Joseph P. Veres

Glenn Research Center, Cleveland, Ohio

Compressor Study to Meet Large Civil

Tilt Rotor Engine Requirements

NASA/TM—2009-215641

August 2009
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Joseph P. Veres

Glenn Research Center, Cleveland, Ohio

Compressor Study to Meet Large Civil

Tilt Rotor Engine Requirements

NASA/TM—2009-215641

August 2009

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

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NASA/TM—2009-215641 1 

Compressor Study to Meet Large Civil Tilt Rotor Engine Requirements 

Joseph P. Veres 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Summary 

A vehicle concept study has been made to meet the 

requirements of the Large Civil Tilt Rotorcraft vehicle 

mission. A vehicle concept was determined, and a notional 

turboshaft engine system study was conducted. The engine 

study defined requirements for the major engine components, 

including the compressor. The compressor design-point goal 

was to deliver a pressure ratio of 31:1 at an inlet weight flow 

of 28.4 lbm/sec. To perform a conceptual design of two 

potential compressor configurations to meet the design 

requirement, a mean-line compressor flow analysis and design 

code were used. The first configuration is an eight-stage axial 

compressor. Some challenges of the all-axial compressor are 

the small blade spans of the rear-block stages being 0.28 in., 

resulting in the last-stage blade tip clearance-to-span ratio of 

2.4 percent. The second configuration is a seven-stage axial 

compressor, with a centrifugal stage having a 0.28-in. 

impeller-exit blade span. The compressors’ conceptual designs 

helped estimate the flow path dimensions, rotor leading and 

trailing edge blade angles, flow conditions, and velocity 

triangles for each stage. 

Introduction 

A notional study of the Large Civil Tilt Rotorcraft (LCTR) 

vehicle mission has been done in Reference 1. To meet the 

LCTR vehicle thrust requirements at the takeoff and cruise 

conditions, a thermodynamic cycle study of a notional 

turboshaft engine was performed in Reference 2 with the 

Numerical Propulsion System Simulator (NPSS) 

thermodynamic cycle code. The results of the engine system 

model are illustrated by the schematic diagram in Figure 1. 

Utilizing the NPSS system model, the compressor flow and 

pressure ratio requirements to produce a pressure ratio of 31:1 

at a corrected airflow rate of 28.4 lbm/sec were determined. 

The focus of this study is to perform a conceptual sizing 

study of the compressor to meet the pressure ratio and flow 

requirements of the turbine engine for the LCTR vehicle. The 

conceptual design was done with a mean-line flow 

methodology and focuses on the compressor flow path and 

key aerodynamic parameters of the rotor and stage at the 

design point condition. Off-design performance is also 

estimated for the 100 percent speed line with the mean-line 

methodology. The purpose of the conceptual design is to have 

an initial estimate of the overall compressor that could meet 

the requirements, and to identify early on any potential 

technical barriers that will need to be considered during the 

preliminary and detailed design phases. 

Compressor Conceptual Design 

The purpose of the compressor conceptual design was to 

identify potential technical challenges towards meeting the 

goals. There are several options that can be considered for the 

shape of the flow path during the conceptual design process of 

multistage compressors. The constant-tip compressor flow 

path provides the most work capability in the fewest number 

of axial stages, but can result in excessively small blade spans 

in the rear stages. A tapered-tip axial-compressor flow path 

can generally produce less work per stage because of the 

reduction of rotor-tip diameter for each subsequent stage, and 

can result in blades with larger spans in the last stages than a 

constant-tip diameter flow path. A combination of a multistage 

axial compressor with a centrifugal compressor in the last 

stage can provide a more axially compact configuration in 

comparison to an all-axial compressor, but could have a larger 

outer diameter due to the impeller and radial diffuser. These 

conceptual design options for the 31:1 compressor were 

explored in this study with the use of a mean-line compressor 

design and analysis code. The mean-line compressor flow 

analysis and design code from Reference 3 was used in this 

study. The axial rotor-blade tip speed was limited to no higher 

than 1500 ft/sec because of anticipated structural limitations, 

based on historical compressor experience (Refs. 4 and 5), as 

well as concerns about having excessively high rotor-inlet tip 

relative Mach number. The diffusion factor as defined by 

Equation (1) was limited at the design point to be on the order 

of 0.53. 

+ σ

− = − +

1 2 1

1 1 2 2

1

2

( )

( ) 1

R R W

R C R C

W

W DF U U (1) 

The diffusion factor is determined from flow and geometric 

quantities at the rotor leading edge (subscript 1) and trailing 

edge (subscript 2). The R refers to root-mean-square radius, 

while C refers to absolute velocity, and W is the relative 

velocity. The symbol σ is the blade solidity. 

As this mean-line code does not generate blade shapes, but 

only estimates leading and trailing blade angles, the values for 

the number of blades and blade solidity at the rotor tip were 

obtained from the energy-efficient engine high-pressure 

compressor of Reference 6. Each rotor was sized to have an 
NASA/TM—2009-215641 2 

inlet absolute Mach number on the order of 0.50; however, in 

the latter stages, this limit was reduced in an effort to keep the 

blade spans as large as possible. The adiabatic efficiency for 

all rotors was input into the mean-line flow code as follows: 

tip: 84, mid: 92, hub: 93. The loss coefficient for the stators of 

0.05 as defined by Equation (2) was also an input parameter. 

2 2

4 2

t S

t t

P P

P P

−

− ω = (2) 

The ω is the pressure-loss parameter through the stator 

Pt4–Pt2 normalized with rotor-exit total pressure Pt2 and rotorexit static pressure PS2. The mean-line compressor design 

analysis method of Reference 3 was used to perform the 

conceptual designs and analyses. The first conceptual design 

iteration focused on an all-axial stage design to estimate the 

overall geometric parameters of the compressor and the 

aerodynamic parameters for each stage. The aerodynamic 

parameters for each individual blade row were calculated at 

the leading edge and trailing edge to determine number of 

stages that would be required to produce the required pressure 

ratio. The aerodynamic performance of each rotor blade row 

was determined, including geometric parameters such as tip 

and hub radii and blade angles. The absolute flow angle 

entering each rotor was 0° at the design point operating 

condition. The work per blade row and the inlet and exit radii 

were varied, and the required exit blade angles were 

calculated. The diffusion factor and relative velocity ratios to 

achieve the prescribed work per blade row were limited to a 

maximum value to reduce the likelihood of flow separation. 

An iterative process was used that required variation of key 

parameters to stay within the limits of maximum diffusion 

factor per stage. The work split between stages was 

determined on the basis of the maximum work that could be 

achieved in a given blade row within the maximum diffusion 

factor limit. Also monitored during the design process was 

work coefficient per rotor as defined by Equation (3), where 

ΔH is the enthalpy rise and UTip is the rotor-inlet tip speed in 

feet per second. 

2 UTip

ΔH

φ = (3) 

For the off-design analysis a loss model correlated to rotor 

incidence was used (Ref. 3), as defined by Equation (4). 

0.0006 0.0185 0.1699 0.5187 3 2 Δη = i − i + i + (4) 

The incidence i at the rotor leading edge is the difference 

between the relative flow angle and the blade angle at the 

mean radius. The efficiency reduction at off-design values of 

rotor incidence is defined by Δη. As the rotor incidence 

changes at off-design flow rates, the rotor efficiency is 

determined by applying the loss model to the input value of 

rotor efficiency. 

All-Axial Compressor 

The blade tip maximum speed limit was 1500 ft/sec and the 

rotor-tip diameters were tapered or reduced in each successive 

stage. The flow rate of 28.4 lbm/sec, the maximum tip speed 

criteria and the hub-to-tip ratio of 0.36 was used to size the 

first-stage axial compressor, resulting in a design speed of 

27 289 rotations per minute. To meet the overall pressure ratio 

requirement of 31:1, the flow path and each blade row were 

designed concurrently in an iterative process using the meanline code. This process determined the pressure ratio that can 

be achieved in each stage. The number of stages that would be 

required to meet the overall pressure ratio was determined by 

an iterative process. The maximum tip speed of 1500 ft/sec 

was considered to be a structurally acceptable limit, as 

advanced blade and disk materials could support these speeds. 

The maximum tip speed limit was also imposed because of 

concerns about high rotor-inlet tip relative Mach number 
NASA/TM—2009-215641 3 

particularly in the first transonic stage. The taper of the tip 

flow path was kept at a rate such that the pressure ratio 

requirements could be met with eight stages without the 

addition of a ninth stage. Pressure ratios of this magnitude 

result in high gas temperatures at the compressor exit. The 

tapered-tip flow path axial-compressor configuration is 

favorable from a structural perspective, as the latter stages that 

experience the highest temperatures also operate at reduced tip 

speeds in comparison to the front-block compressor stages. 

The hub diameter was allowed to vary to keep the absolute 

Mach number near 0.50 at the rotor inlet, and within the limit 

of rotor diffusion factor at the design point. The flow path and 

the rotor designs were arrived at iteratively while keeping 

within limits of rotor diffusion factor and rotor-inlet Mach 

number. 

The resulting flow path for the eight-stage axial compressor 

with tapered-tip diameter flow path is shown in Figure 2. The 

blade span of the last stage for this design is 0.281 in. at a tip 

diameter of 11.53 in. The rotor-tip clearance will be on the 

order of 0.011 in., resulting in the ratio of clearance to span in 

the first stage of 0.3 percent and the last stage having a 

clearance-to-span ratio of 4 percent. 

Table I shows the results obtained with the mean-line 

compressor flow analysis and design code for the eight-stage 

axial compressor with tapered-tip flow path. The pressure ratio 

per stage tapers off in each subsequent stage, but the work 

split between the eight stages resulted in a relatively even 

distribution of power of near 1350 hp per stage, as shown in 

Table I. 

Except for the tip speed and tip relative Mach number, all 

other rotor parameters in the table are calculated at the rootmean-square diameter. The rotor-inlet root-mean-square 

radius of each successive rotor increases from 4.72 in. in the 

first rotor inlet through the first five stages and then gradually 

becomes constant at 5.71 in. for the last three stages. The first 

five stages are transonic at the tip. The absolute Mach number 

at the inlet of each rotor was on the order of 0.50. The relative 

flow angle at each rotor inlet is on the order of 62° from the 

axial direction. The rotor exit blade angles are on the order of 

40° to 46° for all rotors. The resulting peak overall adiabatic 

efficiency at 28.4 lbm/sec is 80.3 percent. The power required 

to drive this compressor is estimated to be 10 781 hp. The 

flow path coordinates of the rotor tip and hub radial and axial 

positions are listed in Appendix A. Note that the flow path is 

based on a mean-line conceptual flow analysis, and may 

change as the design and analysis progress to higher levels of 

fidelity. The design point performance for each stage is listed 

in Appendixes C and D. As a follow on to this study, it is 

planned to generate preliminary blade shapes with a 

turbomachinery design code that includes the losses due to tip 

leakages. Structural conceptual design, material selection, and 

analyses of the blades and disks is also planned, as the exit 

temperature of 1568 °R is high, and proper material selection 

will determine the viability of this compressor. 
NASA/TM—2009-215641 4 

TABLE I.—SUMMARY OF THE EIGHT-STAGE ALL-AXIAL COMPRESSOR 

WITH TAPERED-TIP FLOW PATH 

 1 2 3 4 5 6 7 8 

Rotor inlet 

Flow rate, corrected, 

lbm/sec 

28.40 15.65 9.61 6.52 4.59 3.38 2.57 1.99

Mach absolute 0.53 0.54 0.47 0.45 0.47 0.44 0.40 0.36

Relative Mach at tip 1.50 1.34 1.19 1.10 1.05 0.98 0.92 0.86

Tip speed (U), ft/sec 1495 1475 1452 1434 1419 1407 1399 1392

Relative flow angle, deg 62.7 62.9 64.3 63.8 61.7 61.9 62.8 64.0

Blade angle, deg 57.6 56.7 58.4 57.5 55.6 55.4 56.2 57.6

Rotor exit

Blade angle, deg 40.5 43.5 46.0 44.2 41.0 41.7 42.3 43.8

Absolute flow angle, deg 47.8 48.2 45.7 42.3 41.1 40.8 42.4 43.2

Flow deviation, deg 4.5 4.4 4.0 3.9 4.0 4.0 4.0 4.0

Diffusion factor 0.50 0.55 0.52 0.49 0.50 0.49 0.51 0.51

Relative velocity ratio 1.95 1.85 1.69 1.61 1.55 1.55 1.57 1.57

Exit temperature, °R 657.9 794.0 922.1 1054.0 1184.0 1313.0 1442.0 1568.0

Stage

Pressure ratio 2.056 1.789 1.587 1.519 1.440 1.387 1.354 1.314

Temperature ratio 1.268 1.207 1.161 1.143 1.123 1.109 1.099 1.087

Efficiency, adiabatic 85.5 87.6 87.2 87.5 87.1 87.3 87.5 87.6

Work coefficient, φ 0.388 0.395 0.385 0.410 0.416 0.423 0.436 0.435

Horsepower 1391 1364 1290 1343 1336 1339 1367 1350

 

All-Axial Compressor 

Off-Design Operation 

The mean-line methodology was used to perform an offdesign analysis of this all-axial compressor at the 100 percent 

speed line. The criteria that were used for predicting surge in 

the analysis are a maximum value of a diffusion factor of 0.60. 

The pressure ratio versus flow is shown in Figure 3, while the 

adiabatic efficiency versus flow rate is shown in Figure 4. 

Figure 3 illustrates the steep pressure rise characteristic as the 

flow is reduced from the design value of 28.4 lbm/sec. The 

off-design characteristics at a range of part speed operating 

conditions will be studied in more detail in the future, as it is 

important to understand the operability of this high-pressure 

compressor. In addition to the variable inlet guide vanes 

(IGV), there will likely be a need to have several stator vanes 

variable as well, to aerodynamically match the stages at part 

speed operating conditions with acceptable levels of rotor 

incidence. 
NASA/TM—2009-215641 5 

Since the last stage blade span of 0.281 in. is low in 

comparison to traditional designs, a steeper tip-taper rate will 

also be studied in the future in order to increase the span of the 

last stage, possibly requiring additional axial stages. 

Axial-Centrifugal Compressor 

The second configuration that was studied focused on an 

axial-centrifugal compressor, since traditionally, rotary 

winged aircraft engines typically feature axial-centrifugal 

compressors. This design utilized a first-stage axial 

compressor that is close to the first stage of the previous allaxial case shown in Figure 2, but the tip flow-path taper 

through the subsequent axial stages is even steeper. The 

design shaft speed is 27 289 rotations per minute. This 

conceptual design iteration focused on adding one centrifugal 

stage to the back end of the compressor to take the place of 

several axial stages. In addition, the taper of the tip flow path 

was increased to make the rotor blade spans as large as 

possible. This further reduced the tip clearance-to-span ratio of 

the axial rotors, as well as their hub-to-tip ratio. This reduction 

was considered to be particularly important for the centrifugal 

impeller, as its efficiency can be negatively influenced by a 

high inlet hub-to-tip ratio. The increased rate of tip flow-path 

taper through the axial stages accommodates the transition to 

the centrifugal impeller with a hub-to-tip ratio as low as 

possible. 

As the tapered-tip flow path already provided this transition 

to the centrifugal impeller, there was no need to use a 

transition duct after the last axial stage to further reduce the 

inlet hub-to-tip ratio. This was done in an effort to reduce 

pressure losses normally experienced in the “goose neck” of 

traditional axial-centrifugal compressors. 

The centrifugal compressor can effectively produce enough 

pressure ratio to take the place of several rear-block axial 

stages. However, there are limitations on the impeller pressure 

ratio such as specific speed (Ref. 7), which is a normalized 

aerodynamic parameter that can be used to estimate the flow 

and pressure ratio where the centrifugal compressor will 

operate most efficiently. Specific speed NS is defined by the 

following Equation (5). 

4

3

2

1

H

N Q NS

Δ





 





= (5) 

Q is the volumetric flow rate in cubic feet per minute, while 

ΔH is the enthalpy rise through the rotor, and N is the shaft 

speed in rotations per minute. The range of centrifugal 

impeller specific speed that typically has the highest potential 

level of efficiency is 80 to 90. If the centrifugal is not designed 

to be in this range of specific speed, the maximum attainable 

efficiency will be limited. Another limitation of centrifugal 

compressors is the structural and material limitation of tip 

speed at high operating temperatures. 

A mechanical design study for the LCTR engine has not yet 

been done, as it is currently only in the study phase. 

Rotordynamics analyses would be done on the engine system 

after the complete shaft assembly, including the turbines, 

bearings, and seals, has been sized. Likely, there will be 

mechanical and rotordynamics considerations that will 

influence the evolution of the final compressor flow path. The 

cross section for the compressor that resulted from this 

conceptual design is shown in Figure 5. 
NASA/TM—2009-215641 6 

The wall curvature of the flow path leading into the 

centrifugal stage has been sized to reduce curvature effects on 

local velocity, which has an effect on the pressure losses 

encountered in the duct. The flow path coordinates of the rotor 

tip and hub radial and axial positions are listed in Appendix A. 

Note that these coordinates are based on a mean-line 

conceptual flow analysis, and may change as the design and 

analysis progress to higher levels of fidelity. In addition, 

subsequent structural analyses may indicate that the limits on 

maximum tip speed may be different from 1500 ft/sec used in 

this conceptual design. The design point performance for each 

stage is listed in Appendixes C and D. 

The work split among the seven axial stages was tailored to 

be as even as possible and within the limits on rotor diffusion 

factor. The specific speed of the centrifugal compressor stage 

is 56 and is lower than the level considered to be near the 

optimum value for maximum efficiency potential. To be closer 

to what is considered to be the optimum specific speed, more 

of the work would have to be done by the axial stages, or more 

stages would be required, and the centrifugal compressor 

pressure ratio would need to be reduced (lower impeller exit 

diameter, and/or more back swept blade angle). Another way 

to increase the specific speed of the centrifugal stage is to have 

a two-spool compressor, as suggested in Reference 2. The 

second spool would rotate at a faster speed, allowing more 

flexibility to optimize the specific speed of the centrifugal 

stage. The two-spool compressor will be studied in more detail 

in the future. 

Since the tip flow path taper is steeper than the previous allaxial compressor case, the blade spans for the rear block axial 

stages are larger, culminating with a 0.61 in. blade span of the 

seventh (last axial) stage rotor. This significant increase in 

blade span from the previous all-axial case was caused by two 

factors (1) the reduced rotor-tip diameters resulted in reduced 

pressure ratio per stage from the previous case and (2) as the 

rotor-tip diameter was reduced, the annular area in the flow 

path required a larger span to accommodate the increased 

volume flow. Even with a 0.62-in. impeller inlet span that was 

the result of reduced tip flow path through the axial stages, 

this design produced a high impeller inlet hub-to-tip ratio of 

0.85. The centrifugal impeller exit height is 0.280 in., and the 

exit blade angle has a 20° back sweep from the radial 

direction. The tip clearance of this impeller will likely vary 

from inlet to exit. Table II summarizes the key stage-by-stage 

mean-line performance parameters for the seven-stage axial 

compressor followed by the centrifugal stage. 

A reasonable axial running clearance of this impeller at the 

exit tip is on the order of 0.005 in., or near 2 percent of the 

exit span. It will be necessary to maintain tight axial 

clearances to minimize tip leakages and prevent a reduction in 

impeller efficiency. 

TABLE II.—THE AXIAL-CENTRIFUGAL COMPRESSOR DESIGN POINT STAGE-BY-STAGE 

AERODYNAMIC PERFORMANCE 

 1 2 3 4 5 6 7 8 Centrifugal 

Rotor inlet 

Flow rate, corrected, 

lbm/sec 

28.4 15.81 9.86 6.82 5.05 3.92 3.18 2.65

Mach absolute 0.53 0.50 0.49 0.44 0.41 0.40 0.37 0.29

Relative Mach at tip 1.50 1.31 1.16 1.03 0.93 0.85 0.78 0.65

Tip speed (U), ft/sec 1495 1452 1379 1319 1248 1169 1106 1595 (exit)

Relative flow angle, deg 62.7 63.6 62.0 62.4 61.6 59.7 59.2 60.6

Blade angle 57.6 57.4 55.7 56.2 55.2 53.2 53.0 54.5

Rotor exit 

Blade angle, deg 40.1 41.5 42.0 40.5 38.0 36.0 34.0 20.0

Absolute flow angle, deg 50.1 47.5 44.2 42.5 43.1 40.2 41.1 63.8

Flow deviation, deg 4.8 4.5 4.2 4.2 4.4 4.2 4.4 18.2

Diffusion factor 0.43 0.55 0.53 0.52 0.53 0.51 0.52 0.74

Relative velocity ratio 2.00 1.88 1.77 1.69 1.70 1.65 1.67 1.54

Exit temperature, °R 654.9 785.7 906.0 1017. 1119. 1212. 1298. 1577.

Stage 

Pressure ratio 2.031 1.755 1.553 1.431 1.351 1.284 1.241 1.818

Temperature ratio 1.263 1.200 1.153 1.123 1.101 1.083 1.071 1.215

Efficiency, adiabatic 85.8 87.5 87.4 87.2 87.3 87.1 86.8 82.8

Work coefficient, φ 0.379 0.403 0.408 0.418 0.438 0.450 0.470 0.718

Horsepower 1361 1296 1211 1126 1053 957 897 2935

 
NASA/TM—2009-215641 7 

The reduced tip radii of each subsequent axial rotor are also 

good from a structural perspective, since the operating 

temperatures of each subsequent axial rotor are higher from 

the previous stage, and a reduced blade tip speed in the latter 

rotors can result in a more structurally acceptable design. The 

material selection and structural design of the centrifugal 

impeller needs careful consideration as its tip speed is 

1595 ft/sec at an exit temperature of 1565 °R. The feasibility 

of this rotor needs to be verified with structural and thermal 

analyses. 

The resulting overall compressor efficiency at 28.4 lb/sec is 

79.6 percent adiabatic. The power required to drive it is 

estimated to be 10 850 hp. The results from Tables I and II are 

shown in Appendix B as plots comparing the values obtained 

from the all-axial compressor study and the axial-centrifugal 

compressor. 

As shown in Table II, the diffusion factor of the axial rotors 

at the design point ranges between 0.43 and 0.55. These values 

indicate that there is a surge margin available. The centrifugal 

impeller design point relative velocity ratio from inlet to exit 

is 1.54. The value of diffusion factor for the centrifugal 

impeller is higher at the design point (0.74) than the axial rotor 

diffusion factors, but this is not unexpected as centrifugal 

impellers are typically more highly loaded than axial blades. 

Structural conceptual design, material selection, and analyses 

of the blades and disks are planned to determine whether this 

design is feasible, as the exit temperature of 1577 °R and 

impeller tip speed of 1595 ft/sec may be challenging with 

current material capabilities. The development of a material that 

can support these high tip speeds at high temperatures will 

determine the viability of this compressor. 

Axial-Centrifugal Off-Design 

The mean-line methodology of Reference 3 was used to 

perform an off-design analysis of this axial-centrifugal 

compressor at the 100 percent speed line. The criteria that 

were used for modeling the onset of surge were a maximum 

value of rotor diffusion factor of 0.60 for the axial rotors and 

relative velocity ratio of 1.95. Figure 6 and 7 show the 

compressor performance along the 100 percent speed line. As 

the flow rate is reduced to 27.7 lbm/sec, the maximum relative 

velocity ratio that is experienced in the axial rotors is 1.9 and 

1.77 in the centrifugal impeller. Based on the relative velocity 

criteria, it appears that at the 100 percent speed, stall will be 

initiated in the axial rotors and not in the centrifugal impeller. 

However, the diffusion factor limit for the centrifugal 

compressors at surge needs further validation. 

Further analyses of this compressor are planned to determine 

the variable geometry schedule that will be necessary to operate 

it with an acceptable surge margin at part speed. 

The pressure ratio versus flow rate of the axial-centrifugal 

compressor is illustrated in Figure 6. 

As illustrated in Figure 6, the pressure rise characteristic of 

this compressor is shallow as the flow is reduced from the 

design point value. The adiabatic efficiency versus flow rate is 

shown in Figure 7. The axial-centrifugal compressor pressure 

rise characteristic was compared to the all-axial compressor 

pressure rise characteristic in Figure 8. The axial-centrifugal 

compressor has less pressure rise as the flow is reduced, than 

does the all-axial compressor case. The axial-centrifugal 

compressor appears to have higher flow margin before surge is 

encountered, in comparison to the all-axial case. The reduced 

pressure rise to surge of the axial-centrifugal configuration 

and the additional flow margin are likely due to the centrifugal 

compressor. Figure 9 illustrates the efficiency characteristics 

of the all-axial versus the axial-centrifugal compressor. The 

peak efficiency of both compressors is at a flow rate of 

28.3 lbm/sec. 
NASA/TM—2009-215641 8 

 

Conclusions 

A conceptual design study was made on two potential 

compressors to meet the Large Civil Tilt Rotorcraft (LCTR) 

engine pressure ratio and flow requirements. An all-axial 

compressor was sized with a tapered-tip flow path, resulting in 

the last-stage rotor span of 0.281 in., and a tip-to-span 

clearance ratio of 4 percent. The hub-to-tip ratio of the last 

rotor is high at 0.95. Further study of the all-axial 

configuration is planned to increase the blade span of the last 

stage by increasing the tip taper and likely requiring additional 

axial stages. A second compressor was studied, featuring an 

axial-centrifugal compressor configuration with seven axial 

stages followed by a centrifugal stage. The centrifugal 

compressor has a specific speed of 56, which is low even for 

centrifugals. The impeller inlet also has a high hub-to-tip ratio 

of 0.85 and an exit height of 0.28 in. The high hub-to-tip ratio 

and low exit height will require tight control of eccentricity 

and axial clearance for this compressor configuration to keep 

rotor-tip leakages at acceptable levels. It is planned to do a 

conceptual design study of a two-spool compressor in the 

future that can provide additional flexibility to increase the 

specific speed of the centrifugal stage. The result of this study 

shows that from an aerodynamic perspective, it may be 

feasible to meet the requirements of a 31:1 pressure ratio 

compressor for the LCTR engine with either the all-axial 

compressor with a tapered-tip flow path, or with an axialcentrifugal compressor featuring seven axial stages, followed 

by a centrifugal stage. The high rotor-tip speeds of the latter 

compressor stages operating at temperatures up to 1580 °R 

will require advanced alloy materials be selected that are 

typically used to make turbines. The selection of materials for 

these high-tip-speed rotors operating in a high-temperature 

environment is a technology that will enable the feasibility of 

these compressors. Further study is planned with higher 

fidelity turbomachinery design and flow analysis codes to 

estimate the compressor efficiencies with clearance effects. 

Preliminary structural and thermal analyses of the blades and 

vanes are needed to determine if the required tip speeds from 

this conceptual design study can be achieved with currently 

available advanced alloy rotor materials. The mechanical 

design of the compressor and turbine shafts will also influence 

the feasibility of the flow paths in this study. 
NASA/TM—2009-215641 9 

Appendix A.—Compressor Coordinates 

The coordinates listed below describe the hub and tip flow path contour for both the eight-stage all-axial compressor and the 

seven-stage axial-centrifugal compressor. LE is the blade or vane leading edge and TE is the blade or vane trailing edge.

Eight-stage all-axial compressor

 Tip Hub 

X R X R 

Rotor 1 LE 0.2218 6.2778 0.0000 2.2474 

Rotor 1 TE 1.7847 6.2778 1.9957 3.8000 

Stator 1 LE 2.1728 6.2634 2.1728 3.9049 

Stator 1 TE 3.4517 6.2120 3.4517 4.3956 

Rotor 2 LE 3.7025 6.2009 3.6033 4.4439 

Rotor 2 TE 4.4893 6.1646 4.5798 4.7246 

Stator 2 LE 4.7238 6.1534 4.7238 4.7615 

Stator 2 TE 5.4735 6.1162 5.4735 4.9363 

Rotor 3 LE 5.6755 6.1059 5.6149 4.9659 

Rotor 3 TE 6.2246 6.0770 6.2972 5.0945 

Stator 3 LE 6.3877 6.0682 6.3877 5.1098 

Stator 3 TE 6.9196 6.0390 6.9196 5.2288 

Rotor 4 LE 7.0800 6.0300 7.0314 5.2503 

Rotor 4 TE 7.5312 6.0080 7.5753 5.3435 

Stator 4 LE 7.6681 6.0001 7.6681 5.3576 

Stator 4 TE 8.0874 5.9756 8.0874 5.4143 

Rotor 5 LE 8.2775 5.9643 8.2373 5.4320 

Rotor 5 TE 8.6236 5.9434 8.6770 5.4674 

Stator 5 LE 8.7729 5.9343 8.7729 5.4744 

Stator 5 TE 9.0998 5.9140 9.0998 5.5051 

Rotor 6 LE 9.2801 5.9146 9.2282 5.5051 

Rotor 6 TE 9.5195 5.9041 9.5425 5.5234 

Stator 6 LE 9.6297 5.8993 9.6297 5.5280 

Stator 6 TE 9.8973 5.8875 9.8973 5.5411 

Rotor 7 LE 10.0574 5.8805 10.0159 5.5464 

Rotor 7 TE 10.2404 5.8725 10.2656 5.5564 

Stator 7 LE 10.3702 5.8668 10.3702 5.5601 

Stator 7 TE 10.6039 5.8565 10.6039 5.5675 

Rotor 8 LE 10.7352 5.8507 10.7071 5.5704 

Rotor 8 TE 10.8959 5.8437 10.9208 5.5755 

Stator 8 LE 11.0220 5.8381 11.0220 5.5776 

Stator 8 TE 11.1900 5.8308 11.1900 5.5804 

Seven-stage axial-1 centrifugal compressor

 Tip Hub 

X R X R 

Rotor 1 LE 0.2218 6.2778 0.0000 2.2474 

Rotor 1 TE 1.7858 6.2647 2.0067 3.6718 

Stator 1 LE 2.1728 6.2483 2.1728 3.7529 

Stator 1 TE 3.4517 6.1508 3.4517 4.1597 

Rotor 2 LE 3.7072 6.1313 3.6019 4.1848 

Rotor 2 TE 4.4767 6.0112 4.5760 4.4313 

Stator 2 LE 4.7146 5.9741 4.7146 4.4598 

Stator 2 TE 5.4592 5.8578 5.4592 4.5707 

Rotor 3 LE 5.6627 5.8261 5.5970 4.5902 

Rotor 3 TE 6.1991 5.7423 6.2887 4.6127 

Stator 3 LE 6.3668 5.7161 6.3788 4.6141 

Stator 3 TE 6.8933 5.6171 6.8933 4.6103 

Rotor 4 LE 7.0470 5.5695 6.9848 4.6075 

Rotor 4 TE 7.4863 5.4694 7.5454 4.5781 

Stator 4 LE 7.6303 5.4339 7.6303 4.5717 

Stator 4 TE 8.0420 5.3223 8.0420 4.5016 

Rotor 5 LE 8.2207 5.2718 8.1611 4.4815 

Rotor 5 TE 8.5639 5.1714 8.6409 4.3913 

Stator 5 LE 8.7134 5.1262 8.7134 4.3791 

Stator 5 TE 9.1205 4.9985 9.1205 4.2940 

Rotor 6 LE 9.2980 4.9408 9.2135 4.2734 

Rotor 6 TE 9.5245 4.8652 9.5654 4.1882 

Stator 6 LE 9.6303 4.8291 9.6303 4.1756 

Stator 6 TE 9.9066 4.7329 9.9066 4.0952 

Rotor 7 LE 10.0604 4.6780 9.9850 4.0714 

Rotor 7 TE 10.2467 4.6101 10.2979 3.9663 

Stator 7 LE 10.3712 4.5639 10.3712 3.9402 

Stator 7 TE 10.6099 4.4735 10.6099 3.8536 

Rotor 8 LE 12.0406 4.1509 11.9423 3.5313 

Rotor 8 TE 13.3540 6.7000 13.6340 6.7000 

Stator 8 LE 13.3722 7.2814 13.6340 7.2814 

Stator 8 TE 13.3722 9.4314 13.6340 9.4314 

NASA/TM—2009-215641 11 

Appendix B.—Plots of Compressor Parameters at the Design Point 

Figures 10 to 23 present plots comparing the values obtained from the all-axial compressor study and the axial-centrifugal 

compressor (see Tables I and II in text). 
NASA/TM—2009-215641 12 
NASA/TM—2009-215641 13 
NASA/TM—2009-215641 14 
NASA/TM—2009-215641 15 
NASA/TM—2009-215641 16 
NASA/TM—2009-215641 17 
 
NASA/TM—2009-215641 19 

Appendix C.—All-Axial Compressor Data 

Below are the all-axial compressor mean-line flow analysis results at the design point operating condition. 

 LCTR 8 Axial Compressor Stage 100% Speed 

 COMPRESSOR INLET CONDITIONS, STAGE 1 

RESET = 0.000 BLEED = 0.000 DP Inc 5.000 Kg/sec = 12.90425 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.39 27289.000 14.613 518.670 1.100 0.800 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 28.39 27289.789 1.402 0.249 53.349 28.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 1 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 6.28 0.00 -0.02 636.07 -0.22 636.07 0.59 

 MEAN 4.72 0.00 -0.02 578.24 -0.20 578.24 0.53 

 HUB 2.25 0.00 -0.02 462.60 -0.16 462.60 0.42 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 66.96 62.00 4.96 1495.53 1625.38 11.58 485.21 1.50 

 MEAN 62.76 57.66 5.10 1123.15 1263.44 12.05 490.74 1.16 

 HUB 49.17 44.20 4.97 535.10 707.46 12.95 500.97 0.64 

 

 ROTOR EXIT CONDITIONS, STAGE 1 

 

 B2 axial THK AeroBl 

 0.30 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 6.28 890.20 702.10 547.27 1222.90 0.73 

 MEAN 5.19 874.11 647.99 586.67 1185.39 0.74 

 HUB 3.80 1082.43 829.37 695.56 1143.21 0.95 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1495.53 963.86 793.43 0.79 

 MEAN 1236.03 830.64 588.04 0.70 1.96 

 HUB 904.94 699.65 75.57 0.61 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 33.52 2.29 23.55 687.41 1.33 621.24 0.83 

 MEAN 29.62 2.03 20.63 647.37 1.25 583.59 0.90 

 HUB 28.63 1.96 16.06 639.26 1.23 541.65 0.91 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 52.06 55.40 51.50 3.90 0.93 0.57 1.32 

 MEAN 47.84 45.07 40.50 4.57 0.93 0.50 1.75 

 HUB 50.01 6.20 1.00 5.20 0.93 0.21 3.68 

 

 STAGE EXIT CONDITIONS, STAGE 1 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.855 30.065 2.057 1.269 267.791 2.076 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 867478.44 0.388 1391.703 1191539.375 

 

 

 
NASA/TM—2009-215641 20 

 COMPRESSOR INLET CONDITIONS, STAGE 2 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.39 27289.000 30.065 658.012 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 15.63 24228.629 1.401 0.249 53.349 38.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 2 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 6.20 0.00 -0.02 720.54 -0.25 720.54 0.59 

 MEAN 5.39 0.00 -0.02 655.04 -0.23 655.04 0.54 

 HUB 4.44 0.00 -0.02 589.54 -0.20 589.54 0.48 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 63.97 57.82 6.15 1475.29 1642.07 23.76 615.16 1.34 

 MEAN 62.96 56.70 6.26 1283.25 1440.97 24.73 622.24 1.18 

 HUB 60.85 54.60 6.25 1056.87 1210.36 25.73 629.32 0.98 

 

 ROTOR EXIT CONDITIONS, STAGE 2 

 

 B2 axial THK AeroBl 

 0.20 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 6.16 885.95 663.11 587.53 1341.71 0.66 

 MEAN 5.49 881.34 657.35 587.06 1325.89 0.66 

 HUB 4.72 900.82 641.19 632.74 1303.07 0.69 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1466.71 995.48 803.60 0.74 

 MEAN 1306.45 875.19 649.09 0.66 1.85 

 HUB 1123.55 795.64 482.36 0.61 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 56.83 1.89 42.40 813.88 1.24 748.46 0.84 

 MEAN 55.75 1.85 41.44 795.64 1.21 730.90 0.92 

 HUB 51.09 1.70 37.11 773.46 1.18 705.83 0.93 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 48.46 53.83 50.00 3.83 0.93 0.55 1.27 

 MEAN 48.23 47.87 43.50 4.37 0.93 0.55 1.46 

 HUB 45.38 37.32 32.50 4.82 0.93 0.50 1.77 

 

 STAGE EXIT CONDITIONS, STAGE 2 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.876 53.845 1.791 1.207 213.512 1.655 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 851000.19 0.396 1365.266 794049.563 

 

 
NASA/TM—2009-215641 21 

 COMPRESSOR INLET CONDITIONS, STAGE 3 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.39 27289.000 53.845 794.328 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 9.59 22051.896 1.398 0.251 53.349 50.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 3 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 6.10 0.00 -0.02 699.40 -0.24 699.40 0.52 

 MEAN 5.56 0.00 -0.02 635.82 -0.22 635.82 0.47 

 HUB 4.96 0.00 -0.02 572.24 -0.20 572.24 0.42 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 64.29 58.20 6.09 1452.66 1612.48 44.87 754.12 1.19 

 MEAN 64.35 58.40 5.95 1323.90 1468.86 46.28 760.77 1.09 

 HUB 64.16 58.40 5.76 1181.18 1312.67 47.71 767.41 0.96 

 

 ROTOR EXIT CONDITIONS, STAGE 3 

 

 B2 axial THK AeroBl 

 0.10 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 6.07 873.06 624.63 609.97 1449.37 0.60 

 MEAN 5.60 861.81 618.16 600.50 1440.31 0.60 

 HUB 5.09 848.71 568.87 629.84 1423.84 0.60 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1445.76 1022.89 821.12 0.71 

 MEAN 1333.91 934.29 715.75 0.65 1.69 

 HUB 1211.78 900.01 642.91 0.63 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 88.57 1.64 69.33 938.26 1.18 875.00 0.84 

 MEAN 88.57 1.64 69.55 925.75 1.17 864.11 0.92 

 HUB 82.37 1.53 64.79 904.19 1.14 844.42 0.93 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 45.68 53.39 49.70 3.69 0.93 0.52 1.22 

 MEAN 45.83 50.00 46.00 4.00 0.93 0.52 1.34 

 HUB 42.09 45.59 41.50 4.09 0.93 0.46 1.51 

 

 STAGE EXIT CONDITIONS, STAGE 3 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.873 85.573 1.589 1.162 179.754 1.393 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 806058.38 0.386 1293.166 650593.000 

 

 

 
NASA/TM—2009-215641 22 

 COMPRESSOR INLET CONDITIONS, STAGE 4 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.39 27289.000 85.573 922.735 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 6.50 20460.078 1.395 0.253 53.349 60.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 4 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 6.02 0.00 -0.02 724.66 -0.25 724.66 0.50 

 MEAN 5.65 0.00 -0.02 658.78 -0.23 658.78 0.45 

 HUB 5.24 0.00 -0.02 592.91 -0.20 592.91 0.40 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 63.20 57.00 6.20 1434.56 1607.43 72.33 879.86 1.10 

 MEAN 63.91 57.50 6.41 1344.94 1497.83 74.41 886.96 1.03 

 HUB 64.61 58.00 6.61 1248.91 1382.69 76.53 894.04 0.94 

 

 ROTOR EXIT CONDITIONS, STAGE 4 

 

 B2 axial THK AeroBl 

 0.10 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 6.00 926.43 632.40 677.01 1543.43 0.60 

 MEAN 5.68 902.12 610.45 664.20 1536.64 0.59 

 HUB 5.34 916.89 621.65 673.98 1530.54 0.60 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1429.32 1045.68 796.93 0.68 

 MEAN 1352.53 995.92 742.08 0.65 1.62 

 HUB 1271.10 935.97 649.46 0.61 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 132.14 1.54 103.67 1065.67 1.15 994.92 0.84 

 MEAN 132.17 1.54 104.75 1053.29 1.14 986.21 0.93 

 HUB 130.41 1.52 102.40 1047.67 1.14 978.38 0.94 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 43.05 49.65 45.80 3.85 0.93 0.51 1.19 

 MEAN 42.59 48.17 44.24 3.93 0.93 0.50 1.27 

 HUB 42.69 43.94 39.70 4.24 0.93 0.49 1.37 

 

 STAGE EXIT CONDITIONS, STAGE 4 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.876 130.176 1.521 1.144 148.354 1.150 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 840325.50 0.411 1348.141 567954.750 

 

 

 
NASA/TM—2009-215641 23 

 COMPRESSOR INLET CONDITIONS, STAGE 5 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.39 27289.000 130.176 1055.544 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 4.57 19129.678 1.389 0.256 53.349 70.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 5 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 5.96 0.00 -0.02 798.75 -0.28 798.75 0.51 

 MEAN 5.70 0.00 -0.02 726.14 -0.25 726.14 0.47 

 HUB 5.43 0.00 -0.02 653.52 -0.23 653.52 0.42 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 60.63 54.10 6.53 1418.85 1628.47 108.87 1003.97 1.05 

 MEAN 61.85 55.60 6.25 1356.98 1539.27 112.20 1012.50 0.99 

 HUB 63.18 56.80 6.38 1292.16 1448.22 115.61 1021.01 0.92 

 

 ROTOR EXIT CONDITIONS, STAGE 5 

 

 B2 axial THK AeroBl 

 0.05 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 5.94 983.70 648.18 739.95 1634.40 0.60 

 MEAN 5.70 961.63 635.91 721.35 1630.91 0.59 

 HUB 5.46 931.93 562.69 742.88 1619.14 0.58 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1414.08 1064.96 765.90 0.65 

 MEAN 1358.52 1021.04 722.61 0.63 1.56 

 HUB 1300.59 1047.07 737.90 0.65 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 191.71 1.47 150.33 1198.90 1.14 1119.92 0.84 

 MEAN 193.99 1.49 153.56 1190.66 1.13 1115.18 0.92 

 HUB 183.68 1.41 146.96 1170.00 1.11 1099.12 0.93 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 41.22 45.99 42.00 3.99 0.93 0.51 1.21 

 MEAN 41.40 45.05 41.00 4.05 0.93 0.50 1.27 

 HUB 37.14 44.81 41.00 3.81 0.93 0.42 1.33 

 

 STAGE EXIT CONDITIONS, STAGE 5 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.872 187.818 1.443 1.124 129.361 1.003 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 837881.94 0.419 1344.221 517435.875 

 

 
NASA/TM—2009-215641 24 

 COMPRESSOR INLET CONDITIONS, STAGE 6 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.39 27289.000 187.818 1186.523 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 3.36 18042.957 1.383 0.259 53.349 80.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 6 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 5.91 0.00 -0.02 793.95 -0.27 793.95 0.48 

 MEAN 5.71 0.00 -0.02 721.77 -0.25 721.77 0.44 

 HUB 5.50 0.00 -0.02 649.60 -0.22 649.60 0.39 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 60.57 53.90 6.67 1407.11 1615.88 160.59 1136.17 0.98 

 MEAN 62.03 55.43 6.60 1359.21 1539.18 164.89 1144.50 0.93 

 HUB 63.62 57.10 6.52 1309.56 1462.03 169.25 1152.81 0.88 

 

 ROTOR EXIT CONDITIONS, STAGE 6 

 

 B2 axial THK AeroBl 

 0.05 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 5.90 981.22 654.68 730.87 1723.20 0.57 

 MEAN 5.71 951.16 626.87 715.37 1719.28 0.55 

 HUB 5.52 926.24 571.24 729.11 1711.00 0.54 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1404.58 1047.15 749.90 0.61 

 MEAN 1360.01 1024.33 733.14 0.60 1.56 

 HUB 1313.92 1040.76 742.68 0.61 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 265.99 1.42 214.00 1328.46 1.12 1250.81 0.85 

 MEAN 267.36 1.42 217.66 1318.11 1.11 1245.14 0.93 

 HUB 257.59 1.37 211.50 1302.37 1.10 1233.18 0.94 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 41.85 45.74 41.70 4.04 0.93 0.52 1.21 

 MEAN 41.23 45.70 41.70 4.00 0.93 0.50 1.25 

 HUB 38.08 45.53 41.70 3.83 0.93 0.44 1.30 

 

 STAGE EXIT CONDITIONS, STAGE 6 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.874 261.184 1.391 1.109 113.106 0.877 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 841334.13 0.426 1349.759 468731.500 

 

 
NASA/TM—2009-215641 25 

 COMPRESSOR INLET CONDITIONS, STAGE 7 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.39 27289.000 261.184 1316.314 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 2.55 17130.342 1.376 0.263 53.349 82.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 7 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 5.87 0.00 -0.02 761.01 -0.26 761.01 0.44 

 MEAN 5.71 0.00 -0.02 691.83 -0.24 691.83 0.40 

 HUB 5.54 0.00 -0.02 622.65 -0.21 622.65 0.36 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 61.46 54.80 6.66 1398.96 1592.79 229.53 1270.68 0.92 

 MEAN 63.04 56.20 6.84 1359.76 1525.85 234.57 1278.23 0.88 

 HUB 64.74 57.90 6.84 1319.40 1459.13 239.67 1285.76 0.83 

 

 ROTOR EXIT CONDITIONS, STAGE 7 

 

 B2 axial THK AeroBl 

 0.05 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 5.87 963.51 659.92 702.04 1807.01 0.53 

 MEAN 5.71 938.35 639.43 686.75 1804.28 0.52 

 HUB 5.55 915.64 595.68 695.39 1798.28 0.51 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1397.06 1017.95 737.13 0.56 

 MEAN 1359.94 995.37 720.51 0.55 1.58 

 HUB 1321.78 1005.38 726.10 0.56 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 358.21 1.37 296.07 1456.51 1.11 1382.65 0.85 

 MEAN 361.80 1.39 301.75 1448.54 1.10 1378.49 0.93 

 HUB 352.33 1.35 296.03 1436.04 1.09 1369.33 0.94 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 43.23 46.40 42.30 4.10 0.93 0.53 1.22 

 MEAN 42.96 46.37 42.30 4.07 0.93 0.51 1.26 

 HUB 40.58 46.24 42.30 3.94 0.93 0.47 1.30 

 

 STAGE EXIT CONDITIONS, STAGE 7 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.876 354.474 1.357 1.099 98.583 0.764 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 860064.69 0.441 1379.809 427515.625 

 
NASA/TM—2009-215641 26 

 COMPRESSOR INLET CONDITIONS, STAGE 8 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.39 27289.000 354.474 1447.031 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 1.97 16338.298 1.368 0.267 53.349 84.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 8 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 5.84 0.00 -0.02 721.08 -0.25 721.08 0.39 

 MEAN 5.71 0.00 -0.02 655.53 -0.23 655.53 0.36 

 HUB 5.56 0.00 -0.02 589.98 -0.20 589.98 0.32 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 62.62 55.70 6.92 1391.89 1567.80 319.10 1406.66 0.86 

 MEAN 64.25 57.60 6.65 1358.91 1508.97 324.77 1413.34 0.83 

 HUB 66.00 59.00 7.00 1325.11 1450.70 330.49 1420.00 0.79 

 

 ROTOR EXIT CONDITIONS, STAGE 8 

 

 B2 axial THK AeroBl 

 0.05 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 5.84 933.48 648.48 671.46 1885.48 0.50 

 MEAN 5.71 912.48 632.67 657.54 1883.46 0.48 

 HUB 5.57 892.03 596.42 663.33 1878.86 0.47 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1390.20 1000.50 741.72 0.53 

 MEAN 1358.65 979.50 725.98 0.52 1.59 

 HUB 1326.35 986.31 729.93 0.52 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 469.94 1.33 398.85 1581.94 1.09 1513.62 0.85 

 MEAN 475.61 1.34 406.42 1575.66 1.09 1510.38 0.93 

 HUB 466.31 1.32 400.91 1565.40 1.08 1503.02 0.94 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 44.00 47.85 43.80 4.05 0.93 0.53 1.21 

 MEAN 43.90 47.83 43.80 4.03 0.93 0.52 1.24 

 HUB 41.96 47.74 43.80 3.94 0.93 0.48 1.28 

 

 STAGE EXIT CONDITIONS, STAGE 8 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.877 467.192 1.318 1.088 88.784 0.688 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 851134.44 0.440 1365.482 393019.219 

 

 

 OVERALL EXIT CONDITIONS; ALL 8 STAGES 

 

 Del Enthalpy GHP EFFICIENCY PR TR 

 6755277.50 10837.5469 0.8032 31.7818 3.0353 
NASA/TM—2009-215641 27 

Appendix D.—Axial-Centrifugal Compressor Data 

Below are the axial-centrifugal compressor mean-line flow analysis results at the design point operating condition.

 LCTR 7 Axial 1 Centrifugal Stage 100% Speed 

 

 COMPRESSOR INLET CONDITIONS, STAGE 1 

RESET = 0.000 BLEED = 0.000 DPInc 5.000 Kg/sec = 12.90909 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.40 27289.000 14.613 518.670 1.100 0.800 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 28.40 27289.789 1.402 0.249 53.349 28.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 1 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 6.28 0.00 -0.02 636.97 -0.22 636.97 0.59 

 MEAN 4.71 0.00 -0.02 579.07 -0.20 579.07 0.53 

 HUB 2.25 0.00 -0.02 463.25 -0.16 463.25 0.42 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 66.93 62.00 4.93 1495.05 1625.29 11.57 485.12 1.50 

 MEAN 62.72 57.56 5.16 1122.83 1263.54 12.04 490.66 1.16 

 HUB 49.12 44.20 4.92 535.10 707.89 12.94 500.92 0.64 

 

 ROTOR EXIT CONDITIONS, STAGE 1 

 

 B2 axial THK AeroBl 

 0.30 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 6.28 864.45 685.22 527.02 1222.52 0.71 

 MEAN 5.14 870.86 667.80 558.97 1188.55 0.73 

 HUB 3.67 1048.17 801.40 675.59 1140.89 0.92 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1495.05 966.22 809.83 0.79 

 MEAN 1224.70 789.04 556.90 0.66 2.00 

 HUB 874.41 679.52 73.00 0.60 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 32.96 2.26 23.60 683.30 1.32 620.90 0.83 

 MEAN 30.05 2.06 21.02 650.09 1.25 586.78 0.91 

 HUB 27.51 1.88 15.93 631.26 1.22 539.70 0.92 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 52.44 56.95 53.20 3.75 0.93 0.57 1.32 

 MEAN 50.07 44.89 40.10 4.79 0.93 0.53 1.75 

 HUB 49.87 6.17 1.00 5.17 0.93 0.23 3.68 

 

 STAGE EXIT CONDITIONS, STAGE 1 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.858 29.677 2.031 1.263 272.485 2.112 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 848001.94 0.379 1360.967 1192642.500 

 

 
NASA/TM—2009-215641 28 

 COMPRESSOR INLET CONDITIONS, STAGE 2 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.40 27289.000 29.677 654.883 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 15.81 24286.434 1.401 0.249 53.349 38.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 2 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 6.10 0.00 -0.02 679.69 -0.23 679.69 0.56 

 MEAN 5.23 0.00 -0.02 617.90 -0.21 617.90 0.50 

 HUB 4.18 0.00 -0.02 556.11 -0.19 556.11 0.45 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 64.93 59.00 5.93 1452.66 1604.02 24.06 616.74 1.31 

 MEAN 63.62 57.42 6.20 1245.69 1390.71 24.93 623.05 1.14 

 HUB 60.84 54.50 6.34 996.62 1141.44 25.83 629.35 0.92 

 

 ROTOR EXIT CONDITIONS, STAGE 2 

 

 B2 axial THK AeroBl 

 0.20 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 5.98 865.82 633.85 589.80 1331.49 0.65 

 MEAN 5.26 873.23 644.05 589.68 1316.63 0.66 

 HUB 4.43 941.87 700.49 629.63 1297.19 0.73 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1424.09 986.07 790.23 0.74 

 MEAN 1253.29 847.87 609.24 0.64 1.88 

 HUB 1055.20 722.67 354.71 0.56 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 53.91 1.82 40.57 799.56 1.22 737.07 0.84 

 MEAN 53.30 1.80 39.67 784.25 1.20 720.69 0.92 

 HUB 51.19 1.72 36.04 773.34 1.18 699.41 0.93 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 47.06 53.26 49.47 3.79 0.93 0.54 1.27 

 MEAN 47.52 45.93 41.46 4.47 0.93 0.55 1.48 

 HUB 48.05 29.40 24.00 5.40 0.93 0.54 1.85 

 

 STAGE EXIT CONDITIONS, STAGE 2 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.875 52.097 1.755 1.200 219.481 1.701 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 816707.63 0.403 1310.742 812599.375 

 

 
NASA/TM—2009-215641 29 

 COMPRESSOR INLET CONDITIONS, STAGE 3 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.40 27289.000 52.097 785.718 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 9.86 22172.389 1.399 0.251 53.349 50.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 3 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 5.79 0.00 -0.02 729.10 -0.25 729.10 0.54 

 MEAN 5.22 0.00 -0.02 662.81 -0.23 662.81 0.49 

 HUB 4.59 0.00 -0.02 596.53 -0.21 596.53 0.44 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 62.14 56.00 6.14 1378.84 1559.96 42.62 742.01 1.16 

 MEAN 61.96 55.70 6.26 1244.19 1409.93 44.09 749.24 1.05 

 HUB 61.38 55.00 6.38 1093.07 1245.43 45.60 756.46 0.92 

 

 ROTOR EXIT CONDITIONS, STAGE 3 

 

 B2 axial THK AeroBl 

 0.15 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 5.71 868.09 609.75 617.89 1433.13 0.61 

 MEAN 5.19 855.65 596.31 613.63 1422.29 0.60 

 HUB 4.61 895.51 633.67 632.78 1411.88 0.63 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1359.79 971.77 750.04 0.68 

 MEAN 1236.09 886.48 639.78 0.62 1.77 

 HUB 1098.55 785.19 464.88 0.56 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 83.00 1.59 64.79 917.93 1.17 855.37 0.84 

 MEAN 82.05 1.57 64.26 903.25 1.15 842.47 0.92 

 HUB 80.47 1.54 61.39 896.71 1.14 830.13 0.93 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 44.62 50.52 46.64 3.88 0.93 0.54 1.22 

 MEAN 44.18 46.19 42.01 4.18 0.93 0.53 1.36 

 HUB 45.04 36.30 31.47 4.83 0.93 0.53 1.54 

 

 STAGE EXIT CONDITIONS, STAGE 3 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.874 80.924 1.553 1.153 192.065 1.489 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 754495.94 0.408 1210.898 703666.375 

 

 

 
NASA/TM—2009-215641 30 

 COMPRESSOR INLET CONDITIONS, STAGE 4 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.40 27289.000 80.924 905.963 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 6.82 20648.596 1.395 0.252 53.349 60.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 4 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 5.54 0.00 -0.02 697.68 -0.24 697.68 0.48 

 MEAN 5.09 0.00 -0.02 634.26 -0.22 634.25 0.44 

 HUB 4.61 0.00 -0.02 570.83 -0.20 570.83 0.39 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 62.12 56.00 6.12 1318.83 1492.21 69.06 866.19 1.03 

 MEAN 62.40 56.20 6.20 1213.05 1369.05 70.93 872.77 0.95 

 HUB 62.52 57.00 5.52 1097.12 1236.91 72.84 879.34 0.85 

 

 ROTOR EXIT CONDITIONS, STAGE 4 

 

 B2 axial THK AeroBl 

 0.10 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 5.44 863.90 584.23 636.39 1519.69 0.57 

 MEAN 5.03 852.54 575.81 628.70 1512.51 0.56 

 HUB 4.58 879.96 603.22 640.68 1505.43 0.58 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1295.49 954.40 711.26 0.63 

 MEAN 1197.27 884.01 621.46 0.58 1.69 

 HUB 1090.24 804.77 487.02 0.53 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 117.54 1.45 94.47 1025.79 1.13 964.21 0.84 

 MEAN 117.44 1.45 94.72 1015.10 1.12 955.13 0.92 

 HUB 115.96 1.43 92.08 1010.07 1.11 946.18 0.93 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 42.55 48.18 44.26 3.92 0.93 0.52 1.19 

 MEAN 42.49 44.67 40.51 4.16 0.93 0.52 1.29 

 HUB 43.28 37.24 32.62 4.62 0.93 0.52 1.43 

 

 STAGE EXIT CONDITIONS, STAGE 4 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.872 115.820 1.431 1.123 172.586 1.338 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 701661.88 0.418 1126.104 637040.500 

 

 

 

 
NASA/TM—2009-215641 31 

 COMPRESSOR INLET CONDITIONS, STAGE 5 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.40 27289.000 115.820 1016.988 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 5.05 19488.920 1.391 0.255 53.349 70.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 5 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 5.24 0.00 -0.02 690.70 -0.24 690.70 0.45 

 MEAN 4.88 0.00 -0.02 627.91 -0.22 627.91 0.41 

 HUB 4.48 0.00 -0.02 565.12 -0.19 565.12 0.37 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 61.04 54.80 6.24 1247.86 1426.47 100.90 978.31 0.93 

 MEAN 61.60 55.20 6.40 1161.06 1320.17 103.26 984.70 0.86 

 HUB 62.10 55.50 6.60 1067.23 1207.79 105.66 991.09 0.78 

 

 ROTOR EXIT CONDITIONS, STAGE 5 

 

 B2 axial THK AeroBl 

 0.10 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 5.14 849.48 575.31 625.00 1597.25 0.53 

 MEAN 4.78 844.07 576.55 616.48 1592.18 0.53 

 HUB 4.39 854.35 579.87 627.43 1585.22 0.54 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1224.05 900.82 648.73 0.56 

 MEAN 1138.39 834.09 561.85 0.52 1.70 

 HUB 1045.75 781.48 465.88 0.49 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 158.30 1.37 130.71 1127.51 1.11 1068.43 0.84 

 MEAN 159.28 1.38 131.68 1120.00 1.10 1061.66 0.93 

 HUB 156.10 1.35 128.25 1112.16 1.09 1052.39 0.94 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 42.63 46.07 42.00 4.07 0.93 0.53 1.21 

 MEAN 43.08 42.35 38.00 4.35 0.93 0.53 1.30 

 HUB 42.74 36.59 32.00 4.59 0.93 0.52 1.42 

 

 STAGE EXIT CONDITIONS, STAGE 5 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.873 156.506 1.351 1.101 159.814 1.239 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 655986.81 0.438 1052.800 604502.250 

 

 

 
NASA/TM—2009-215641 32 

 COMPRESSOR INLET CONDITIONS, STAGE 6 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.40 27289.000 156.506 1119.889 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 3.92 18571.986 1.386 0.257 53.349 80.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 6 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 4.91 0.00 -0.02 703.82 -0.24 703.82 0.44 

 MEAN 4.60 0.00 -0.02 639.84 -0.22 639.84 0.40 

 HUB 4.27 0.00 -0.02 575.85 -0.20 575.85 0.36 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 58.96 52.20 6.76 1169.27 1364.97 137.44 1080.07 0.85 

 MEAN 59.73 53.20 6.53 1096.05 1269.33 140.47 1086.66 0.79 

 HUB 60.50 54.20 6.30 1017.58 1169.39 143.54 1093.23 0.72 

 

 ROTOR EXIT CONDITIONS, STAGE 6 

 

 B2 axial THK AeroBl 

 0.10 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 4.83 845.97 548.72 643.88 1661.44 0.51 

 MEAN 4.52 833.86 537.94 637.13 1656.75 0.50 

 HUB 4.18 864.62 579.20 641.95 1653.46 0.52 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1151.18 881.78 602.46 0.53 

 MEAN 1076.35 834.15 538.40 0.50 1.65 

 HUB 995.91 765.34 416.70 0.46 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 202.28 1.29 169.75 1218.08 1.09 1159.98 0.85 

 MEAN 202.48 1.29 170.58 1209.89 1.08 1153.45 0.93 

 HUB 202.83 1.30 168.62 1209.54 1.08 1148.86 0.94 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 40.44 43.10 39.00 4.10 0.93 0.52 1.21 

 MEAN 40.18 40.20 36.00 4.20 0.93 0.51 1.29 

 HUB 42.06 32.99 28.36 4.63 0.93 0.52 1.39 

 

 STAGE EXIT CONDITIONS, STAGE 6 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.871 200.887 1.284 1.083 154.647 1.199 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 596158.88 0.450 956.781 593802.938 

 

 

 
NASA/TM—2009-215641 33 

 COMPRESSOR INLET CONDITIONS, STAGE 7 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.40 27289.000 200.887 1212.503 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 3.18 17848.609 1.382 0.260 53.349 82.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 7 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 4.65 0.00 -0.02 683.27 -0.24 683.27 0.41 

 MEAN 4.37 0.00 -0.02 621.16 -0.21 621.16 0.37 

 HUB 4.07 0.00 -0.02 559.04 -0.19 559.04 0.33 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 58.31 52.10 6.21 1106.55 1300.70 179.46 1175.31 0.78 

 MEAN 59.16 53.00 6.16 1040.27 1211.79 182.88 1181.46 0.72 

 HUB 60.04 53.80 6.24 969.47 1119.28 186.35 1187.61 0.66 

 

 ROTOR EXIT CONDITIONS, STAGE 7 

 

 B2 axial THK AeroBl 

 0.10 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 4.58 826.13 547.75 618.44 1721.39 0.48 

 MEAN 4.28 813.61 535.13 612.86 1717.06 0.47 

 HUB 3.97 836.11 563.38 617.81 1713.53 0.49 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1090.45 822.79 542.70 0.48 

 MEAN 1020.07 781.52 484.95 0.46 1.67 

 HUB 944.47 725.89 381.09 0.42 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 251.41 1.25 215.16 1304.43 1.08 1249.53 0.84 

 MEAN 251.46 1.25 216.03 1296.52 1.07 1243.27 0.92 

 HUB 250.66 1.25 213.42 1294.39 1.07 1238.15 0.93 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 41.53 41.27 37.00 4.27 0.93 0.54 1.22 

 MEAN 41.13 38.35 34.00 4.35 0.93 0.52 1.30 

 HUB 42.36 31.67 27.00 4.67 0.93 0.53 1.39 

 

 STAGE EXIT CONDITIONS, STAGE 7 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.05 0.868 249.360 1.241 1.071 148.361 1.150 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 558713.00 0.470 896.684 584503.813 

 

 

 
NASA/TM—2009-215641 34 

 COMPRESSOR INLET CONDITIONS, STAGE 8 

 

 W act RPM act Pt Tt 'POTS POTH AeroBl 

 28.40 27289.000 249.360 1298.448 1.100 0.900 0.980 

 

 W cor RPM cor GAMMA Cp 'R NBLAD THK 

 2.65 17247.789 1.377 0.262 53.349 24.000 0.030 

 

 ROTOR LEADING EDGE CONDITIONS, STAGE 8 

 

 R1 Stator Alfa C1 CU1 Cm1 Abs MACH 

 TIP 4.10 0.00 -0.02 562.70 -0.19 562.70 0.32 

 MEAN 3.81 0.00 -0.02 511.55 -0.18 511.55 0.29 

 HUB 3.50 0.00 -0.02 460.39 -0.16 460.39 0.26 

 

 BetaFlo BetaBlade Incid U1 W1 Ps1 Ts1 Rel Mach 

 TIP 60.05 54.00 6.05 976.38 1127.09 232.26 1273.45 0.65 

 MEAN 60.60 54.50 6.10 907.75 1042.12 235.03 1277.59 0.60 

 HUB 61.09 55.00 6.09 833.50 952.33 237.81 1281.71 0.54 

 

 ROTOR EXIT CONDITIONS, STAGE 8 

 

 B2 axial THK AeroBl 

 0.28 0.030 0.950 

 

 R2 C2 Cu2 Cm2 Ao2 Mach2 

 TIP 6.70 1286.38 1137.24 601.22 1846.69 0.70 

 MEAN 6.70 1281.72 1149.70 566.57 1849.22 0.69 

 HUB 6.70 1281.17 1151.12 562.42 1849.51 0.69 

 

 U2 W2 Wu2 Mach Rel2 Ws1/W2 

 TIP 1595.55 755.98 458.31 0.41 

 MEAN 1595.55 720.96 445.85 0.39 1.54 

 HUB 1595.55 716.82 444.42 0.39 

 

 Pt2 PR Ps2 Tt2 TR Ts2 Eff2 

 TIP 455.63 1.83 330.98 1574.88 1.21 1442.97 0.84 

 MEAN 483.37 1.94 352.20 1577.91 1.22 1446.95 0.92 

 HUB 486.92 1.95 354.91 1578.25 1.22 1447.40 0.93 

 

 Alfa2 Beta FLO Beta BLADE Deviat Slip F. Diff Fct Solidity 

 TIP 62.14 37.32 20.00 17.32 0.85 0.75 1.50 

 MEAN 63.77 38.20 20.00 18.20 0.85 0.74 1.61 

 HUB 63.96 38.32 20.00 18.32 0.85 0.70 1.76 

 

 STAGE EXIT CONDITIONS, STAGE 8 

 

 DIFF LOSS Effic Pdisch PR TR Ns Ns nondim 

 0.17 0.828 453.331 1.818 1.215 55.924 0.434 

 

 Del Enthalpy Del_H/U^2 GHP Reynolds# 

 1828923.63 0.718 2935.257 554689.563 

 

 

 OVERALL EXIT CONDITIONS; ALL 8 STAGES 

 

 Del Enthalpy GHP EFFICIENCY PR TR 

 6760649.50 10850.2324 0.7961 30.8389 3.0405
NASA/TM—2009-215641 35 

References 

1. Johnson, Wayne; Yamauchi, Gloria K.; and Watts, 

Michael E.: NASA Heavy Lift Rotorcraft Systems 

Investigation. NASA/TP–2005-213467, 2005. 

2. Snyder, C.: Gas Turbine Characteristics for a Large Civil 

Tilt-Rotor (LCTR). AHS 65th Annual Forum & 

Technology Display, Grapevine, Texas, 2009. 

3. Veres, Joseph: Axial and Centrifugal Compressor Mean 

Line Flow Analysis Method. AIAA–2009–1641, 2009. 

4. Reid, L.; and Moore, R.D.: Design and Overall 

Performance of Four Highly Loaded, High Speed Inlet 

Stages for an Advanced High-Pressure-Ratio Core 

Compressor. NASA TP–1337, 1978. 

5. Steinke, Ronald J.: Design of 9.271-Pressure-Ratio FiveStage Core Compressor and Overall Performance for First 

Three Stages. NASA TP–2597, fig. 22, 1986. 

6. Holloway, P.R., et al.: Energy Efficient Engine High 

Pressure Compressor Detail Design Report. NASA CR–

165558, 1982. 

7. Balje, O.E.: Turbomachines: A Guide to Design Selection 

and Theory. John Wiley & Sons, New York, NY, 1981.
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OMB No. 0704-0188 

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PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-08-2009 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Compressor Study to Meet Large Civil Tilt Rotor Engine Requirements 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Veres, Joseph, P. 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 877868.02.07.03.01.02.02 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-16952 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/TM-2009-215641 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Categories: 02 and 07 

Available electronically at http://gltrs.grc.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

A vehicle concept study has been made to meet the requirements of the Large Civil Tilt Rotorcraft vehicle mission. A vehicle concept was 

determined, and a notional turboshaft engine system study was conducted. The engine study defined requirements for the major engine 

components, including the compressor. The compressor design-point goal was to deliver a pressure ratio of 31:1 at an inlet weight flow of 

28.4 lbm/sec. To perform a conceptual design of two potential compressor configurations to meet the design requirement, a mean-line 

compressor flow analysis and design code were used. The first configuration is an eight-stage axial compressor. Some challenges of the allaxial compressor are the small blade spans of the rear-block stages being 0.28 in., resulting in the last-stage blade tip clearance-to-span ratio 

of 2.4 percent. The second configuration is a seven-stage axial compressor, with a centrifugal stage having a 0.28-in. impeller-exit blade 

span. The compressors’ conceptual designs helped estimate the flow path dimensions, rotor leading and trailing edge blade angles, flow 

conditions, and velocity triangles for each stage. 

15. SUBJECT TERMS

Compressor; Turbomachinery; Aerodynamics 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

33 

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) 

a. REPORT

U 

b. ABSTRACT

U 

c. THIS 

PAGE

U 

19b. TELEPHONE NUMBER (include area code)

443-757-5802 

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18