NASA Contractor Report 185222 

PW FR-20668 

Assessment of High Temperature 

Superconducting (H‘TS) Electric 

Motors for Rotorcraft Propulsion 

.Jay Doernhach 

C’n it cd Tech rio log ies Co r’poru t ion 

Prott CG I.i’liitticv 

It‘est Pnlm Bcwch, Noritlo 

April 1990 

Prepared for 

Lewis Research Center 

Under Contract NAS3-25 1 17 (‘Task 009) 

Narional Aeronautics and 

Soace Aaministratron 
NASA Contractor Report 185222 

PW FR-20668 

Assessment of High Temperature 

Superconducting (HTS) Electric 

Motors for Rotorcraft Propulsion 

Jay Doembach 

United Technologies Corporation 

Pmtt & Whitney 

?Vest Palm Beach, Florida 

April 1990 

Prepared for 

Lewis Research Center 

Under Contract NAS3-25117 (Task 009) 

National Aeronautics and 

Soace Administration 
Table of Contents 

ABSTRACT .................................................................. 1 

PREFACE ..................... ...... 2 

1.0-Introduction .............................................................. 3 

2.0 . Technical Requirements ..................................................... 5 

3.0 . Summary of Results ........................................................ 6 

4.0 - Design Criteria ........................................................... 12 

5.0 - HTS MotorlGenerator Concepts .............................................. 14 

6.0 . Closed Loop Cooling Concepts .............................................. 18 

7.0 - Expendable Refrigerant Concepts ............................................. 21 

8.0 . Room Temperature HTS Systems ............................................. 25 

9.0 . Mission Studies .......................................................... 28 

10.0 . Weight. Power. Performance and Cooling Audits ................................ 33 

11.0 . Conclusions and Recommendations .......................................... 41 

12.0 . List of Abbreviations ...................................................... 44 

13.0-References ............................................................. 45 

'l'ahle of Contents PRECEDING PAGE BLANK NOT FILMED ... 111 
FR-20668 

ABSTRACT 

The successful development of high temperature superconductors (HTS) could have a 

major impact on future aeronautical propulsion and aeronautical flight vehicle systems. 

A prelimiiiary examination of the potential application of HTS for aeronautics indicates 

that significant benefits may be realized through the development and implementation 

of these newly discovered materials. 

Applications of high temperature superconductors have been envisioned for ceveral 

classes of aeronautical systems, including subsonic and supersonic transports. 

hypersonic aircraft, V/STOL aircraft, rotorcraft, and solar, micirowave and laser powered aircraft. This study examines the potential application and benefits of applying 

HTS to helicopter drive trains. 

Three existing helicopters were chosen to determine the feasibi1it:y of using HTS motorq 

and generators to replace their conventional mechanical power drive systems. Helicopter selection was made to cover the range of applications from mall to very large to be 

sure all size categories would be considered including the 500 antl 3000 HP applications 

specified in the contract requirements. Motorigenerator designs were created antl cooling systems to keep them at cryogenic temperatures were sized. The helicopter’s gas 

turbine power turbines were customized to drive the electrical generators at acceptable 

speeds while the existing helicopter rotor speeds were used as the design speed requirements for the electrical motors. Mission studies using the chosen helicopters were then 

done using a helicopter sizing computer program (HESCOMP) developed for NASA by 

Boeing Vertol Co. HESCOMP was used to determine the impact of the HTS weight and 

power changes on helicopter takeoff gross weight (TOGW). The TOGW impacts 

showed which applications would be good choices for HTS motors and genera tors to 

replace the mechanical transmissions. 

I 
FR-20668 

PREFACE 

This report documents a contracted study to investigate conceptual HTS power train 

systems for a helicopter with a Technology Availability Date (TAD) year of 2000. 

System weights as well as 'TOGWs were generated for helicopters with gas turbincs rated 

in the 500 through 13000 HP range having missions designcd using the HESCOMP 

computer code. This study was conducted by Pratt 22 Whitney for NASA during the 

time period from February 1989 through July 1989. 

The NASA project manager for this study was George Turney, NASA Lewis Research 

Ccnter, Cleveland, Ohio. Roger Luidens of NASA Lewis was the technical consultant. 

George Champagne was program manager at Pratt Rr Whitney. Kcy technical ctmtributions came from Osman Mawarcli, President of Collaborative Planners Inc. in 

Cleveland, Ohio; Frank Biancartli of United Technologies Rcsearch Center in Ea5t 

Hartford, Connccticut; Marv Glickstcin, Bill Grunske and Bruce Trembly of Pratt & 

Whitney at West Palm Beach, Florida. Jay Doernbach of Pratt & Whitney at Wcst 

Palm Beach coordinated the efforts, proviclctl technical contributions and wrote the final 

report. 

2 
FR-20668 

1.0 = Introduction 

The recent discovery of high temperature superconductors (HIS) has creatcd a major 

stir of excitement in the qcientific community. Thc potential hencfits which may bc 

provided through application of these newly discovcrcd superconducting materials arc 

now being explored by investigators. Studies have bcen initiated to aqsess the potcn tial 

advantage of HTS in a number of areas including clectrical powcr transmiqsion and 

storage, communications, computer systems, mcdical diagnostic equipment, and ground 

transportation systems (magnetically levitated high speed trains). 

Despite the claims and projections made for thesc newly discovered superconducting 

materials, it is gcnerally agreed that a significant effort will be required before thcse 

materials can be devcloped to a state whcrc they can be succes*;fuIly used in practical 

applications. Nevertheless, the feeling of many rescarchcrs seems to be that most or all 

of the currently recognized problems with HTS can be rcsolvcd through applied rewrch 

efforts. Based on this premisc, HTS will, in time, find it's way into applications of aeron aut ical sys tems . 

With successful demonstration of superconductivity at a temperature of 95K, which is 

in excess of the liquid nitrogen boiling point (77K), potential applications for 

aeropropulsion syqtems appear feasible. As the technology for HTS develops, cxtending 

the critical temperatures to higher levels, thc potential bcnefits of such applications appear more interesting. With evidence of potential superconducting material exhibiting 

critical temperatures approaching room temperature (300K), a variety of potcntial 

aerosystem applications can be envisioned, one of which is helicoptcr drive system motors and generators to replace the mechanical drive train. The ohjective of this task was 

to define the characteristics of high temperature superconducting motors and generators 

and to asscss their bencfits in helicopter main lift systcms. This study looked at thrce 

existing hclicopter systems; namely, the 5mnll Bell .let Ranger, thc medium Sikorsky 

Black Hawk and the largc Sikorsky Super Stallion. 

Collaborative Planners,Inc of Cleveland Hcigh t5,Ohio tlitl the HTS electric 

motor/generator design work and specified thc cooling requircrncn ts for all thrce sizes 

of applications investigated. 

1 .O - Introrluction 
FR-20668 

Two electrical motor constructions, callcd SROC and SRASC, arc mcntioncd frequcn tly 

in this study and somc explanation here will bc hclpful. Thc superconducting rotor only 

construction (SROC) is uscd for the motors and gencra tors having superconductors only 

in the rotors with the resulting electrical efficicncies of ncar 99%. The superconducting 

rotor and stator construction (SRASC) is uscd for the motors and generators haVing 

superconductors in both the rotors and stators with the resulting clcctrical efficicncics 

greater than 99%. 

United Technologies Research Ccntcr of East Hartford, Connecticut sized thc 

cryocoolers needed to meet the cooling requirements 5pccificd for various sizes of applications investigatcd. 

Pratt & Whitney of Wcst Palm Beach performed the helicopter tradc factor studics using 

the HESCOMP code developed by Boeing Vertol Co. Pratt & Whitney of West Palm 

Bcxh, sized the expendable cooling equipment needcd to meet the cooling rcquircmcn ts 

specified for the various sizcs of applications investigatcd, and also coordinated the entire study and compiled and wrote the final report. 

I .O - Introduction 4 
2.0 - Technical Requirements 

The main technical requircmen t of this contract was to investiga tc HTS motor/gcncrator 

characteristics consisting of power output, speed, size, efficiency and system weight and 

to apply these to the main lift system of helicopters designed with TAD’S in the year 

2000. The HTS operating temperatures of 95K and 300K were to be used for 500 & 

3000 HP applications with an individual component efficiency of at least 97 to 98%. 

The cryogenic cooling systcms to be conyidercd were closed loop refrigeration and cxpendable liquid nitrogen systems. Cooling systems are necessary to keep the superconductors at operating tempcratures consistent with the electrical wperconductivity which 

makes the motors and generators much smallcr than prescnt statc of the art equipment 

and thus provides a weight benefit for flight weight equipment. 

Weight and performance characteristics of conventional and HTS drive systems werc to 

be compared for each application. Missions and airframe performance and weight 

modcls were to be developed to assess the HTS and conventional systems’ performances. 

2.0 - Technical Requirements 
~ 

FR-20668 

3.0 - Summary of Results 

This study showed that larger helicopters provide thc best potential for application 

of the HTS motorjgenerator drive systems. Figure 3-1 is a schematic of the mechanical 

drive system that is used to powcr the large MH-53E helicoptcr. The mechanical drive 

systcm shown in figure 3-1 is very similar to that of the CH-S3E helicopter which wnq 

one of the systems investigated in this study. 

Figlire 3-1. Mechanical Power Train for 7'11~ MIi-53E Ilelicopter 

3.0 - Sirmmary of Results 
FR-20668 

Figure 3-2 again shows the same MH-53 drive train with modifications to include the 

HTS motor and generator for main rotor power. The three gas turbines have been rcplaced with two having the power of three and the electrical gencr-ator is attached to the 

powcr turbine shaft of each. The large three engine gearbox has been eliminated but an 

accessory/tail rotor gearbox is still needed. Thc motor, generator and cryocoolcr shown 

arc only conceptual. 

MAIN ROTOR 

MOTORIGENERATOR (aM&] 

TRANSFER CONDUIT - 

Figure 3-2. Conceptual II’I’S Powcr rrain for the C11-53E f Iclicopter. 

Table 3-1 includes the kcy study item5 conqisting of cooling loads. system weight<. 

power requirements and, electrical cfficiencies used in this study and providcs the system 

comparisons including drive train specific power (HP/LB) and aircraft takeoff grow 

weight (TOGW) changes. 

3.0 - Srirnrnnry of Results 
FR-20668 

KEY ITEMS 

HTS ELECTRlCAL 

SYSTEM SUMMARY 

Cooling Loads 

Refrigeration 

1 

I Max Load (K W) 

SizelPower (HP) 

Motor 

Generator 

Speeds (RPM) 

Motor 

Generator 

Efficiencies (?A) 

Motor 

Generator 

SYSTEM WEIGHTS 

Power Drive Train 

Main Rotor (LBS.) 

1. Mechanical Sys. 

A. Main Rotor 

2. HTS Systems : 

A. Total CILoop 

B. Total Expend. 

C. Total Air cool 

PERFORMANCE 

Specific Power 

(HPILB) 

1. Mechanical Sys. 

2. HTS Systems 

A. ClLoop System 

8. Expnd LN2 Sys 

C. Air Cooled Sys 

TOGW CHANGE-) 

Due to Changes in 

Weight and Power 

A. CILoop System 

8. Expnd LN2 Sys. 

C. Air Cooled Sys. 

Table 3-1 - HTS Key /terns Summary 

Helicopter Power Plant : Gas TiirhinP Maximrim HP Rating at Sea Level 

502 HP 3606 HP 13407 HP 502 HP 3606 HP 13407 HP 

SROC MotartGenerator SRASC MotorlGener-ator 

0.474 

318 

318 

350 

2 1500 

99.9 

99.9 

182 

499 

380 

309 

1.747 

0.637 

0.887 

1.029 

16.5 

7.1 

4.5 

3.229 

2165 

2165 

25 7 

17900 

99.9 

99.9 

1387 

1815 

1230 

943 

1.561 

1.193 

1.917 

2.296 

5.2 

-5.1 

-7.4 

11.77 

7894 

7894 

179 

11900 

99.9 

99.9 

573 1 

5380 

3 798 

2908 

1377 

1467 

2.290 

2.715 

-2.1 

-14.1 

- 17.0 

0.104 0.597 2.124 

318 2165 7894 

318 2165 7894 

350 257 179 

2 1500 17900 11900 

99.98 99.98 99.98 

99.98 99.98 99.98 

182 1387 5731 

312 990 2950 

35 1 912 2626 

325 839 2427 

1.747 1.561 1.377 

1.019 2.187 2.676 

0.919 2.427 3.084 

0.978 2.580 3.253 

5.4 -6.5 -16.5 

6.7 -8.1 - 18.9 

5.6 -8.7 -1.9.6 

3.0 - Summary of Rrsiilts 
FR-20668 

Figure 3-3 shows that the helicopter main rotor drive weights for convcnlional mechanical drives are lighter than any of the HTS candidates for helicopters requiring less 

than 501) Horscpower. The crossover point for larger drives can be seen to depend 

strongly on the electrical motor/generator construction,the cooling load,and the type of 

cooling devices used. 

a 

a 

-I 

LL 

0 

a 

0 

z 

a 

a 

3 

0 

r 

3 

z 

a 

t 

m 

W 

L 

a 

a 

a 

5 

a 

I 

0 

e 

0 

I I 1 I I I I 

0 2 4 6 8 10 12 14 

HELICOPTER HORSEPOWER IN THOUBANDS 

, 

HT8 SYSTEM8 USED TWO EFFICIENCY LEVELS 

Figure 3-3. 1 lelicopter Main Rotor Drive System I17 S Replacetrirnt of blain Rotor hive. 

3.0 - Summary of Results 9 
FR-20668 

Figure 3-4 is similar to 3-3 exccpt that the main rotor drive systcm weight is plotted 

against horsepower transmitted through thc main rotor rather than the helicoptcr total 

ratcd horscpower. Figure 3-4 shows that under about 500 horscpowcr, the mcchanical 

drives arc lighter and abovc about 5500 horscpowcr the HTS Yyqterns are lightcr. i 

........................... 

............................................................................................. 

0 2 4 6 

MAIN ROTOR HORSEPOWER (ThOU8andd 

8 

FIGURE 3-4 

* SROCLN? - SRASCLN2 -+ SROC AIR/C -+- SROCCRYOCOOLE 

+ 8RA8C AIR/C MECH.DRIVE -+ SRASC CRYOCOOLER 

I 

, HT9 8Y8TEMS USED TWO EFFICIENCY LEVELS 

Figure 3-4. I lelicopter hlain IColor Drive System 11'1s Replaccnirnl of blain Rotor Ilrive. 

3.0 - Summary of Results 10 
FR-20668 

3- 

2.6 

2- 

1.6 

1- 

0.6 

0'- 

Figure 3-5 shows the qpccific power tranqmitted to thc main rotor vs helicoptcr total 

ra tcd horscpower. 

The 5tutly is summarizcd using two level< of electrical generating efficiencies: one of 

99.9?'0 (not including cooling rcquiremcntq) and a qecond level of 99.99 + "1'. This much 

higlicr efficiency according to CPI ic thoright to be po5sible by the year 2000. 

- 

- 

- .. 

I 1 I I I I I 

3,6 I I 

PIQURE 3-6 

-e- SROC LN2 --*r 9ROC CRYO/W 9AOC AIR/C --yC MECH.DRV. 

-%- SRASC LNP -&- SRASC cRYO/e 8RASC AIR/C I 

I 

/I HTS $Y$TEM8 USED TWO EfFlClENW LEVELS 

, 

Figure 3-5. I lelicopter Main Rotor Drivc System Pnwcr 'l'mnsniittrd to Main Rotor Ihive. 

3.0 - Summary of Results 
f R-20668 

! 

I 

i 4.0 - Design Criteria 

Electric motor and generator design speeds: 

This study was begun hy making a helicopter power survey, an excerpt of which ic 

shown in table 4-1. References 2 and 6 were used to survey the gcncral power cpxtr'urn 

of present day helicopters. This survey yielded thc three helicopter choices uscd in thiy 

study and satisfied the study intent to get rotorcrart in the SO0 and 3000 horwpowcr 

categoriec. Main rotor speeds needed for sizing the electric motorc were also necdcd arid 

these were obtained initially from various sourccs including conversations with the 

manufacturers. Finally, these data were updated where neccswry using data from rcferences 3, 4, and 5. 

Table 4-1 -Helicopter Specifications 

(From Aviation Week 8; Space Technoloav March 24, 7988 and March 9, 79871 

Popular Rotor Empty Gross Maximum Powerplant Total 

Name Diameter Weight Weight Speed Gas Turbine Horsepower 

(Ft.) (Lbs.) (Lbs.) (MPh) Engines (HP) 

Bell Jet 33.3 1635 3200 140 1 .Allison 420 

Ranger 2.50-C2OJ 

206 B -JR3 

S ikorsk y 53.7 10900 16450 184 2 GE 

Black Hawk 71700-GE- 700 3120 

UH-GOA 

Sikorsky 79.0 33226 69750 196 3 GE 13140 

Stallion 

Super T64-4 16 

CH-53E 

4.0 - Design Criteria 12 
FR-2 0 6 6 8 

The electric gencrator was to be drivcn by the power turbine of the gas gencrator. 

To makc thc speeds compatible with the electrical gcncrator’s tlc\ircd design speed<, t he 

powr turbines werc customizcd for thesc applications. Powcr turhinc spceds for cach 

power catcgory wcrc calculated by Pratt & Whitncy to be fcasihlc and compatible with 

thc electrical gencrator speed of 20000 RPM or less, :is rcqucstcd by CPI. To kccp thcw 

speeds down, all dcsign speeds were calculated basccl on threc stage powcr turbines (assumed to rcquire no weight addition) with the resulting spcccls shown in Table 4-2. 

Table 4-2 - Design Ground Rule Summary 

APPLICATION Motor Generator Generator 

1 Turbine 2 Turbine 

Engine Engines 

Speed Speed Speed 

(RPW WM) (RPW 

500 HP 350 21500 Probably too 

Small for two 

3000 HP 257 17900 I9 700 

13140 HP I 79 11900 15000 

Drive system power utilization: 

The HESCOMP program providcd the existing hclicoptcr power distributions used in 

thi.; study which are shown in Tablc 9-2. The hclicoptcr accc.;sory drivc rcquircmcnts 

are the same for both the HTS and convcntional systcrns and so werc not adjustctl. 

Power extraction necdcd for the cryocoolers was .;pccified by UTRC and these value.; 

~vverc scaled for the actual equipment sizcs. CPI specified that the SRASC machincry 

requires a cooling fan for air circulation and this power uwgc is also accountctl for. 

Since it was assumccl that supcrconducting cablcs/conciiritc wcrc to be uscd, the electrical power transmission los~cs wcrc conyidcred to I~c negligible as was suggested by 

CPI. A 3:‘n loss of power transmitted was uwl for thc convcntion:1I mechanical po~cr 

transmissions due to hearing sen1 and gear train friction a.; included in the HESCOMP 

code. A kcy item power audit tlerivctl from this .;tudy is shown in Tahlc 10-2. 

System weights: 

The helicopter mechanical drive systcm wights were obtained from references 3, 4, 

and 5 mentioned previously. Thc wcights for the main rotor c1riL.e systems are the ones 

u.;cd in this study since the clcctric motors and generators werc uscti to replace the niairi 

drivc5 in the HTS versions of these helicopters. A key items wight audit i5 shown in 

Table 10-1. Tlic same references were alw uwd to obtain the TOGW for the threc helicopters in this study. 

4.0 - Ihign Critcria 
FR-20668 

5.0 - HTS MotorGenerator Concepts 

I 

Table 5-1. shows the geometries and wcightq for key itcms in thc HTS elcctrical qystcm 

along with the cooling systcrn rcquirerncnts uscd for both rnotor/generator efficicncics 

for three lcvels of power transmission. 

I 

I 

Table 5-1 - MotorlGenerator Sizing from CpI 

Motor Dimensions 

Stator Diameter (In.) 

Rotor Diameter (In.) 

Length (In .) 

Weight (Lbs.) 

Generator Dimensions 

Stator Diameter (In.) 

Rotor Diameter (In.) 

Weight (Lbs.) Length (In .) 

MotorlGenerator 

Total Weight 

SROC (Lbs.) 

SRASC (Lbs.) 

Cooling Load SRASC 

Requirements 

A.Heat Transfer (WATTS) 

B.Ohmic Losses (WATTS) 

C.Electromagnetic (In1 9 11s) 

Heat Losses 

(Use Fan Blower) 

500 HP 

21 

14 

14 

224 

7 

3.5 

14 

155 

3 79 

303 

23 

130 

746 

3000 HP 

49.79 

38.3 

15.32 

62 7 

8.02 

4.01 

20.05 

62 7 

1254 

1003 

54 

776 

746 

13140 HP 

39.78 

30.6 

12.24 

1871 

10.28 

5.14 

25.7 

1871 

4678 

3742 

69 

341 1 

1497 

In gcldition to thc teclinicnl rcquirementq in qcct ion 2 :~hovc, thc design rcquircmcnts 

for the mofor/gencrntor ivcrc qpccificd by rcfcrcncc X 3s Yhown previously in Tahle 4-2. 

Thc approach was to qizc an clcctxical gcncrator drivcn hy a ga5 turbinc to supply powcr 

to an electric motor having a dircct drive to the helicopter rotor. 

5.0 - 1 ITS iWotor/Gencrator Concepts 
FR-20668 

Collaborative Planncrs Inc. gcneratcd thc informntion sliou~n in Table 5-1 for the SO0 

HP ,3000 HP . and 13140 HP motors and gcncrator-c. Early indications from thic study 

wcre that that largcr horsepower drives ~vould be bcttcr suited to HTS application5 antl 

so the IN4O HP size was added to thc originally 5pccifictl vnlucy of SO0 and 3000 

horsepower to covcr the higher portion of the powcr spectrum of prescnt helicnptcrs antl 

to more fully investigate the HTS application feasihility. Table 5-1 alw showy a dctnilctl 

weight breakdown of all the options sizcd by CPl for this studJ7. Figure 5-1 shows thc 

weights of the motors and gencrators tle~igncd by CPI for thc superconducting rotor 

only construction (SROC) and the SRASC designs. 

FIGURE 6-1 

X SROC 4 SRASC - EQ.6-1 EQ.5-2 

33UATION8 SHOWN UBED FOR 8CALINQ DATA 

Figure 5-1. (:Pi Original Designs - Electric Motor Plris Gerieralor Weight. 

5.0 - 117's Motor/C.cncr:itor Concrpts 
FR-20668 

Motors and gcncrators finally used in this study wcrc much \mailer than those initially 

requestcti from CPI because as the study progrcssctl and the mission studies were completed it was found that the power rcquircd was only about AO"/r, of the maximum sen 

level rating of the gas turbine engines; and the tail rotor powcr was small relativc to the 

main rotor power and so better transmitted mcchanically. Therefore the 

motor/gcnerator weights and cooling loads had to be scaled to the: smaller sizes. Several 

types of losses were identified: A type losses arc cold section thermal losses. B typc 

losses arc wiring ohmic losses and C typc losses arc electromagnetic shield thermal 

losses. Scaling was done to represent thc type A,R,and C losses for thc SRASC equipment while for equipment with efficiencies of lcss than or equal to 99.9% only typc A 

and I3 losses were calculatcd since type C losses arc negligible for thc SROC machincry. 

Elcctrical conduits and cnblcs were assumctl to be of thc supcrconducting types and 

power transmission losses were assumed to be ncgligible. 

Motor/Gcnerator weight scaling was done using the following cquations: 

SRASC motor and gencrator weight = 0.27S*HP + 163 Ibs. Equation 5- I. 

The maximum fit error in equation 5-1 is less than 1.51'0. 

SROC motor and gcnerator weight = 0.343*HP + 200 Ibs. Equation 5-2. 

The maximum fit error in equation 5-2 is less than 2.0%. 

Motor/Generator heat load estimates and fan power rcquircments were 

obtained from the following equations: 

SRASC rnotor/gcnerator heat load =0.2666*HP + 19.7 watt5 Equation 5-3. 

for the type A plus B losscs in the machincry with the SRAC. 

The maximum fit error is less than 1.3%. 

SRASC motor/gencrator fan load =0.06SS*HP + 631 watts Equation 5-4. 

which is required for thc typc C losses in the machinery with the 

SRASC. The maximum fit error is lcss than I I.'%, which is high because 

the same size fan,%lower was specified for both thc 500 and the 3000 HP 

machines and the fit gives an equal plus and minus error to accomodatc 

thi? data. 

SROC motor/gencrator heat load = 74S.7*2 HP*( I -(ETA/100)) watts Equation 5-5. 

in watts for the machincry with thc SROC. This includes both thc types 

A and B losses since typc C losses are negligiblc for SROC nixhines. 

5.0 - 1 ITS Motor/C;enerxtnr ('oncrpts I6 
FR-20668 

Figures 5- I and S-2 show that tlic equations juqt prcscntctl arc good rcprcwn t n tion5 

of the CPI supplied data. 

4000 

3000 

2000 

1000 

0 

-0 

/--- 

................... : ......................... ;, .................. ; ..... ...................... :. ........ :... ... ...................... 

/VH x 

............................ ................... .......................... .......................................................................................... 

X , ...... ,.2- 

0 2 4 6 0 IO 12 14 16 

CPI DESIQN HORSEPOWER (ThOu8andd 

FIGURE 6-2 

0 8RA8C A.* 1. Loeeer EIQ.S-3 

X 8RASC C LOSSES EQ.S-4 

- 

17 

5.0 - I ITS Motor/Gcnerator Conccpts 
FR-20668 

6.0 - Closed Loop Cooling Concepts 

Cryocooler weight and power scaling were done using data supplied by UTRC and 

5hown in Figures 6-1 and 6-2, respectively. UTRC suggested tlint we use the Stirling 

cycle data so that is what was done for this study. The data did not cover the cooling 

capacity range needed and so the data was cxtrapolatctl 3s can be wen by the points on 

these figures marked as scaled data. Reference 17 was used by UTRC to guide their 

recom mknd a tions. 

10 

1 

...................... ............................. 

0.1 

........................... 

0 

UJ 0.01 a 

0) 1 10 100 1000 10000 100000 1000000 

k 

COOLING CAPACITY (WTTB) 

FIQURE 6-1 

- STIRLING CYCLE - ERAYTON CYCLE * SCALED DATA 

~ ~~ 

Figure 6-1. Cryocooler Specific Weight vs. Cooling Capacity 

6.0 - <:lased I,oop Cooling Concepts 

I 
FR-20668 

0.1 

0.0 1 

0.301 

COP (WTTS COOLING PER WTT POWER INPUT) 

1 10 100 1000 10000 100000 100000 

COOLINQ CAPACITY (WTTS) 

FIQURE 8-2 

- STIRLING CYCLE - BAAYTON CYCLE * SCALED DATA 

Figure 6-2. Cryocooler (:oefTtcirnt of Pcrfnrrnance vs. <:ooling 

6.0 - Closed I,oop Cooling Conccpts 
FR-20668 

The spccific weight data uscd ran from about 200 to 600 pounds of cryocooler pcr 

kilowatt of cooling rcquircd. The power input rcquircd to achicvc thc cooling effect ran 

from 12 to 18 kilowatts of power input required for cach kilowatt of cooling rcquirccl. 

Inversely then, the cocfficicnt of performance (COP) is 0.056 to 0.083. 

The cooling requirements from Collaborativc Pla ti ners, Inc. wrc .;calcd and uscd to 

size the cryocooler dcsigns shown in Tahlc 6- I. Sizing was tlonc u.;ing 100% powcr a.; 

the design requirement and thus this sizing shows that cach cooling load required a relatively large power input to be supplied to thc cryocooler system which is also relati\cly 

heavy. 

Table 6-7 - Stirling Cvcle Scaled Data Cryocooler Characteristics for HTS Applications. Superconducting Rotor and Stator Construction (SRASC) and Superconducting Rotor Onlv Construction (SROC) 

Helicopter Power Plant: Gas Turbine Maximum HP Rating at Sea Level 

502 HP 3606 HP 13407 HP 502 HP 3606 HP 13407 HP 

Key Items Summary SROC MotorlGenerator SRASC MotorlGenerator 

Cooling Required 0.4742 3.229 11.773 

at Max Power (K W) 

0.104 0.596 2.124 

Cooler Weiglit (LB) 190 872 24 72 62.4 232 616 

Power Input (K W) 7.075 44.85 150.9 1.825 8.765 29.92 

Power Input (HP) 9.49 60.14 202.4 2.45 11.75 40.12 

Table 6-2 shows that cryocoolcr system weights increase rapidly as electrical efficiency 

is reduced below the 99.9% level used in this study. 

Table 6-2 - Crvocooler Characteristics for H TS Applications. Comparison at Lower 

Electrical Efficiencv Levels Usinn the 7894 HP Case with Superconducting Rotor Only 

Construction (SROC). 

Efficiencies (76) 

Motor 

Generator 

Cooling Required 

at Max Power (KW) 

Power Input (K W) 

Power Input (HP) 

Cooler Weight (LB) 

MotorlGen. Wt. (LB) 

TOTAL SYS. WT. (LB) 

99.9 

99.9 

11.77 

151 

202 

24 72 

2908 

5380 

99.7 

99.7 

35.32 

447 

600 

7064 

2908 

9972 

99.5 

99.5 

58.87 

740 

993 

I I185 

2908 

14093 

99.0 98.0 

99.0 98.0 

117.7 235.5 

1 4 72 2907 

1973 3898 

21 780 42383 

2908 2908 

24688 4529 1 

6.0 - Closed Imop Cooling C'nnrepts 20 
FR-20668 

7.0 - Expendable Refrigerant Concepts 

, 

I Expendable cooling LN2 t.cquircments and the cquipment rcquircd was sized using a 

computer code written for this application. Liquid nitrogen wnc, uscd as the expendable 

coolant and the mission average power was used to 4zc the LN2 volume required. A 

systcrn boiloff allowance was added to cover the potential hcat gain by the tank, 

plutnbing, and motor/gencrator structure. A sphcrical tank was chosen to minimize tlic 

outside surface area for a given volume of tank. Thc tank wall thickness of 70 mils w;1s 

used to accommodate the system pressure and provitlc 9tructur;ll integrity. Urethane 

foam insulation was assumed to cover the LN2 storage tank and providc a thermal 

barrier to the surroundings. Thirty percent of the LN2 cooling \,olume was pro$-ided for 

tank pressurization space (ullage). Plumbing weights were estirnatcd to be 2.5'!'0 of the 

tank weight. 

Chilldown of the motor,gcncrator, tank, plumbing and insulation was assumed to be 

done beFore the tank was topped off to minimize the LN2 necdccl to be carried in flight 

and only thc amount of LN2 necessary for steady cooling and boiloff from tank heat 

gain would be supplied in flight. 

An ambient air blower for cooling the SRASC motor/gcncrntor was also included and 

the size used is shown in Figure 8-2. The power for this hlomer can then he an electric 

motor drawing current from the electrical generator. 

Table 7-1 providcs a summary of expendable LN2 system qizing for all three power 

level applications. Some added benefit due to thc LN2 use (wcigh t loss) was accnun tecl 

for in that the aircraft wight reduces throughout the mission dtic to the LN2 coolant 

expended for motor/ gcncrator coding. 

I 

7.0 - Evpendahle Refrigernnt Concrpts 21 
FR-20668 

Table 7-7 - Expendable Refrigerant Summary for Systems Usinq Superconductinq 

Rotor and Stator Construction (SRASC) and Superconducting Rotor Onlv Construction (SROC). 

Helicopter Power Plant : Gas Turbine Maximrim HP Rating at Sea Level 

502 HP 

KEY ITEMS SUMMARY 

Cooling Required 

At Max Power (K W) 

Flight Time 

(Minutes) 

Mission Average 

Power Factor 

(?4 of Max Power) 

LN2 Boil Off 

Margin 

(06 LN2 Vol) 

LN2 Volume 

(Cu Ft.) 

Internal Tank 

Volume (Cu ft.) 

Internal Tank 

Radius (Inches) 

External Tank 

Radius (Inches) 

External Insul. 

Radius (Inches) 

LN2 Weight 

(Lbs.) 

Tank Weight 

(Lbs.) 

Insulation 

Weight (LBSI 

Plumbing and 

Valve Wt.(LBS) 

Air 0lower (10s.) 

Tot. Hdware. Weigh1 

(NO LN2) (LBS.) 

Tot. Cooler Weight 

Inc. LN2 (LBS.) 

0.4742 

200 

47 

30 

0.764 

0.99 

7.43 

7.50 

13.5 

38.5 

14.01 

14.82 

3.5 

None 

32.33 

70.83 

3606 HP 13407 HP 

SROC MotorlGenerator 

3.229 

160 

51 

15 

3.99 

5.19 

12.89 

12.96 

18.96 

20 1.3 

40.03 

33.74 

10.5 

None 

86.28 

287.6 

11.773 

I60 

51 

10 

13.93 

18.1 1 

19.07 

19.62 

25.6 

702.1 

96.5 

67 36 

24.12 

None 

188.0 

890.0 

502 HP 3606 HP 13407 HP 

SRASC MotorlGenerator . 

0.104 

200 

47 

60 

0.206 

0.268 

4.79 

4.86 

10.87 

10.39 

5.88 

8.5 

1 .47 

75.0 

90.9 

101.2 

0.596 

160 

51 

25 

0.80 1 

1.04 

7.54 

7.6 1 

13.61 

40.39 

14.5 

15.1 

3.62 

81.0 

I 14.0 

154.6 

2.124 

t60 

51 

15 

2.63 

3.42 

I1.2 

11 27 

17.3 

132.0 

31.8 

27.1 

7.95 

93.0 

160.0 

292.3 

7.0 - Expendahle Refrigerant Concepts 22 
FR-20668 

Figure 7-1 shows the cffccts of cooling load on weight for thc cxpcndabie LN2 cooling 

system hardware,the LN2 itself and the total of thc two as a total system weight. 

1000 

800 

600 

400 

200 X 

n 

0 I 

0 2 4 6 8 10 12 14 

COOLlNd LOAD IN KILOWTTS 

Figure 7-1 

++- SRASC HARDWRC- SRA8CLN2 - SRASC SYSTEM 

a $ROC LN2 SROCHARDWRE 8ROC SYSTEM 

'1 TANK WLL 18 70 MIL8 WITH 6' INSULATION 

1igw-e 7-1. Expcndnhle Rcrripcrnnt Systcm Wriplil 

Table 7-2 shows that expcndablc LN2 system weights incrcrrse rapidly as electrical 

efficiency is rcduccd below the 99.9%) lcvcl used in this study. 

7.0 - Evpentluhle Refrigernnt Conrrpts 2.3 
FR-20668 

Table 7-2 - Cooler System Weights for HTS Applications at Various Electrical Efficiency Levels usinq the 7894 HP Case with Superconducting Rotor Onlv Construction (SROC). 

Efficiencies (76) Motor. 99.9 99.7 99.5 97.0 98.0 

Generator 99.9 99.7 99.5 97.0 98.0 

Cooling Required 11.77 35.32 58.87 117.7 235.5 

at Max Power (KW) 

Cooler Weight (LB) 890 2378 3863 7269 1426 1 

MotorlGen. Wt. (LB) 2908 2908 2908 2908 2908 

Total System Wt. (LE) 3798 5286 6771 10204 17169 

7.0 - Eupendahle Refrigerant Concepts 24 
FR-20668 

8.0 - Room Temperature HTS Systems 

Thc development of room ternperaturc wpcrcontfuctors could producc revolutionary 

changcs in aeronautical propulsion systcms. With room tempera tiire (300K) superconductors, the LN2 cryogcnic cooling systcm would no longer be nccdctl. And cooling of 

room temperature superconductors would be accomplishcd with air supplied by ;I 

blower or fan. 

Figure 8- I shows the weights of air coolcd HTS niotor/gcnern tor propulsiori 5ystcmc 

for helicopters as a funchn of helicopter engine power. Figure 8-1 show that thc 

SROC weighs roughly 20 pcrccnt more than the SRASC over a considerable power 

range. The difference between the lower two curves in Figure 8-1 rcprescnts the wight 

of the blower system. And, as wen in this figure, thc weight of' thc air blower is relatively small. 

8.0 - Room 'Temperature I ITS Systems 2s 
FR-20668 

3000 /x 

I I I I I I 1 

0 2 4 6 8 10 12 14 16 

HELICOPTER HORSEPOWER IN THOUSAND8 

FIGURE 8-1 

-H- 8ROC M+Q * 8RASC M+Q + 8RASC M+Q+BLOWFR SYS 

AIR BLOWERS INCLUDED FOR 8RA8C SYSTEM8 

Figure 8- 1. Room Temperature (300°K) FITS System Electric Motor Plus Generator Weight. 

8.0 - Room Temperature 1I'l.S Sgstcms 26 
FR-20668 

Figure 8-2 shows the electrical air blower weight. The blower systcm weights are for 

induction motors, blowers and ducting. The blowers and ducting were estimated to hc 

20% of the motor weight. Thus, the air cooling system weight is the sum of the these 

weights as shown in Figure 8-2. 

3 4 6 6 7 0 9 10 0 : 1. 2 

MOTOR HORSEPOWER 

MOTOR WEIGHTS FROM MARKS' HANDBOOK 

Figure 8-2. Cooling Air Blower System Weight - lndriction Electric Motor and Air Blower. 

8.0 - Room 'I cmpcrature 11'15 Systems 27 
FR-20668 

9.0 = Mission Studies 

Simulations using the Bell .Jet Ranger, the Sikorsky Black H<Twk and the Sikorcky 

Super Stallion helicopter.; were executed on the HESCOMP hklicoptcr design and performance program to develop trade factory in the 500, 3000, 4ind 13000 total chnft 

horcepowcr (SHP) range. The trade factors were devclopcd for ’’1 uhbcrized” versions of 

the helicopters to show the total magnitude in terms of change in grosy nTeight,fucl 

burned and delta SHP required due to a perturbation in transmission weight wlicn the 

helicopter is designcd to perform a given mission scenario. 

The HESCOMP program scalcs all components of the helicoptcr’s structure and propulsion system (except cabin size for a gi\m payload requirement) to develop a new 

configuration with the same disc loading , radiu~jrange capability,and design hoverira le 

of climb capability at thc design sizing ambient altitude and tcmpcrature day. Trade 

factors can then be determined relative to the base configuration in the follon-ing manner: A change in transmission weight or powcr required (both qimulated by multiplying 

factors) is made to the HESCOMP model and the mission is Ilown to determine the 

change on TOGW, fuel burned, and SHP requircmcnts. These arc uscd as the tradc 

factors for that particular helicopter. Tables 9-1A, R, and C are the results of these 

HESCOMP mission ”flights” and provide the basis for the tradc factors. As noted prcviously, the results of thcse mission studicy show the influencc of‘ engine weight and fucl 

flow perturbations of TOGW, required fuel weight and ?haft HF’. 

9.0 - Mission Studies 28 
FR-20668 

Table 9-1A - AEP HESCOMP Trade factor Study ... Bell Jet Ranger with Allison 

250-C30 Engine. 

Take Off 

Gross Wt. 

TOGW (Lbs.) 

Total 

Engine WT. 

(Lhs.) 

fuel Req'd 

(Lbs.) 

Delta TOGW 

(%) 

Delta TOGW 

(Lbs.) 

Delta 

Engine Wt. 

(Lbs.) 

Delta fuel 

Req'd (%) 

Delta Fuel 

Req'd (Lbs.) 

Max. Shaft 

Horsepower 

Delta Shaft 

Horsepower (96) 

Base 

3117 

178 

5.39 

Base 

Base 

Base 

Base 

Base 

502 

Base 

+ 20% 

Eng. Wt. 

3184 

2T8 

547 

+2.150 

+ 67 

+ 40 

+ 1.48 

+8 

512 

+ 1.99 

-20% 

Eng. Wt. 

3052 

140 

53 I 

-2.085 

-65 

-38 

-1.48 

-8 

492 

-1 99 

-4O?G 

Eng. Wt. 

3000 

103 

530 

-3.754 

-117 

- 75 

- I .67 

-9 

485 

-3.39 

+ 5% 

Fuel 

Flow 

3170 

181 

5 70 

+ I .7OO 

f 53 

+3 

+ 5.75 

+ 31 

510 

+ 1.59 

-506 

Fuel 

Flow 

3065 

176 

509 

- 1.668 

+ 52 

-2 

-5.57 

-30 

494 

-1.59 

9.0 - Mission Stiltlies 
FR-20668 

Table 9-76 - AEP HESCOMP Trade Factor Study ... UH-60 Black Hawk with GE T-700 

Take Off 

Gross Wt. 

TOG W (Lbs.) 

Total 

Engine WT. 

(Lbs.) 

Fuel Req'd 

(Lbs.) 

Delta TOGW 

(06) 

Delta TOGW 

(Lbs.) 

Delta 

Engine Wt. 

(Lbs.) 

Delta Fuel 

Req'd (%) 

Delta Fuel 

Req'd (Lbs .) 

Max. Shaft 

Horsepower 

Delta Shaft 

Horsepower (%) 

9.0 - Mission Studies 

Base 

16747 

8 72 

2135 

Base 

Ease 

Base 

Base 

Base 

3606 

Base 

+ 20x 

Eng. Wt. 

17210 

1075 

2191 

+ 2.765 

+ 463 

+ 203 

+ 2.62 

+ 56 

3 702 

+ 2.66 

-2006 

Eng. Wt. 

16363 

682 

2089 

-2.293 

-384 

-190 

-2.15 

-46 

3526 

-2.22 

-40'6 

Eng. W1 

15969 

500 

2040 

-4.645 

-778 

-372 

-4.45 

-95 

3445 

-4.46 

+ 5s:; 

Fuel 

Flow 

170Pl 

890 

2283 

+ 2.054 

+ 3'14 

+ 18 

+ 6.93 

+ 148 

36 7'7 

+ 1.97 

-5?6 

Fuel 

Flow 

16471 

859 

1998 

- 1.648 

-276 

-13 

-6.42 

-137 

3589 

-I .58 

30 
FR-20668 

Table 9-7C - AEP HESCOMP Trade Factor Study ... CH-53 Super Stallion with 

T65-GE-47 5 Engines. 

Take Off 

Gross Wt. 

TOG W (Lbs.) 

Total 

Engine WT. 

(Lbs.) 

Fuel Req’d 

(Lbs.) 

Delta TOGW 

(04) 

Delta TOGW 

(16s.) 

Delta 

Engine Wt. 

(Lbs.) 

Delta Fuel 

Req’d (%) 

Delta Fuel 

Req’d (Lbs.) 

Max. Shaft 

Horsepower 

Delta Shaft 

Horsepower (5%) 

Base 

49022 

2204 

6066 

Base 

Base 

Base 

Base 

Base 

13407 

Base 

+ 20% -2006 

Eng. Wt. Eng. Wt. 

50405 47741 

2718 1718 

6235 5917 

+ 2.82 -2.61 

+ 1383 -128r 

+514 -486 

+ 2.79 -2.46 

+ 169 -149 

13779 13062 

+ 2.77 -2.57 

-401?6 

Eng. Wt 

46.525 

1256 

5774 

-5.09 

-2497 

0 48 

-4.81 

-292 

1 2 735 

-5.01 

+ 506 

Fuel 

Flow 

50 103 

2252 

649 1 

t 7.21 

t 1081 

+ 48 

+ 7.01 

+ 425 

13697 

+2.16 

-5% 

Fuel 

Flow 

480 12 

2159 

5669 

-2.06 

-1010 

-45 

-6.54 

-397 

13135 

-2.03 

9.0 - Mission Stiirlics 
FR-20668 

The baseline helicopter configurations were obtained from the actual geometric characteristics of each hclicopler and the weight delta LVV;IF taken from detailed weigh t 

breakdowns furnished by the helicoptcr contractor';. Mision profilcq were thow clcvcloped for previous U.S. Army contracted TAGG-M propulsion system Wdics at Pr-att 

& Whitncy. Rotor drive system weights and power rcquiremcnts uscd in thiy study mere 

from the HESCOMP computer program information and are shown in Table 9-2. 

Power Rating" 353 2489 9178 

Total at Rotors 

Table 10-4 shows power train performance as well as thc TOGW for the h3rdware 

weight changes and the TOGW due to the powcr extraction tliffercnces. Atltlitinnal 

weight changes due to the diffcrences in power use were calculated and are also Thomn 

on table 10-4. 

Table 9-2 - HESCOMP Weight and Horsepower Summary 

HELlCOP TER Bell Jet Sikorsky Sikorsky 

WEIGHT Ranger Black Hawk Super Stallion 

SUMMARY (LBS.) 206-8-JR3 UH-GOA CH-53E 

Take Off Gross 31 17 16747 49022 

Weight (TOG W) 

Engine(s) 178 (I) 872 (2) 2204 (3) 

Drive Systems: 

Main Rotor 182 1387 5731 

Tail Rotor 25 97 8 72 

Total 20 7 1484 6603 

HELlCOP TER 

HORSE P 0 WE R 

s UM M A R Y (npi 

Engine (s) 

Maximum Rating 502 3606 13407 

at Sea Level 

, 

I Main Rotor 318 2165 7894 

Tail Rotor 35 304 1284 

Accessories 33 100 153 

Main Rotor 11 74 283 

Transmission 

Losses (3%) 

"WHEN SIZED FOR THE FOLLOWlNG CONDlTlONS PER MlSSlON SPEClFlCATlONS ; 

BELL 2068 

SIK. UH-60 

POWER RATING AT 95% MAX POWER, 95"DAY, 2000 FT., 300 FPM CLIMB. 

POWER RATING A7 95% MAX POWER, 95"DAY, 4000 FT., 500 FPM CLlMB. 

SIK. CH53-E POWER RATING AT 95% MAX POWER, 95"DAY, 4000 FT., 500 FPM CLIMB 

9.0 - Mission Stiltlies I 

r 

32 
FR-20668 

10.0 - Weight, Power, Performance and Cooling Audits 

Table 10-1 shows the wights of thc kcy components of the ~tutiy which wrc calculated and providcs an ovcrview of thc study result%. The 5ystcm weights were ohtaincd 

from compiling all the individual weights from the indi\4tlu~1 contributors as were the 

powcr extractions sliown on Table 10-2. 

Table 10-1 - HTS System Weights Key Items Summary 

Helicopter Power Plant : Gas Turbine Maximiin7 HP Rating at Sea Level 

502 HP 3606 HP 13407 HP 502 HP 3606 HP 13407 HP 

KEY ITEMS SROC MotorlGenerator SRASC MotorlGentvator 

SYSTEMS WEIGHTS 

AUDIT (LBS.) 

Gas Turb. Engines 

Power Drive Train 

I. Mechanical Sys. 

A. Main Rotor 

8. Tail Rotor 

Total Train 

2. HTS Systems : 

A. Closed Loop 

Cryocooler 

MotorlGen. 

Total ClLoop 

Tank 

Insulation 

Plum 0 ing 

fan 

MotorlGen. 

Hardware lot. 

LNZ 

Total Expend. 

C. Air Blower SYS 

f anlBlwrlDuct 

MotorlGen 

Total Air cool 

B. Expendable LN2 

3. WEIGHT CHANGE 

FROM BASE (IA.) 

a. CILoop System 

b. Expnd LN2 Sys 

c. Air Cooled Sys 

~- 

178 

182 

25 

20 7 

190 

309 

499 

14 

15 

4 

none 

309 

34 1 

39 

380 

none 

309 

309 

317 

198 

127 

8 72 

1387 

97 

1484 

8 72 

943 

1815 

40 

34 

11 

none 

943 

1029 

20 1 

1230 

none 

943 

943 

428 

-157 

-444 

10.0 - Weight, Power, Performance sild Cooling Aiitlits 

2204 

5731 

8 72 

6603 

24 72 

2908 

5380 

97 

67 

24 

nolie 

2908 

3096 

702 

3798 

none 

2908 

2908 

-35 I 

- 1933 

-2823 

178 8 72 2204 

182 1387 5731 

25 97 8 72 

20 7 1484 6603 

62 

250 

312 

6 

9 

1.0 

75 

250 

34 1 

10 

35 1 

75 

250 

325 

130 

169 

143 

232 

758 

990 

14 

15 

4.0 

81 

758 

8 72 

40 

912 

81 

758 

839 

-397 

-475 

-548 

616 

2334 

2050 

32 

27 

8.0 

93 

2334 

2494 

132 

7626 

93 

2334 

2427 

-2781 

-3105 

-3304 
Table 70-2 - HTS System Power Extraction Key Items Summary 

Helicopter Power Plant : Gas Turhine Maximum HP Rating at Sea Level 

KEY ITEMS 

POWER AUDIT 

{HORSEPOWER) 

Power Requirement 

Maximum to Main 

Maximum to Tail 

Maxiniiim Total 

Main Rotor Power 

I. TRANSMlSSlON 

A. Mech Drive 

€3. HTS Conduit 

(3% Loss) 

(0% Loss) 

2. HTS Sys. Rqmts 

A. Closed loop 

(Crycool Pwr) 

8. Expend LN2 

(Air Fan) 

C. Air Cool fan 

3. HTS-Mech Loss 

A. Closed loop 

B. Expend LN2 

C. Air Cool fan 

502 HP 3606 HP 13407 HP 

SROC MotorlGenerator 

3 18 2165 7894 

35 304 1284 

353 2469 9178 

9.54 64.95 236.8 

Negl. NPgl. Negl. 

9.49 60.14 202.4 

none none none 

none none none 

-0.52 -4.81 -34.4 

-9.54 -65.0 -237.0 

-9.54 -65.0 -237.0 

502 HP 3606 HP 13407 HP 

SRASC MotorlGenerator 

318 2165 7894 

35 304 1284 

353 2469 9178 

9.54 64.95 236.8 

Negl. Negl. Negl. 

3.32 12.79 4 1.66 

0.874 1 .04 1.54 

0.874 1.04 1.54 

-6.2 1 -52.2 -195.2 

-8.66 -63.9 -235.0 

-8.66 -63.9 -235.0 

10.0 - \Vcight, I’arvcr, rcrformancc and Cooling Autlits 34 
FR-20668 

The valucs shown haw bcen scalccl to tlic rcquiretl horscpowcr conditions from thosc 

supplicd by the individual contractors whcrc ncccssary. Tahlc IO- I shows wcights for 

the gas turbine engines uscd in thc threc hclicopters sclcctcd for this study. The values 

shown are thc total cngiric weights for thc aircraft. Thc power drivctrain weights shown 

have been apportioned to the main and tail rotor Fystcms by HESCOMP. The total 

weights shown for the closed loop cryocoolcr 5ystemF arc thosc .;caIctl valucs necded for 

replacing the main rotor drive with an HTS motorigcncrator, cryocooler and air blowcr 

systcm where requircd. The total weight shown for the cxpendablc LN2 systcm incluclcs 

HTS motor and generator, the supporting cquipnicnt and a full tank of LN2. 

The total wcights shown for the air coolcd fan (room tcmperaturc 300°K HTS) option 

arc thosc scaled values necdcd for replacing the main rotor drivc with an HTS niotor. 

gcnerator and an air blowcr wherc nceded. 

Item 3 on Table 10-1 shows the wcight change from the basc mcchanical main dri\~ 

systcm mxlc by replacing it with an HTS systcm to transmit main drivc systcm powcr 

from the gas turbine cngine to thc mgin rotor of each hclicoptcr. 

Table 10-2 showy a power audit of the key items usccl in this study. Maximum powcr 

requirements are shown for the main and tail rotors. The main rotor values are thc 

horsepowers used for designing the HTS systems to rcplace the con\. entional mechanical 

system. Powcr transmission losses are shown for cach system but not for the HTS systems since they are considcred negligible. All thc FfTS system power requiremcnts are 

shown but the most significant powcr usagcs are thc ones ncetlcd to drive the cryocoolers 

for the lower cfficiency (99.9%) SROC motor generators. Item 3 on Table 10-2 shows 

the powcr uqagc differences between the mechanical and the FITS systems; thc nicchanical losses are for geared transmissions while thc only significant HTS lossc5 arc for 

the cryocoolers. 

10.0 - Weight, I’ower, I’trformancc and Cooling Audits 
FR-20668 

Tablc 10-3 shows thc ~cnlcd MTS motor/gcncrator cooling load rcquirctncnts (typc 

A.and B.lo~scs) as ivcll as thc motor/gcncrator cfficicncics rcla tivc to thcsc losscs. 

Table 10-3 - HTS Electrical System Cooling Requirements Key ltems Summary 

Helicopter Power Plant: Gas Turbine Maximum HP Rating at Sea Level 

Helicopter Power Plant: Gas Turlline Maximum HP Rating at Sea Level 

KEY ITEMS SROC MotorlGenerator SRASC MolorlGeneralor 

COOLiNG 

SYSTEM AUDIT 

502 HP 3606 HP 13407 HP 502 HP 3606 HP 13507 HP 

* 

Efficiencies (?6) 

Motor 99.9 99.9 99.9 99.98 99.98 99.98 

Generator 99.9 99.9 99.9 99.98 S9.98 99.98 

Cooling Required 

Max Load (KWJ 

Max Load (HP) 

0 474 3 229 11 77 0 104 0 597 2 12 

0 636 4 33 15 79 0 140 0 80 2 85 

10.0 - Weight, I’nwcr, Performancc mil Cooliiig Aridits 
FR-20668 

Table 10-4 provides a drive systcm pcrformance comparison of spccific power 

(HPiLB) for the conventional mechanical systcms and the HTS systcms. Alw included 

in this table are the TOGW changes rclatctl to each of thc systcms. The weight changes 

are presented in pounds and in percent of TOGW. Thc cffccts of system weight and 

power on TOGW arc shown separately. Figure 10-1 shows the hclicopter TOGW 

changes due to each HTS system. The HTS changcs can bc secn to be beneficial at 

helicopter horscpowers above about 1000 HP. But for the SROC with cryocoolcr, HTS 

is not a good choice at any power levcl undcr I1000 HP. The cxpcndablc LN2 and thc 

room teinperaturc HTS SROC systems become competitive from 2000 to 3000 FIP 

where the helicopter TOGW begins to drop. 

10 

0 

-10 

-20 

0 2 4 6 8 10 12 14 16 

HELICOPTER HORSEPOWER IN THOUSAND8 

0 

Z 

< 

X 

HTB SYSTEMS USED TWO EFFICIENCY LEVEL8 

* 8ROC AIR COOLED 

+ SRASC AIR COOLED 

Figlire 10-1. llclicopter TOGW Comparisons vs. Powrr I1 1’s Rcpl;~crmcr~t or Main Rotor Ilrivc. 

10.0 - Weight, Power! Performance and Cnoling Aiitlits 37 
FR-20668 

Table 70-4 - HTS System Performance and TOGW Key Items Summary 

Helicopter Power Plant : Gas Turbine Maximrim HP Rating at Sea Level 

502 HP 3606 HP 13407 HP 502 HP 3606 HP 13407 HP 

KEY ITEMS SROC MotorlGenerator SRASC MotorlGenerator 

MAIN ROTOR DRIVE 

PERFORMANCE 

Specific Power 

(H PIL B) 

1. Mechanical Sys 1.747 

2. HTS Systems 

a. CILoop System 0.637 

., 6. Expnd LN2 Sys 0.887 

c.Air Cooled Sys 1.029 

1.561 1.377 1747 1.561 1377 

1.019 2.187 2.676 

0.919 2.427 3.084 

0.978 2.580 3.253 

1.193 1.467 

1.917 2.290 

2.296 2.715 

TOG W CHANGE (LBS) 

Due to Changes 

In System Weight 

a. ClLoop System 

6. Expnd LN2 Sys 

c. Air Cooled Sys 

go 7 -930 

-546 -6053 

-94 1 -7481 

212 

26 7 

233 

-842 

-1049 

-1162 

-7370 

-8756 

-8403 

517 

288 

20 7 

-0.3 

-64.2 

-64.2 

516 

223 

143 

16.5 

9.2 

6.6 

0.0 

-2.0 

-2.0 

16.5 

71 

4.5 

Due to Changes 

In Power Used a. ClLoop System 

6. Expnd LN2 Sys 

c. Air Cooled Sys 

Due to Changes in 

Weight and Power 

a. CILoop System 

6. Expnd LN2 Sys 

c. Air Cooled Sys 

TOG W CHANGE I%) 

Due to Changes 

In System Weight 

a. CILoop System 

b. Expnd LN2 Sys 

c. Air Cooled S ys 

. Due to C/?anges 

In Power Used 

a. CILoop System 

6. Expnd LN2 Sys 

c. Air Cooled Sys 

Due to Changes In 

Weight and Power 

a. CILoop System 

6. Expnd LNP Sys 

c. Air Cooled S ys 

-41.7 

-58.2 

-58.2 

- 726 

-875 

-875 

-23.1 -128 

-313 -88 1 

-313 -88 1 

-25 I 

-308 

-308 

-8096 

-963 1 

-9278 

884 -1058 

-a59 -6933 

-1254 -8362 

~ 170 

209 

175 

- 1093 

-1357 

-1470 

5.4 -1.8 

-3.2 -12.3 

-5.6 -15.2 

6.7 

8.5 

7.4 

-5.0 

-6.2 

-6.9 

-15.0 

-17.1 

-1 7.8 

-1.3 

-1.8 

-1.8 

-1.4 

-1.7 

-1.7 

-0.1 -0.2 

-1.8 -1.7 

-1.8 -1.7 

-1.5 

-1.8 

-1.8 

- 16.5 

-18.9 

-19.6 

5.2 -2.1 

-5.1 -14.1 

-7.4 - 17.0 

5.4 

6.7 

6.6 

-6.5 

-8.1 

-8.7 

.? 8 10.0 - Wright, rowcr, PcrFormnncc and Cooling Aidits 

~- ~ - ~- ~ ~- 
FR-20668 

A 

Figurcq 10-2 and 10-3 show the actual systcm wcights for both the SROC and the 

SRASC rnotor/generator construction. Thc mcchanical dri\.c syqlcm weight is alco includcd to gic-c a rcfcrcncc for the con\cntional systcm wcight wl1ic.h thc FITS cystcms 

would replacc. Tlicce figures also show that wmc HTS cyctcni wcighty are bcncficial 

relative to mechanical drivc systcm weights at hciicoptcr powcr Icvcls of about 1000 FTP. 

- i 

Figure 10-2. 111s SROC and Mechanical Drive System 1 IIS Rcplaccmcnt nf hfain Rotor Drive. 

10.0 - Weight, I’orvcr, Performance and (’ooling Audits 39 
FR-20668 

6- 

4- 

3- 

2- 

1- 

0 

10 12 14 16 0 2 4 6 8 

HELICOPTER HORSEPOWER (ThOUmndd 

Figure 10-3 

* MECHMAIN ROT.DRIVI 

* EXPENDABLE LN2 $Ye. 

-+ CRYOC30LER SYSTEM 

-+ AIR COOLED 8% , 

j SYS.WT8.1NCLUDE HTS MOTORS & QENERATORS - 

Figure 10-3. I I1 S SROC and hlechanical Drice System 117s Rrplaccment nf Main Rntor Drive. 

10.0 - Weight, Power. I’crformance and Cooling ,Aidits 40 
~~ 

FR-20668 

11 .O = Conclusions and Recommendations 

Figure 10-1 shows that the FITS concept has weight benefits nhcn applied to a helicopter drive system. The hest applications (ncgatiic changes in TOGW) arc the Iargcihorsepower ones as can bc wen in the 2000 and the 13140 HP cascs where a take off 

gross weight -saving is apparent. The closed loop refrigeration systcms arc too hcavy for 

all cases exccpt the high horsepower, high efficiency oncs. Thc expendable LN2 and the 

air cooled (300°K HTS) arc applicable to lonw horscpowcr hclicopters with air coolccl 

being the most wiciely acceptable for the Iielicoptcn in this study. 

The original powcr sizing requiremcnts were chosen as tlic gas turbine power horscpower ratings. As the study progressed, two obscrvations changctl the initial elcctric 

motor/genera tor sizing rcquircments. Tlic first wac that the HTS powcr tr;ln-sriiiscion 

equipment in the lower horscpower sizes woulcl be hcackr than mechanical trancmi5qion 

equipment so the EJTS clri\.cn tail rotor motors which rcquirc rclativcly littlc powcr 

would not be beneficial escept perhaps for the CH-SJE which could go either way depending of motor efficiency and cooling methods. This lowcrccl the initial power rcyuirernents by froni 10?0 to 14?'0. Then as the qtudy progressed a second major 

observation was that the mission ytudies showed thc main rotor maximum power use 

was from only 59 to 63"h of the maximum ggq turbine power so the sizings wrc scaled 

down to the mission required levels of main rotor horscpower. 

41 
FR-20668 

Electrical efficiencies used in this study are quite high (09.9"/0 and 99.9 + "'h) and 

thought to he attainable by thc year 2000; hovmer, some questions naturally ar+ie 

aboirt what the effccts lower cfficicncics would have on the rcsults of this study. Consideration was given to this in Figiirc 11-1. which cliowc; that for the largest aircraft the 

I-ITS room tcmpcratture (aircoolcd) yystcm i? a good choice in tcrms of weight compared 

to mechanical drives at thc lower cfficiencies. Flou.cver, the LN2 system hccomcs too 

heavy at efficiencies ~CIOM~ 99..5'!4 and thc cryocooler system is too hcavy even at Y9.7"/0. 

26 

26 

15 

13 

3 

3 

z 99.9 %I 99.7 % 99.6 %J 99.0 9b 

2 

MOTOR b QENERATOR ELECTRICAL EFFICIENCY CD 

w 

a FIQURE 11-1 0 

= MECHANICAL DRIVE8 AIR COOLED SYSTEM 

0 EXPEND.LN2 SYSTEM CRYOCOOLER 8Y8TEM 

1 

I 3ROC MACHINERY FOR ALL EFFICIENCY LEVEL6 

ripre I 1-1. 1 argest I lelicoptrr Ilrive System Weights I'or Sqqtciric witli I ower I fliciciicy I cvcls 

I I .O - Cnncliisions and Ilccommcndations 42 
FR-20668 

1 

Information gained from this qtudy suggcqts that any futurc cffortr qhould bcgiii vifli 

r2 preliminary design For a large helicopter with tail rotor power requirements grcatcr 

than 500 HP so that it may be a candidate for an HTS tail rotor motor as well as a main 

rotor motor. Cooling candidates should include cryocoolcrs, and expendable LN2. For 

room temperature superconductors, an air hlower may be iised for cooling. 

t 

1 

I 

I 

I 1 .O - Concliisinns and Rccommcntlations 
FR-20668 

-- SYMBOL 

B TU 

COP 

CILoop 

CPI 

CU FT 

CU M 

D 

($000) 

ENG. 

ETA 

EXPND. 

FPM 

FT 

G 

HESCOMP 

HP 

HR 

HTS 

IN 

K 

"K 

LB(V 

KW 

LN2 

M 

MAX. 

MIL 

MIL-STD 

MPH 

NEGL 

YO 

REQ'D 

RPM 

SP.WT. 

SP .POWER 

SRASC 

SROC 

SYS. 

TAD 

TOG W 

TOT. 

UTRC 

WT 

12.0 - List of Abbreviations 

__ UNlTS 

British Thermal Unit 

Coefficient of Perfortnance for Cooling 

Closed Loop Refrigeration System (Cryocooler) 

Collaborative Planners .Inc. 

Cubic Feet 

Cubic Meter 

Change (Delta) 

Thousands of Dollars 

Engine 

Motor or Generator Efficiency 

Expendable 

Feet per Minute . Feet 

Generator 

Helicopter Conil.wter Sizing Prograrn 

Horsepower 

Hours 

Hi!gh Temperature Superconductor 

In 

Thermal Conductivity 

Absolute Temperature 

Pound(@ 

Kilowatt 

Liquid Nitrogen 

Motor 

Maximum 

Thousandths of an Inch 

Military Standard 

Miles per Hour 

Nc.gligible 

Percent 

Required 

Revolutions per Minute 

Specific Weight 

Specific Power 

Superconducting Rotor and Stator Constructior? 

Superconducting Rotor Only Construction 

System 

Technology availability date 

Take Off Gross Weight 

Total 

United Technologies Research Center 

Weight 

UNITS 

B TU 

Watts1 Watt 

CU.FT 

CU.M. 

Dollars 

FTlMin. 

feet 

Horsepower 

Hours 

Inches 

BTU FTIHRISQ. FT.I"F 

Deg. Kelvin 

LB 

Kilowatts 

In1 IO00 

MileslHour 

RPM 

L B .I Watt 

HP.1LB 

Lbs or ?C 

LBS. 
-~ 

FR-20668 

, 

I 

1 

i 

13.0 - 

13.0 = References 

1. National Bureau of Standards Technical Note 129A, 'The Thermodynamic 

properties of Nitrogen from 114 to 540" R Between 1 and 3000 PSIA', 

T.R.Strobridge NBS February 1963. 

cations" March 9.1981, 

2. Aviation Week R Space Technology. Torecast R liiventory Specifi3. MIL-STD- 1374A, Bell Helicopter Textron ,'Group Weight Stalpment for 

the Model OH-58A Helicopter', IJndated. 

4. MIL-STD- 1374A, Sikorsky Aircraft ,'Group Weight Statement for the 

Model UH-GOA Helicopter', October 12,1987. 

5. MlL-STD-l374A, Siltorsky Aircraft ,"Group Weight Statement for the 

Model CH-53E Helicopter", October 12,7987. 

6. Aviation Week & Space Technology, 'Forecast & Inventory Specificat7. Meeting Notes, 'Requirements for Task 009", Meeting at NASA Lewis 

ions' ,March 14, 1988. 

Research Center among NASA, Collaborative Planners, Inc. and Pratt 

& Whitney. December 7,1988. 

8. Pratt & Whitney Memorandum, "Design Speeds for Helicopter Application" J.Doernbach, March 6,1989. 

9. Collaborative Planners,Inc. Progress Letter #I. containing Motor 

IGenerator Sizes for the 500 and 3000 HP Cases ,March 29,19R9. 

1O.Pratt & Whitney Trade Factor Study "Sikorsky UH-GOA Black Hawk Trade 

Factor Study" W.J.Grunske and B.E.Trernbly, April 6,1989. 

Characteristics for HTS Applications, Frank Biancardi, April 6,1989. 

I I .United Teclinologies Research Center Memorandum 'Cryocooler 

12.Pratt & Whitney Trade Factor Study "Bell 2068 Jet Ranger Trade 

13.Pratt R Whitney Trade Factor Study "Sikorsky CH-SRE Super Stallion 

14.Materials Engineering Volume 78 No. 4 , '1974 Materials Selector' 

15.Collaborative Planners,lnc.Feasibility Study 'Superconducting Power 

16.Marks' Mechanical Engineers' Handbook, McGraw-Hill Book Company 

I7.UTRC supplied reference :"Cryocoolers',Graham Walker, Plenum Press 

Factor Study" W..I.C;runske and B.E.Tremhly. May 5. 1989. 

Trade Sttidy' W.J.Grunslte and t3.E.Trembly. June 1,1989. 

Reinhold Publishing Co.,lric. Stamford Conn. September 1973. 

Transmission for Helicopters' June 1989. 

New York, 1958. 

New York, 1983. This includes references to I971 NASA report by 

Jensen et.al. and to a NBS report by Strohridge. 

References 4s 
. - -- 

1. Report No. 

NASA CR 18522 

Report Documentation Page 

2. Government Accession No. 

7. Key Words (Suggested by Author($)) 

7. Author@) 

Jay Doernbach 

18. Distribution Statement 

______~ 

9. Performing Organization Name and Addrota 

United Technologies Corporation 

Pratt & Whitney 

Government Engine Business, PO Box 109600 

West Palm Beach,Fl. 33410-9600 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 44135-3191 

12. Sponsoring Agency Name and Address 

3. Security Classif. (of this report) 

Unclassified 

3. Recipient's Catalog No. 

20. Security Classif. (of this page) 21. No of pages 22. Price' 

Unclassified 58 

1 

5. Report Date i 

8. Performing Organization Report No. 

PW FR 20668 

10. Work Unit No. 

505-69-61 

11. Contract or Grant No. 

NAS3 25117 

Contract Final Report 

13. Type of Report and Period Covered 

14. Sponsoring Agency Code 

~ 

5. Supplementary Notes 

Project Manager, George E. Turney 

Aeropropulsion Analysis Office 

NASA Lewis Research Center 

6.Abstract The successful development of high-temperature superconductors (HTS) could 

have a major impact on future aeronautical propulsion and aeronautical flight 

vehicle systems. 

visioned for several classes of aeronautical systems, including subsonic and 

supersonic transports, hypersonic aircraft, V/STOL aircraft, rotorcraft and 

solar powered aircraft. 

The study presented herein examines the potential of HTS electric motors 

and generators for providing primary shaft power for rotorcraft propulsion. 

Three different sized production helicopters were investigated; namely, the 

Bell Jet Ranger,the Sikorsky Black Hawk and the Sikorsky Super Stallion. 

These rotorcraft have nominal horsepower ratings of 500, 3600, and 13400, respectively. 

Preliminary results fm this study indicated,that an all-electric HTS 

drive system produces an improvement in rotorcraft: TOGW ;for those rotorcraft 

with power ratings above 2000 horsepower. 

are up to 9% for the medium-sized Sikorsky Black Hawk and up to 20% for the 

large-sized Sikorsky Super Stallion. 

experienced a penalty in TOGW with the all-electric HTS drive system. 

Applications of high-temperture superconductors have been enThe predicted TOGW improvements 

The small-sized Bell Jet Ranger,however, 

Superconducting Electric Motors 

Rotocraft 

Propulsion 

Unclassified-Unlimited 

Subject Category 07