TEM 6000kg
D Civil SAR (SAR) TEL
&
TEM
2500kg
6000kg
Figure 8: Reference Mission A, Passenger Mission, Altitude and Cruise speed Mission Profile
Figure 9: Reference Mission C, Utility Oil & Gas Mission, Altitude and Cruise speed Mission Profile
Figure 10: Reference Mission D, Search and Rescue Altitude and Cruise speed Mission Profile
8. RESULTS AND DISCUSSION
8.1 Heat Exchanger Effectiveness
The HE effectiveness is a measure of the heat transfer capability of the HE. A heat exchanger with higher heat transfer ability represents high HE effectiveness. In case of a fixed geometry tubular type HE, the higher effectiveness represents the higher matrix weight of the HE. This is evident from Figure 5. The specific fuel consumption decreases linearly as a function of increased effectiveness. However the “added weight” of the HE increases nonlinearly as a function of increased HE effectiveness. 8.2 Regenerated Helicopter Requirements
To replace an existing simple cycle helicopter engine with a regenerated engine, several design requirements must first be satisfied. Firstly, the gross weight of the reference simple cycle must be maintained, secondly the Take-off
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ght (
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Heat Exchanger Effectiveness ε %
REF TW REG TW (TW=EW+MFB)
Ref EW REG EW (EW= Engine Weight)
REG MFB (MFB = Mission Fuel Burn) REF MFB
power (TO) must be matched and finally the size, volume and fuselage aerodynamic profile of the existing helicopter must also be maintained. The scope of this study covers and satisfies the first two requirements and assumes for both helicopters, that the installation of the HE will have minimal or neglected effects on the helicopter size, volume and fuselage aerodynamic profile and therefore will not affect its performance. 8.3 Break-even Point The point at which the HE “added weight” is exactly compensated by reduction in mission fuel burn, represents an equilibrium or break-even point. In other words, in order for a regenerated helicopter to be economical and beneficial, the fuel carrying capacity of the regenerative rotorcraft must be reduced by an amount equal to the weight added by its installed HE. Once the break-even point is satisfied for a given heat exchanger, any additional fuel reduction can be regarded as reduction in All-UP-Mass
(AUM) denoted as DELTA AUM (ΔAUM).
8.4 All-Up-Mass
For the purpose of this study a positive ΔAUM demonstrates reduction in AUM of the regenerated helicopter compared to its corresponding reference simple cycle helicopter and a negative ΔAUM demonstrates increase in AUM compared to reference simple cycle helicopter (a negative ΔAUM can also be classed as “weight penalty”). The missions that justify the need for regeneration demonstrate positive ΔAUM and missions with negative ΔAUM are not found feasible for regeneration, investigated in this study respectively.
Depending on the evaluation criteria, an obtained positive ΔAUM can either be utilized as mission fuel saving, enabling the helicopter to increase its mission range, or it can be utilized as an increase in useful payload. It is however also possible to utilize the obtained positive “ΔAUM” to offer benefits in both aforementioned areas. 8.5 TWIN ENGINE LIGHT, MISSIONS RESULTS
8.5.1 TEL, Reference Mission A & B, Results
The results for TEL, Reference simple cycle and regenerated helicopter for reference mission A and B are presented in Table 6 and Table 7. Clearly, significant reduction in fuel burn is achieved throughout the assumed range of HE effectiveness. Approximately, 25% to 30% mission fuel burn reduction at HE effectiveness of 40% and up to about 40% to 44% at 80% HE effectiveness, with respect to both missions (A & B). Despite the significant mission fuel burn reduction, a negative ΔAUM is found for
HE Effectiveness of up to 80% with respect to both missions (shown in figure 12 & 13). This is purely due to the fact that the low range 19 nautical miles of the Passenger Air Taxi mission and the 26 nautical miles of the Emergency Medical Service mission is not sufficient to allow and meet the required level of fuel reduction to
achieve a break-even point and essentially results in a
positive ΔAUM.
Table 6: TEL, Reference Simple Cycle and
Regenerated Helicopter, Reference Mission A, Passenger Mission Results
TEL , Reference Mission A Results
Power Plant Cycle Fuel Burn (kg)
Reference SC 57.68
Regenerated ε 40% 44.36
Regenerated ε 50% 41.2
Regenerated ε 60% 37.81
Regenerated ε 70% 34.9
Regenerated ε 80% 32.14
Figure 11: HE Optimization for TEL, Regenerated Helicopter Reference Mission A, Passenger Mission
Figure 12: TEL Regenerated Helicopter, Reference
Mission A, Passenger Mission, ΔAUM
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66.0
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ght (
kg)
Heat Exchanger Effectiveness
HE Weight Added
Fuel Reduction
DELTA AUM
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eigh
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Optimum HE Effectiveness
REF EW REG EW
REF MFB REG MFB
REG TW
REF TW TW=EW+MFB MFB= Mission Fuel Burn
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66
117
85.9
110.9
133.7
153.9
170.8
66.4
85.4 93.2
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ε 40% ε 50% ε 60% ε 70% ε 80%
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HE Weight Added
Fuel Reduction
DELTA AUM
Optimum HE Effectiveness
8.5.2 TEL, Reference Mission D, Results
Compared to low range missions A and B, the 200 nautical mile SAR mission proves to have much positive and favorable results for regeneration potential, presented in Table 8. The long range and the favorable operation of the rotorcraft both justify the need for regeneration. A positive ΔAUM is achieved throughout the entire range of assumed
HE effectiveness (40% - 80%) Figure 14. However, the
amount of “Positive ΔAUM” diminishes above 60% HE
effectiveness, this is due to the fact that weight added by the heat exchanger increases significantly above 60% HE effectiveness. Also, the fixed mission range 200 nautical miles limits the fuel savings. Therefore the optimum HE effectiveness for his particular mission is 60% with a weight saving of 93kg, as shown in Figure 15.
Table 7: TEL Reference Simple Cycle and Regenerated Helicopter, Reference Mission B, EMS Mission Results
TEL RC, Reference Mission B Results
Power Plant Cycle Fuel Burn (kg)
Reference SC 60.19
Regenerated ε 40% 47.75
Regenerated ε 50% 43.2
Regenerated ε 60% 40.18
Regenerated ε 70% 37.19
Regenerated ε 80% 34.53
Figure 13: TEL Regenerated Rotorcraft, Reference
Mission B, EMS, ΔAUM
Figure 14: HE Optimization for TEL, Reference
Mission D, SAR Mission
Figure 15: TEL Regenerated Rotorcraft, Reference Mission D, SAR, ΔAUM
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117.0
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-43.0
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-20
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100
120
140
ε 40% ε 50% ε 60% ε 70% ε 80%
Wei
ght (
kg)
Heat Exchanger Effectiveness
HE Weight Added
Fuel Reduction
DELTA AUM
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REG MFB REF MFB
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REF EW REG EW
EW=Engine Weight MFB = Mission Fuel Burn
Optimum HE Effectiveness
Break- even Point
Table 8: TEL, Reference Simple Cycle and Regenerated Helicopter, Reference Mission D, SAR Mission Results
TEL RC, Reference Mission A Results
Power Plant Cycle Fuel Burn
(kg)
Reference SC 401.2
Regenerated ε 40% 315.28
Regenerated ε 50% 290.35
Regenerated ε 60% 267.51
Regenerated ε 70% 247.27
Regenerated ε 80% 230.43
8.6 TWIN ENGINE MEDIUM, MISSIONS RESULTS
8.6.1 TEM, Reference Mission C The specific weight of the HE is function of engine mass flow. The Higher the mass flow through the engine, the higher the matrix weight of the heat exchanger. The specific Engine A1 of the TEM helicopter (PUMA SA330 Model) with the engine mass flow of 5.9kg enormously increases the gross weight of the HE above 60% HE
effectiveness (See Fig.7). Nevertheless, a positive ΔAUM
is achieved for the HE effectiveness of up to 65% and the break-even point is occurs at HE effectiveness of around 65% to 70% (See Fig.16). Due to higher added weight of heat exchanger above 60% HE effectiveness, the positive
ΔAUM gradually diminishes until the break-even point and
turns negative above 70% HE effectiveness (resulting in weight penalty). The optimum effectiveness with maximum positive ΔAUM for this particular mission is achieved at
50% HE effectiveness with a total weight gain of 59kg (See Figure.17). (Figure 22 shows variations in the SFC as a function of mission time (Reference Mission C) with respect to different flight conditions for reference simple cycle and for regenerated helicopter configuration).
8.6.2 TEM, REFERENCE MISSION D A positive ΔAUM is achieved throughout the range of heat
exchanger effectiveness of (40% - 80%). Figure 18 shows the (REG TW dashed line) lies underneath the (REF TW) and further drops until it meets the optimum efficiency point at 60%. Due to the significant increase in HE weight above
60% HE effectiveness, the (ΔAUM) gradually starts to
diminish beyond 60% HE effectiveness. Despite the
enormous weight added by HE, a positive ΔAUM is
achieved up to 80% HE effectiveness. Maximum ΔAUM of
231kg is achieved at optimum 60% effectiveness, highlighted in Figure 19.
Figure 16: HE optimization for TEM, Reference Mission C, Utility Oil & Gas Mission
Figure 17: TEM, Regenerated Rotorcraft, Reference Mission C, Utility Oil & Gas Mission, ΔAUM
77 100
159
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460
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500
ε 40% ε 50% ε 60% ε 70% ε 80%
Wei
ght (
kg)
Heat Exchanger Effectiveness
HE Weight Added
Fuel Reduction
DELTA AUM
Optimum HE effectiveness
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ght (
kg)
Heat Exchanger Effectiveness ε %
REF EW
REG EW
REF MFB REG MFB MFB = Missoin Fuel Burn
EW = Engine Weight
REFTW REG TW
TW=EW+MFB Optimun HE Effectiveness
0
0.005
0.01
0.015
0.02
0.025
0.03
0.035
0 200 400 600 800 1000 1200 1400 1600 1800
Fuel
Flo
w (k
g/s)
Mission Time (sec)
REF SC REG e40% REG e50%REG e60% REG e70% REG e80%
Idle
Cruise Cruise
Cruise
Idle Idle
Figure 18: HE optimization for TEM, Reference Mission D, SAR Mission
Figure 19: TEM Regenerated Rotorcraft, Reference Mission D, SAR Mission, ΔAUM
Figure 20: TEL, Reference Simple Cycle and
Regenerated Helicopter, Reference Mission A, Fuel Flow Comparison
Figure 21: TEL, Reference Simple Cycle and Regenerated Helicopter, Reference Mission B, Shaft
power Comparison
Figure 22: TEM, Reference Simple Cycle and Regenerated Helicopter, Reference Mission C, SFC
Comparison
25000
65000
105000
145000
185000
225000
265000
305000
345000
-50 100 250 400 550 700 850 1000 1150 1300 1450 1600
Shaf
tpow
er (K
W)
Mission Time (Sec)
REF SCREG e80%
Idle
Cruise Cruise
Idle
Cruise
Idle
50
80
110
140
170
200
230
260
290
0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000
SFC
µg/J
Mission Time (Sec)
REF SCREG e40%REG e50%REG e60%REG e70%REG e80%
Idle Idle
Cuise Cruise
77 100
159
260
460
223
307
390
456
513
146
206 231
196
53
0
50
100
150
200
250
300
350
400
450
500
550
ε 40% ε 50% ε 60% ε 70% ε 80%
Wei
ght k
g
Heat Exchanger Effectiveness
HE Weight Added
Fuel Reduction
DELTA AUM
Optimum HE effectiveness
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10
9. CONCLUSIONS A well-known and effective approach of regeneration is deployed to enhance the integrated performance of the current simple cycle turboshaft helicopter. The methodology is based on incorporating a heat exchanger, which enables the heat transfer between the exhaust gas and compressor delivery air to the combustion chamber. The proposed methodology is implemented by adopting an integrated framework capable of computing the flight mechanics and engine performance of any defined rotorcraft configuration within any designated mission. A Heat exchanger trade-off study has been developed to identify optimum heat exchanger design effectiveness that represents maximum fuel savings applicable to each mission for both helicopter configurations (investigated).
The overall methodology has been applied to four different generic–baseline rotorcraft missions. The results obtained so far indicate that the proposed method is promising in achieving substantial reduction in mission fuel burn for all generic missions investigated with respect to both helicopter configurations. The obtained results also suggest that despite substantial reduction in mission fuel burn, the application of regeneration is not necessarily found suitable for certain types of helicopter missions, reported in this study. Whereas, it can be considered as a promising concept for helicopter missions with long range e.g. Utility Oil & Gas and Search and Rescue missions.
NOMENCLATURE
ACARE Advisory Council for Aviation Research and Innovation in Europe
AUM All Up Mass ∆AUM DELTA All Up Mass DP Design Point EMS Emergency Medical Service FPT Free Power Turbine HE Heat Exchanger HECTOR HeliCopTer-Omni-disciplinary-Research-
platform HPC High Pressure Compressor LPC Low Pressure Compressor MTOW Maximum Takeoff Weight MFB Mission Fuel Burn OPR Overall Pressure Ratio OEW Operational Empty Weight RC Rotorcraft REF TW Reference Total Weight REG TW Regenerated Total Weight REF EW Reference Engine Weight REG EW Regenerated Engine Weight REF FB Reference Fuel Burn REG FB Regenerated Fuel Burn TEM Specific Fuel Consumption SC Simple Cycle (Brayton Cycle) SAR Search And Rescue TET Turbine Entry Temperature TEL Twin Engine Light
TEM Twin Engine Heavy Wa Total Engine Mass flow
REFERENCES 1. Advisory Council for Aeronautics Research in Europe
home page http//www.acare4europe.com 2. Colin F. McDonald, Aristide F. Massardo, Colin
Rodgers, Aubrey Stone, (2008) "Regenerated gas turbine aero-engines. Part I: early development activities, Aircraft Engineering and Aerospace Technology, Vol. 80 Issue: 2, pp.139 – 157
3. Colin F. McDonald, Aristide F. Massardo, Colin Rodgers, Aubrey Stone, (2008) "Regenerated gas turbine aero-engines. Part II: engine design studies following early development testing", Aircraft Engineering and Aerospace Technology, Vol. 80 Issue: 3, pp.280 – 294
4. Colin F. McDonald, Aristide F. Massardo, Colin Rodgers, Aubrey Stone, (2008) "Regenerated gas turbine aero-engines. Part III: engine concepts for reduced emissions, lower fuel consumption, and noise abatement", Aircraft Engineering and Aerospace Technology, Vol. 80 Issue: 4, pp.408 – 426
5. Edward J. Privoznik. Allison T63 Regenerative Program. Annual forum proceedings. Presented at the 24
th Annual Forum of the American Helicopter Society,
May 1968. 6. Ioannis Goulos, Panos Giannakakis, Vassilios
Pachidis, Pericles Pilidis. Mission Performance Simulation of Integrated Helicopter – Engine Systems using an Aeroelastic Rotor Model. Proceedings of ASME Turbo Expo June 2013
7. Jane’s International Aero-engines 8. Nicolas C. Kailos (1967): Increased helicopter
capability through advanced power plant technology Journal of the American Helicopter Society, Volume 12, Number 3, 1 July ,pp. 1-15(15)
9. Grieb, H and Klussman. W, (May 1981), Regenerative helicopter engine- Advances in performance and expected development problems. AGARD Conference Proceedings, helicopter propulsion systems. Volume No 302.
10. Goulos, I., Hempert, F., Sethi, V., Pachidis, V., DIppolito, R., and D Auria, M.,“Rotorcraft Engine Cycle Optimisation at Mission Level”, Proceedings of ASME Turbo Expo 2013, San Antonio, Texas, No. GT2013-94798, June 3-7 2013.
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