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† 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)

Copyright © 2013 by ASME


3

60

80

100

120

140

160

180

200

220

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



4

60

80

100

120

140

160

180

200

220

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 REC = Regenerated e = HE effectiveness

05

10152025303540

40 45 50 55 60 65 70 75 80

Hea

t Exc

hang

er S

peci

fic

Wei

ght .

lbs

/ Wa

Heat Exchanger Effectiveness ε %

Heat Exchanger Specific Weight LOWER LIMITHeat Exchanger Specific Weight UPPER LIMIT

150

165

180

195

210

225

240

255

270

40 45 50 55 60 65 70 75 80 85

Wei

ght F

or T

wo

Reg

ener

ated

En

gine

s B

1 (k

g)


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).



5

500550600650700750800850900950

1000

40 45 50 55 60 65 70 75 80 85

Wei

ght F

or T

wo

Reg

ener

ated

Eng

ine

B2

(kg)


Gross Weight Lower LimitGross Weight Upper Llimit

Engine Mass Flow 5.9kg

010203040506070

075

150225300375450525

0 300 600 900 1200 1500 1800Mission Time (Sec)

Cru

ise

spee

d (m

/sec

)

Alti

tude

(m)

Altitude V_AirCruise speed

010203040506070

0150300450600750900

1050

0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000

Cru

ise

Spee

d (m

/sec

)

Alti

tude

(m)

Mission Time (sec)

Altitude Vel_AirCruise speed

0

10

20

30

40

50

60

70

0

75

150

225

300

375

450

525

0 1250 2500 3750 5000 6250 7500 8750

Cru

ise

Spee

d (m

/sec

)

Alti

tude

(m)

Mission Time (sec)

Altitude Vel_AirCruise speed

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



6

0

50

100

150

200

250

300

40 45 50 55 60 65 70 75 80

Wei

ght (

kg)


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





Figure 11: HE Optimization for TEL, Regenerated Helicopter Reference Mission A, Passenger Mission

Figure 12: TEL Regenerated Helicopter, Reference

Mission A, Passenger Mission, ΔAUM

19.5 25.5 40.5

66.0

117.0

13.3 16.5 19.9 22.8 25.5

-6.2 -9.0 -20.6

-43.2

-91.5

-125

-100

-75

-50

-25

0

25

50

75

100

125

150

ε 40% ε 50% ε 60% ε 70% ε 80% Wei

ght (

kg)

Heat Exchanger Effectiveness

HE Weight Added

Fuel Reduction

DELTA AUM



7

0

50

100

150

200

250

300

350

400

450

500

550

600

650

700

40 45 50 55 60 65 70 75 80W

eigh

t (kg

) Heat Exchanger Effectiveness ε %

Optimum HE Effectiveness

REF EW REG EW

REF MFB REG MFB

REG TW

REF TW TW=EW+MFB MFB= Mission Fuel Burn

19.5 25.5

40.5

66

117

85.9

110.9

133.7

153.9

170.8

66.4

85.4 93.2

87.9

53.8

0

20

40

60

80

100

120

140

160

180

ε 40% ε 50% ε 60% ε 70% ε 80%

Wei

ght (

kg)


HE Weight Added

Fuel Reduction

DELTA AUM


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






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

19.5 25.5

40.5

66.0

117.0

12.4 16.7 20.0 23.0 25.7

-7.1 -8.8

-20.5

-43.0

-91.3 -100

-80

-60

-40

-20

0

20

40

60

80

100

120

140

ε 40% ε 50% ε 60% ε 70% ε 80%

Wei

ght (

kg)


HE Weight Added

Fuel Reduction

DELTA AUM



8

300

375

450

525

600

675

750

825

900

975

1050

1125

1200

1275

1350

40 45 50 55 60 65 70 75 80

Wei

ght k

g


REG MFB REF MFB

REG TW REF TW TW= EW+MFB

REF EW REG EW

EW=Engine Weight MFB = Mission Fuel Burn


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






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

260

460

115

159

201

236

265

39 59

42

-24

-195 -250

-200

-150

-100

-50

0

50

100

150

200

250

300

350

400

450

500

ε 40% ε 50% ε 60% ε 70% ε 80%

Wei

ght (

kg)


HE Weight Added

Fuel Reduction

DELTA AUM

Optimum HE effectiveness



9

300

400

500

600

700

800

900

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

40 45 50 55 60 65 70 75 80

Wei

ght (

kg)


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

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53

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ε 40% ε 50% ε 60% ε 70% ε 80%

Wei

ght k

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HE Weight Added

Fuel Reduction

DELTA AUM

Optimum HE effectiveness



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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