Random Technical Questions

[CH3NO2]

Feb 13, 2019 20:06:07 GMT -5 racket said:

Hi Tony

The turb scroll choke throat should be at the end of the "tongue' close to the wheel where the gases enter it .

If you measure the area at that point and divde by the distance from the "Flow area centre" to the shaft centre , it should give your A/R .

Cheers
John

OK, I was able to do that and it was helpful to understand your description.   Previously I was under the impression that A/R meant scroll inlet area / wheel radius.   Which is very different from choke area at the tongue / the distance to shaft centerline. 

Thank you,
Tony

[racket]

Hi Tony

The 98 mm might be OK for bleed , but probably not for thrust .................somehow remember doing the calcs somewhere for someone .

With bleed the temps will be low initially and as the bleed was instigated the temps will need to rise but the mass flow through the turb stage will drop , sorta keeping things in "balance"

May I suggest you do some calcs using the turbine flow map with T I Ts of between 500C and 1,000 C at your desired 4:1 PR

When I purchased my TV84 with its 3.5"- 89mm inducered comp , the turb stage had an undersized 97/86 mm turb wheel, and a 1.08 A/R scroll which was soon changed to a 1.39 A/R , the engine was equiped with a 73 mm jet nozzle and the engine operated OK up to ~3:1 PR but then went into surge with a thrust of ~30 kgs - 66 lbs. indicating a mass flow of ~1.4 lbs/sec .

The surge knocked off turb blade tip bits at both inducer and exducer from contact with the housing shroud.

Eventually a change to a 110/96 mm turb wheel was made and no more problems encountered .

Cheers
John

[racket]

Hi Tony

there are a couple of good turbo books that might be worth exploring

www.ebay.com/itm/B0006X03GM-Turbochargers/302573905621?hash=item4672cf66d5:g:Mg8AAOSwDkVaPCsj:rk:7:pf:0

a latter edition
www.ebay.com/itm/Turbochargers-by-Hugh-Macinnes-1987-Turbo-Design-Street-Race-Cars-Cycles-Boats/183356069291?epid=439493&hash=item2ab0df95ab:g:vVkAAOSwOfBZvgPJ:rk:9:pf:0

and this one
www.ebay.com/itm/Turbo-Real-World-High-Performance-Turbocharger-Systems-SA123/264148810103?epid=64166225&hash=item3d807ef977:g:LUcAAOSwaPNcFUE-:rk:4:pf:0

Cheers
John

[finiteparts]

Tony,

I measured up my GT55 turbine, which looks to be the same size and I would guess, the same casting as yours...I got a throat area (Plane 44) of:

A44 = 6.051 inches

I am still working on my cycle program, but I thought that I would plug the numbers in to see what I come up with... I am still checking the solver for the gas properties (enthalpy, entropy, Cp, gamma, etc.), but  so far the checks that I have done have lined up well to the NASA published properties. 

This program uses a compressible flow solver to find the local gas properties and then standard relations to solve the momentum/mass/torque exchanges between each station.  I assumed a mechanical efficiency of 98.5% and rotor speed of 76 krpm.  I am still working on the radial turbine module, but I did have it to the point where I was able to solve for the turbine throat plane relative gas properties, so I just added a mass continuity check to solve for the relative Mach number at the turbine throat and with the 6.05 inch throat area I am not seeing any choking.  I have a vane passage filled with RTV to make a mold of the passage so that I can cut it up and measure it more accurately. 

Like I said, I am still working through this program and it is not guaranteed to be correct...just thought I would show my first pass at running the GT55 turbine through it. 

Here is the output that I get...by the way...I am using the SAE AS755F standard station numbering scheme...that means that:

1 - is the inlet plane

2 - is the compressor inlet plane

3 - is the compressor stage discharge plane

4 - is the combustor discharge plane

41 - is the NGV throat

44 - is the turbine rotor throat

45 - is the exit plane of the turbine

5 - is just downstream of the turbine discharge

9 - is the discharge nozzle throat

******************** Compressor Performance ***********************

*********** Compressor Inlet **************

Inlet corrected flow,                                 W2c = 2.33 lbm/s
Inlet physical flow,                                    W2 = 2.33 lbm/s = 139.8 lbm/min
Compressor inlet geometric area,            Ageo2 = 9.292 in^2
Compressor inlet total pressure,                  Pt2 = 14.696 psia
Compressor inlet static pressure,                  p2 = 12.512 psia
Compressor inlet total temperature,              Tt2 = 59.0 deg F
Compressor inlet static temperature,              T2 = 35.696 deg F
Compressor inlet constant pressure specific heat, Cp2 = 0.2393 Btu/lbm R
Compressor inlet total enthalpy,                   ht2 = 124.1 Btu/lbm
Compressor inlet effective area,                    A2 = 9.292 in^2
Compressor inlet Mach number,                     M2 = 0.485
Compressor inlet acoustic velocity                  a2 = 1091.1 ft/s
Compressor inlet mean axial velocity,          Cm2 = 529.2 ft/s

*********** Compressor Discharge ***********

Compressor discharge geometric area,         Ageo3 = 14.186 in^2
Compressor discharge total temperature,          Tt3 = 390.67 deg F
Compressor discharge total pressure,               Pt3 = 58.78 psia
Compressor discharge total enthalpy,               ht3 = 208.0 Btu/lbm
Temperature rise through the compressor    Tt3-Tt2 = 331.67 deg F
Compressor temperature ratio,                  Tt3/Tt2 = 1.639
Compressor specific work,                           WTc3 = 77.5 Btu/lbm
Power required by the compressor,                PwrC = 276.36 hp
Compressor discharge plane Mach number,        M3 = 0.095

******************** Combustor Performance ************************

Fuel to air ratio,                                          FAR = 0.0161
Specified combustor temperature rise,         delT34 = 1209 deg F
Physical fuel flow rate to achieve Tt4 target,      Wf = 0.0374 lbm/s
Volumetric fuel flow rate,                               VFf = 20.266 gal/hr
Combustor annulus Mach number,                   M31 = 0.103

******************** Turbine Performance ***************************

Turbine pressure ratio,                                PRt = 0.552
Turbine expansion ratio,                              ERt = 1.81

************ NGV **************

Turbine inlet total temperature,                 Tt4 = 1600 deg F
Turbine inlet static temperature,               T40 = 1588.9 deg F
Throat static temperature,                       T41 = 1321.7 deg F
Turbine inlet total pressure,                      Pt4 = 55.845 psia
Turbine inlet static pressure,                      p4 = 52.154 psia
Throat static pressure,                             p41 = 30.44 psia

Turbine inlet constant pressure specific heat, Cp4 = 0.2865 Btu/lbm R
Turbine inlet enthalpy,                               ht4 = 590.07 Btu/lbm
Engine temperature ratio,                     Tt4/Tt2 = 3.971

Physical flow at plane 4,                            W4 = 2.3674 lbm/s
Corrected flow at plane 4,                         Wc4 = 1.2415 lbm/s
Turbine critical flow area,                      A41crit = 3.6985 in^2 (slight discrepancy between ideal solver and gas property solver)
NGV Throat effective area,                       Ae41 = 3.692 in^2
NGV absolute discharge velocity,                V41 = 1998.7 ft/s
NGV absolute discharge angle,               alpha41 = 62.7 deg
NGV absolute discharge static temperature,    T41 = 1321.7 deg F
Turbine stage inlet Mach number,                  M40 = 0.186
Throat Mach Number,                                  M41 = 0.997
Throat acoustic speed,                                 a41 = 2004.7 ft/s
Throat mean velocity,                                  V41 = 1998.7 ft/s
Turbine NGV critical pressure ratio,       Pt4/p41crit = 1.841
Turbine NGV actual pressure ratio,              Pt4/p4 = 1.835

********* Turbine Rotor ***************

Turbine discharge total pressure,                  Pt5 = 30.8 psia
Turbine discharge static pressure,                  p5 = 28.99 psia
Turbine discharge total temperature,             Tt5 = 1326.75 deg F
Turbine discharge total enthalpy,                   ht5 = 506.28 Btu/lbm
Turbine total temperature drop,               Tt4-Tt5 = 273.2 deg F
Turbine static temperature drop,                T4-T5 = 123.0 deg F
Turbine NGV static temperature drop,        T4-T41 = 267.2 deg F
Turbine rotor static temperature drop,       T41-T5 = -144.2 deg F
Turbine Total to total Expansion ratio,       Pt4/Pt5 = 1.813
Turbine Total to static Expansion ratio,       Pt4/p5 = 1.927
Turbine temperature ratio,                       Tt4Tt5 = 1.153

Turbine discharge effective area,                    A5 = 12.629 in^2
Turbine discharge critical flow area,          A50crit = 6.246 in^2
Turbine exit Mach number,                            M5 = 0.305
Turbine matching work,                              WT4 = 83.78 Btu/lbm
Turbine ideal enthalpy drop,                  delhideal = 161.3 Btu/lbm
Turbine stage enthalpy drop,                 delht4ht5 = 83.8 Btu/lbm

************************* Exhaust Nozzle ******************************

Nozzle total pressure ratio,                      Pt5/Pt9 = 2.0957
Nozzle static to total pressure ratio,           Pt0/P5 = 0.4772
Nozzle driving pressure difference,             P5-P9 = 14.29 psia
Nozzle discharge Mach number,                      M9 = 1.01

******************************* Performance ***************************

Compressor Power,                                   PwrC = 276.4 hp
Turbine Power,                                         PwrT = 280.6 hp
Turbine Ideal Power,                           PwrTideal = 540.1 hp
Mechanical power absorbed,                  PwrMech = 4.209 hp
Compressor spec work,                         SpecWC = 77.5 Btu/lbm
Turbine spec work,                               SpecWT = 83.8 Btu/lbm

********* Densities ***************

Compressor inlet density,         = 0.0682 lbm/in^3
Compressor outlet density,       = 0.1858 lbm/in^3
Combustor outler density,         = 0.0732 lbm/in^3
NGV Throat density,                 = 0.0461 lbm/in^3
Turbine outlet density,              = 0.0437 lbm/in^3
Nozzle throat density,               = 0.029 lbm/in^3

********* Pressure Ratios ***************

Compressor pressure ratio,        PRc = 4.0
Combustor pressure ratio,    PRcomb = 0.95
Turbine pressure ratio,               PRt = 0.552
Nozzle pressure ratio,               PRn = 0.477
Pressure ratio Product Check            = 1.0

********* Enthalpies ***************

Compressor inlet enthalpy      = 124.129 Btu/lbm
Combustor inlet enthalpy        = 207.98 Btu/lbm
Turbine inlet enthalpy            = 590.067 Btu/lbm
Nozzle inlet enthalpy             = 506.284 Btu/lbm

********* Station Properties ***************

Compressor inlet, Plane 20     Tt2 = 59.0 deg F,     Ts2 = 35.7 deg F,       Pt2 = 14.696 psia       Ps2 = 12.512 psia
Compressor outlet, Plane 31   Tt3 = 390.7 deg F,    Ts3 = 389.2 deg F,     Pt3 = 58.784 psia       Ps3 = 58.417 psia
Combustor outlet, Plane 40     Tt4 = 1600 deg F,     Ts4 = 1588.9 deg F,   Pt4 = 55.845 psia       Ps4 = 52.154 psia
Turbine NGV oulet,Plane 41   Tt41 = 1600 deg F,   Ts41 = 1321.7 deg F,  Pt41 = 55.845 psia     Ps41 = 30.435 psia
Turbine outlet, Plane 45         Tt5 = 1326.8 deg F,   Ts5 = 1300.2 deg F,   Pt5 = 30.799 psia       Ps5 = 28.985 psia
Nozzle outlet, Plane 9            Tt9 = 1326.8 deg F,   Ts9 = 1072.9 deg F,   Pt9 = 25.832 psia      Ps9 = 14.696 psia

********* Turbine Areas ***************

Turbine NGV discharge, Plane 41           Area = 3.692 in^2        Throat Mach Number = 0.997

Turbine discharge relative, Plane 44      Area = 6.051 in^2         Throat Mach Number = 0.859

Good luck

Chris

[finiteparts]

Already spotted an error...density units should be in lbm/ft^3...sorry!

[finiteparts]

Also, I used the RMS blade discharge plane to do these calculations...this the need to fill the turbine passage with silicon and then mark and cut the resulting plug to be able to take accurate measurements of the RMS blade angle. Not to mention, if you use the tip angle to represent the entire flow turning, you will over predict the torque.

Chris

[finiteparts]

One more additional point....notice that the static temperature of the gas entering the turbine is only around 1321 F...I didn't include the turbine inducer total relative temperature, but it is around 1310 F. What is often misunderstood is that the 1600F inlet temperature is not what the metal of the turbine wheel sees...it is exposed to this lower temperature gas temperature by virtue of the gas being in a rotating frame of reference. The NGV sees the full effect of the 1600F gas, but the gas properties in the turbine wheel are more complex due to flow being onboarded to a rotating reference frame.

This is what really drives the complexity of the design model to capture accurately the gas flow at the inducer, throat and exducer planes of the turbine wheel...you can't just assume the conditions that you calculate from the total pressure or total temperature changes across the rotor...they only apply in the stationary reference plane upstream and downstream of the turbine, not in the passages fo the turbine itself.

- Chris

[CH3NO2]

That is pure analytical awesomeness Chris! All stations laid out in detail. Wow.

I can calculate combustion gas chemistry/properties and the properties at a single choke point but its all the other thermodynamic equations in a series of stations I still need to learn.

One could really do some damage with that kind of knowledge.

Tony

[CH3NO2]

Update. I was able to confirm the GT55 turbine wheel and the GTX5533R turbine wheel are different designs.

[philip111]

Hypothetically if we had a gas turbine with a stationary rotor and started pumping compressed air with a compressor in the combustor area which way would the rotor start spinning?

[racket]

One would imagine it would be in the usual direction as theres more flow area , but it might be influenced by a number of variables , guess you'll have to try it to find out ;-)