Enhanced Interactive Rocket Thrust Program (EIRTP)

[MKSheppard]

Really simple Explanation:

What this does is takes in user inputs like nozzle inside diameter, chamber pressure, propellant mix; and calculates thrust and ISP for your rocket. Useful if you want to make accurate rocket engines for KSP!

Version 0.9 download link:

https://www.dropbox.com/s/9oa4t0ejvol5nm7/EIRTPv0-9.zip

Operating Environment:

Windows Command Line Executable (.exe)

Development Platform:

Windows 7 x64 laptop

Language:

C++ (using C++11 standards)

Compiler/IDE:

DevC++ (x32 5.5.3)

using

MinGW GCC 4.7.2 32-bit Release

Opening Scrawl:

**************************************************************************

* Enhanced Interactive Rocket Thrust Program (EIRTP) v0.90 (MAY 2013) *

* Performs One-Dimensional design/analysis of rocket nozzle(s). *

**************************************************************************

This program was originally programmed in Java as a web application back in

2005 by Tom Benson of NASA Glenn Research Center and made available to the

general public via the following URL:

http://www.grc.nasa.gov/WWW/K-12/rocket/ienzl.html

Ported to C++ and improved with more data tables and other additions, such as

supporting (crudely) nuclear thermal rockets, altitude/ISP tables, and support

for Kerbal Space Program in January-February/May 2014 by Ryan Crierie.

LEGAL STUFF: This software is public domain. It may be freely copied and used

in non-commercial products, assuming proper credit to the author(s) are given.

IT MAY NOT BE RESOLD. If you want to use the software for commercial products,

contact the author(s). In no event shall NASA or Ryan Crierie be liable for any

damages resulting from the use of this software.

------------------------------------------------------------------------------

WARNING: The mass-estimator function only works correctly on engines with

'modern' thrust chambers of the configuration(s) used since the mid 1950s. It

returns highly incorrect masses for engines with 'early' thrust chambers of the

V-2/Redstone/RD-100~ era.

------------------------------------------------------------------------------

****************************

Notes:

Yes; I know it says "nuclear thermal rockets" -- there's code in there for NTRs, but it's commented out for this release; because things got a bit too sphagettified, and I needed to 'hit the books' to gather more information on NTRs to make accurate estimations on them.

Case in point:

Simple monopropellant NTRs generate correct ISP figures -- Liquid Hydrogen, Liquid Oxygen, etc; because they're a single gas.

But when you have complex propellants like ammonia or methane for NTRs, the ISP figures are off, because each element of the propellant disassociates at different rates.

Additionally, I need to look again at rocket engine costs for chemical engines. The problem is that engine costs are very proprietary figures in the real world; closely held by corporations; so there are very few real datapoints to base them off of.

Update Probability:

Mildish. I'm kind of burned out on this. Putting it out to see what you all think.

---------------

What does it do?

Simple Command Line Interface (CLI) Program that writes output to files or to a console window.

Here's an example run of EIRTP.exe:

**************************************************************************
*  Enhanced Interactive Rocket Thrust Program (EIRTP) v0.90 (MAY 2013)   *
*  Performs One-Dimensional design/analysis of rocket nozzle(s).         *
**************************************************************************

This program was originally programmed in Java as a web application back in
2005 by Tom Benson of NASA Glenn Research Center and made available to the
general public via the following URL:

http://www.grc.nasa.gov/WWW/K-12/rocket/ienzl.html

Ported to C++ and improved with more data tables and other additions, such as
supporting (crudely) nuclear thermal rockets, altitude/ISP tables, and support
for Kerbal Space Program in January-February/May 2014 by Ryan Crierie.

LEGAL STUFF: This software is public domain. It may be freely copied and used
in non-commercial products, assuming proper credit to the author(s) are given.
IT MAY NOT BE RESOLD.  If you want to use the software for commercial products,
contact the author(s). In no event shall NASA or Ryan Crierie be liable for any
damages resulting from the use of this software.
------------------------------------------------------------------------------
WARNING: The mass-estimator function only works correctly on engines with
'modern' thrust chambers of the configuration(s) used since the mid 1950s. It
returns highly incorrect masses for engines with 'early' thrust chambers of the
V-2/Redstone/RD-100~ era.
------------------------------------------------------------------------------

Input the name of the engine you are simulating: Mainsail

Input Nozzle Exit Diameter (in Inches): 55

Input Nozzle Expansion Ratio: 16


***************
* Nozzle Type *
***************

1.) Regeneratively Cooled (Tube-Wall) Metallic Nozzle (Most Modern Rockets)
2.) Radiatively/Ablatively Cooled Metallic Nozzle (Most RCS Nozzles)
3.) Radiatively/Ablatively Cooled Composite Nozzle (RS-68 Thick Wall)
4.) Radiatively/Ablatively Cooled Composite Nozzle (RL10B-2 Thin Wall)

Input Nozzle Type: 1

Input Chamber Pressure (in PSI): 700
Input Number of Thrust Chambers in Engine (2 for RD-180, etc): 1

***********************
* DATABASES AVAILABLE *
***********************

(1) NASA Glenn IRTP Thermochemical Set; 1 datapoint.
(2) RPA Thermochemical Set (10-6000 PSI Pc); 69 datapoints
-----------------------------------------------------------
Input Database you wish to use for calculations: 2

************************************************************************
*                 PROPELLANTS AVAILABLE (RPA Database)                 *
*   % = Oxidizer/Fuel Tanks are of equal size for this O/F Ratio       *
************************************************************************

1. LOX/75-ALC R:1.24 (V-2)    | 2. LOX/90-ALC R:1.439 (SS-3 SHYSTER)
---------------------------------------------------------------------
3. LOX/RP-1    R:2.7 (RD-180) | 4. LOX/RP-1    R:2.3 (F-1)
5. LOX/Syntin  R:2.7 (RD-180) | 6. LOX/Syntin  R:2.3 (F-1)
7. LOX/Boctane R:2.7 (RD-180) | 8. LOX/Boctane R:2.3 (F-1)
                    9. LF2/Boctane R:2.4
---------------------------------------------------------------------
10. LOX/Methane R:2.7%        | 11. LOX/Methane    R:3.5 (RD-192)
---------------------------------------------------------------------
12.  NTO/MMH R:2.0            | 13. NTO/MMH R:1.9
14. NTO/MMH R:1.6% (STS OMS)  | 15. NTO/MMH R:1.3
---------------------------------------------------------------------
                     16. NTO/UDMH R:1.83%
17. NTO/UDMH R:2.2 (YF-20)    | 18. NTO/UDMH R:2.7 (RD-253/RD-270)
---------------------------------------------------------------------
19. NTO/A-50 R:2.0 | 20. NTO/A-50 R:1.9 (TII SI)| 21% NTO/A-50 R1.6 (LMDE/SPS)
---------------------------------------------------------------------
22. LOX/LH2 R:6.0 (SSME)      | 23. LOX/LH2 R:5.5 (J-2 P/U #1)
24. LOX/LH2 R:5.0 (RL-10)     | 25. LOX/LH2 R:4.5 (J-2 P/U #2)
26. LOX/LH2 R:4.0
---------------------------------------------------------------------
27. LF2/LH2 R:12     | 28. LF2/LH2 R:10    | 29. LF2/LH2 R:8
---------------------------------------------------------------------
30. IRFNA/UDMH R:1.87%        | 31. IRFNA/UDMH R:2.6 (Agena)
---------------------------------------------------------------------
Input Propellant you wish to use for calculations: 11
*************************************
* Engine Packaging/Technology Level *
*************************************

1.) Standard Tech Levels (RL10, F-1, H-1, SSME)
2.) Advanced Development (SpaceX Merlin 1D)
3.) Beyond Bleeding Edge Development

Select Tech Level (and thus cost): 1

*******************************
* Engine Operating Efficiency *
*******************************

This is a combination of thrust chamber efficiency and nozzle efficiency.

Maximum Theoretical Impulse                    : 1.00
(OR-Staged Combustion) Fully Regen. Nozzle     : 0.980-0.990 (RD-180/RD-191)
(Staged Combustion) Fully Regenerative Nozzle  : 0.970-0.975 (SSME)
(Gas Generator) Fully Regenerative Nozzle      : 0.948-0.950 (H-1/J-2/RL10A-3-1)

(Gas Generator) Partially Regenerative Nozzle  : 0.92        (F-1)
(Pressure Fed) Ablatively Cooled Nozzle/Chamber: 0.9026      (RS-18 LMAE)
(Gas Generator) Regen Nozzle, Steering Vanes   : 0.881       (V-2 Engine)
(Pressure Fed) Small Bi-Propellant Thrusters   : 0.795 to 0.850

Input Engine Operating Efficiency (EOE): 0.94

INITAL CALCULATIONS COMPLETED.

Outputting Engine Summary to 'Mainsail_Summary.txt'.

Engine Summary File DONE.

Detailed ISP/Thrust logs will now be computed and made available inside
the following file:

Mainsail_Log.csv

in comma separated variable (CSV) format in intervals of your own chosing.

Input interval between datapoints in feet: 5000

Finished computing detailed altitude/thrust/ISP log.

Holding for acknowledgement.

Mainsail_Summary.txt:

Mainsail engine summary via EIRTP v1.0

PROPELLANT PARAMETERS:
Liquid Oxygen / Liquid Methane -- O/F 3.5
 Temperature of Combustion : 5,856.68F
                     Gamma : 1.18
          Molecular Weight : 21.83
    Oxidizer to Fuel Ratio : 3.50
          Oxidizer Density : 71.17 lb/ft3 (1,140.00 kg/m3)
              Fuel Density : 26.34 lb/ft3 (422.00 kg/m3)
Overall Propellant Density : 51.64 lb/ft3 (827.23 kg/m3)

ENGINE BASIC PARAMETERS:
     Estimated Engine Cost : $3,411,476.69 ($3.41M)
  Propellant Mass Flow (q) : 567.45 lb/sec (257.39 kg/sec)
 Engine Overall Efficiency : 0.94
           Thrust Chambers : 1.00
          Chamber Pressure : 700.00 psi.

***************************************************************************
*          BELL (PARABOLIC) NOZZLE PERFORMANCE DATA                       *
***************************************************************************

ENGINE DIMENSIONS/MASSES:
    Nozzle Expansion Ratio : 16.00
             Nozzle Length : 69.12 inches
    Nozzle Throat Diameter : 13.75 inches.
  Nozzle Throat Area (Ath) : 148.49 square inches.
      Nozzle Exit Diameter : 55.00 inches.
    Nozzle Exit Area (Aex) : 2,375.83 square inches.
            Turbopump Mass : 670.14 lbs (0.30 tonnes)
            Nozzle(s) Mass : 1,215.22 lbs (0.55 tonnes)
         Accessory(s) Mass : 471.34 lbs (0.21 tonnes)
         Total Engine Mass : 2,356.70 lbs (1.07 tonnes)
           Engine T/W (SL) : 60.72
          Engine T/W (Vac) : 75.54

ALTITUDE 	 THRUST 	 SPECIFIC IMPULSE
Sea Level: 	143,102.63	 lbf 	252.19	 sec 	 (100% of Atmosphere)
2,850 ft: 	146,551.78	 lbf 	258.27	 sec 	 (90% of Atmosphere)
6,000 ft: 	150,037.53	 lbf 	264.41	 sec 	 (80% of Atmosphere)
9,500 ft: 	153,536.89	 lbf 	270.57	 sec 	 (70% of Atmosphere)
13,340 ft: 	157,051.53	 lbf 	276.77	 sec 	 (60% of Atmosphere)
17,950 ft: 	160,545.76	 lbf 	282.93	 sec 	 (50% of Atmosphere)
23,250 ft: 	164,040.44	 lbf 	289.09	 sec 	 (40% of Atmosphere)
29,750 ft: 	167,530.71	 lbf 	295.24	 sec 	 (30% of Atmosphere)
38,350 ft: 	171,022.01	 lbf 	301.39	 sec 	 (20% of Atmosphere)
52,800 ft: 	174,524.64	 lbf 	307.56	 sec 	 (10% of Atmosphere)
67,250 ft: 	176,273.39	 lbf 	310.64	 sec 	 (5% of Atmosphere)
81,800 ft: 	177,143.40	 lbf 	312.18	 sec 	 (2.5% of Atmosphere)
96,750 ft: 	177,581.54	 lbf 	312.95	 sec 	 (1.25% of Atmosphere)
112,000 ft: 	177,799.47	 lbf 	313.33	 sec 	 (0.625% of Atmosphere)
Vacuum: 	178,017.69	 lbf 	313.72	 sec

**************************************************************************
*         BELL (PARABOLIC) NOZZLE PERFORMANCE DATA (KSP Ready)           *
**************************************************************************

Propellant Data:
****************************************************
    Oxidizer to Fuel Ratio : 3.50
          Oxidizer Density : 0.0011400 (KSP cfg Units)
              Fuel Density : 0.0004220 (KSP cfg Units)
Overall Propellant Density : 0.0008272 (KSP cfg Units)
****************************************************

PART.CFG information below:

// --- editor parameters ---
cost = 341.15

// --- standard part parameters ---
mass = 1.07 // tonnes.

// maxTemp = 1,381 // Roughly Equilibrium Temperature
// maxTemp = 2,071 // 75% Overheat Bar
maxTemp = 1,864 // 95% Overheat Bar

MODULE
{
	name = ModuleEngines
	thrustVectorTransformName = NozzleTransform
	exhaustDamage = true
	ignitionThreshold = 0.1
	minThrust = 0
	maxThrust = 636.55 	 // kN - sea level
	// maxThrust = 791.86 	 // kN - vacuum
	heatProduction = 291.21
	PROPELLANT
	{
		name = LqdMethane
		ratio = 0.44
		DrawGauge = True
	}
	PROPELLANT
	{
		name = LiquidOxygen
		ratio = 0.56
	}
atmosphereCurve
		{
		key = 1 	252.19
		key = 0.9 	258.27
		key = 0.8 	264.41
		key = 0.7 	270.57
		key = 0.6 	276.77
		key = 0.5 	282.93
		key = 0.4 	289.09
		key = 0.3 	295.24
		key = 0.2 	301.39
		key = 0.1 	307.56
		key = 0.05 	310.64
		key = 0.025 	312.18
		key = 0.0125 	312.95
		key = 0.00625 	313.33
		key = 0 	313.72
		}
}

-------------------------------------------------------------------------------------------
-------------------------------------------------------------------------------------------

***************************************************************************
*          AEROSPIKE (PLUG) NOZZLE PERFORMANCE DATA                       *
***************************************************************************

ENGINE DIMENSIONS/MASSES
Nozzle Aerodynamic Expansion Ratio : 16.00
  Nozzle Geometric Expansion Ratio : 19.93
                   Engine Diameter : 57.75 inches.
                     Nozzle Length : 38.57 inches
             Nozzle Geometric Area : 2,375.83 square inches.
      Nozzle Geometric Throat Area : 119.22 in2.
                    Turbopump Mass : 670.14 lbs. 	0.30 tonnes
                    Nozzle(s) Mass : 684.06 lbs. 	0.31 tonnes
                 Accessory(s) Mass : 338.55 lbs. 	0.15 tonnes
                 Total Engine Mass : 1,692.76 lbs. 	0.77 tonnes
                   Engine T/W (SL) : 73.10
                  Engine T/W (Vac) : 87.53

ALTITUDE 	 THRUST 	 SPECIFIC IMPULSE
Sea Level: 	123,738.83	 lbf 	271.59	 sec 	 (100% of Atmosphere)
2,850 ft: 	125,176.96	 lbf 	274.75	 sec 	 (90% of Atmosphere)
6,000 ft: 	126,855.03	 lbf 	278.43	 sec 	 (80% of Atmosphere)
9,500 ft: 	128,706.24	 lbf 	282.50	 sec 	 (70% of Atmosphere)
13,340 ft: 	130,741.13	 lbf 	286.96	 sec 	 (60% of Atmosphere)
17,950 ft: 	132,982.71	 lbf 	291.88	 sec 	 (50% of Atmosphere)
23,250 ft: 	135,523.94	 lbf 	297.46	 sec 	 (40% of Atmosphere)
29,750 ft: 	138,494.80	 lbf 	303.98	 sec 	 (30% of Atmosphere)
38,350 ft: 	141,582.43	 lbf 	310.76	 sec 	 (20% of Atmosphere)
52,800 ft: 	144,696.06	 lbf 	317.59	 sec 	 (10% of Atmosphere)
67,250 ft: 	146,282.37	 lbf 	321.07	 sec 	 (5% of Atmosphere)
81,800 ft: 	147,100.92	 lbf 	322.87	 sec 	 (2.5% of Atmosphere)
96,750 ft: 	147,537.90	 lbf 	323.83	 sec 	 (1.25% of Atmosphere)
112,000 ft: 	147,775.36	 lbf 	324.35	 sec 	 (0.625% of Atmosphere)
Vacuum: 	148,168.63	 lbf 	325.21	 sec


**************************************************************************
*         AEROSPIKE (PLUG) NOZZLE PERFORMANCE DATA (KSP Ready)           *
**************************************************************************

Propellant Data:
****************************************************
    Oxidizer to Fuel Ratio : 3.50
          Oxidizer Density : 0.0011400 (KSP cfg Units)
              Fuel Density : 0.0004220 (KSP cfg Units)
Overall Propellant Density : 0.0008272 (KSP cfg Units)
****************************************************

PART.CFG information below:

// --- editor parameters ---
cost = 426.43

// --- standard part parameters ---
mass = 0.77 // tonnes.

// maxTemp = 1,381 // Roughly Equilibrium Temperature
// maxTemp = 2,071 // 75% Overheat Bar
maxTemp = 1,864 // 95% Overheat Bar

MODULE
{
	name = ModuleEngines
	thrustVectorTransformName = NozzleTransform
	exhaustDamage = true
	ignitionThreshold = 0.1
	minThrust = 0
	maxThrust = 550.42 	 // kN - sea level
	// maxThrust = 659.09 	 // kN - vacuum
	heatProduction = 334.90
	PROPELLANT
	{
		name = LqdMethane
		ratio = 0.44
		DrawGauge = True
	}
	PROPELLANT
	{
		name = LiquidOxygen
		ratio = 0.56
	}
atmosphereCurve
		{
		key = 1 	271.59
		key = 0.9 	274.75
		key = 0.8 	278.43
		key = 0.7 	282.50
		key = 0.6 	286.96
		key = 0.5 	291.88
		key = 0.4 	297.46
		key = 0.3 	303.98
		key = 0.2 	310.76
		key = 0.1 	317.59
		key = 0.05 	321.07
		key = 0.025 	322.87
		key = 0.0125 	323.83
		key = 0.00625 	324.35
		key = 0 	325.21
		}
}


***********************************************************
EOF

Mainsail_Log.csv output:

NOTE: Aerospike Plug nozzles really don't make sense until you get into crazy high expansion numbers, like 100+

Edited June 6, 2014 by MKSheppard

[MKSheppard]

Forgot about the other program included in EIRTP:

EIRTP Utilities.

It uses the basic EIRTP code, but in special iteration loops to find values that you're looking for:

It's actually kind of cool, in a nerdy way.

**************************************************************************
*  Enhanced Interactive Rocket Thrust Utilities (EIRTP) v0.90 (MAY 2013)   *
*                   Providing Enhanced Functionality                       *
**************************************************************************

The core of this program was originally programmed in Java as a web app back
in 2005 by Tom Benson of NASA Glenn Research Center and made available to the
general public via the following URL:

http://www.grc.nasa.gov/WWW/K-12/rocket/ienzl.html

Ported to C++ and improved in January-February/May 2014 by Ryan Crierie.

LEGAL STUFF: This software is public domain. In no event shall NASA or
Ryan Crierie be liable for any damages resulting from the use of this software.


**************************************************************************
*            SELECT THE PROGRAM FUNCTION THAT YOU WANT TO RUN            *
**************************************************************************

1.) Size Engine by Thrust desired at a Specific Altitude.
2.) Size Engine by Thrust desired, constrained by a given diameter.
3.) Compute Optimum Expansion Ratios from Sea Level to 280,000 ft.
4.) Compute the O/F Ratio resulting from a given O/F tank ratio.
5.) Compute the Engine Operating Efficiency of an engine from known ISP values.
6.) Compute ISP from known thrust and mass flow values.
7.) Compute Vac Thrust when you know vac ISP and sea level thrust/ISP.

**************************************************************************
Enter Menu Choice:2

This selection sizes your engine for thrust desired, using a given diameter for
a stage that the engine will be under; and iterates until it finds the right
chamber pressure to match thrust levels.

Input Altitude (in feet) that you wish to size engine for: 0

Choose Unit of Force for Thrust.
---------------------------------------------------------------------
Units supported are: lbf, klbf, mlbf, kgf, tonnes, N, kN, and MN.
Input is somewhat case insensitive -- all caps is recommended.
---------------------------------------------------------------------
Input Unit of Force to use: TONNES

Input Desired Thrust: 25

Desired thrust is: 55,115.50 lbf.

************************************************************************
* Stage Diameter that your Engine Configuration is Constrained By      *
************************************************************************
Saturn V S-IC/S-II   : 396 inch     (33 feet) Diameter
Saturn V S-IVB       : 260 inch     (21.7 ft) Diameter
Delta IV CBC         : 200 inch     (5.1m)    Diameter
Atlas V CCB          : 150 inch     (12.5 ft) Diameter
KSP Extra Large Part : 147.638 inch (3.75m)   Diameter
Centaur D-1T         : 120 inch     (10 ft)   Diameter
KSP Large Part       : 98.4252 inch (2.5m)    Diameter
KSP Small Part       : 49.2126 inch (1.25m)   Diameter
KSP Tiny Part        : 24.6063 inch (0.625m)  Diameter
************************************************************************
Input Stage Diameter in Inches: 49.21

************************************************************************
NOTE:
Regarding Useable Diameter -- The Saturn V's S-IC stage diameter was
396 inches, and it had a 'rim' which was approximately 10.4 inches thick,
giving a useable inside diameter of 385.6 inches, or 0.973737374 of actual
stage diameter. This program further reduces it to 0.968 to provide extra
margin of safety for stage separation and protection against vibrations.
************************************************************************

Useable diameter available to the engine is 47.64 inches.

This program supports up to 20 thrust chambers for iterative sizing. Please
input the number of thrust chambers (nozzles) your engine has: 4

Maximum Possible Exit Diameter for each engine nozzle: 18.683

************************************************************************
NOTE:
The program will attempt to size an engine for your chamber pressure, but
if it can't meet the thrust requirements within dimensional constraints, it
will begin iteratively increasing chamber pressure until thrust requirements
are met.
************************************************************************

Input Chamber Pressure (PSI): 700
Input Expansion Ratio of Engine: 8

*******************************
* Engine Operating Efficiency *
*******************************

This is a combination of thrust chamber efficiency and nozzle efficiency.

Maximum Theoretical Impulse                    : 1.00
(OR-Staged Combustion) Fully Regen. Nozzle     : 0.98 to 0.99 (RD-180/RD-191)
(Staged Combustion) Fully Regenerative Nozzle  : 0.97 to 0.975 (SSME)
Nuclear Thermal Rockets                        : 0.949 to 0.957 (0.953 Avg)
(Gas Generator) Fully Regenerative Nozzle      : 0.948-0.95 (H-1/J-2/RL10A-3-1)
(Gas Generator) Partially Regenerative Nozzle  : 0.92       (F-1)
(Pressure Fed) Ablatively Cooled Nozzle/Chamber: 0.9026     (RS-18 LMAE)
(Pressure Fed) Small Bi-Propellant Thrusters   : 0.795 to 0.850

Input Engine Operating Efficiency (EOE): .94


***********************
* DATABASES AVAILABLE *
***********************

(1) NASA Glenn IRTP Thermochemical Set; 1 datapoint.
(2) RPA Thermochemical Set (10-6000 PSI Pc); 69 datapoints
-----------------------------------------------------------
Input Database you wish to use for calculations: 2

************************************************************************
*                 PROPELLANTS AVAILABLE (RPA Database)                 *
*   % = Oxidizer/Fuel Tanks are of equal size for this O/F Ratio       *
************************************************************************

1. LOX/75-ALC R:1.24 (V-2)    | 2. LOX/90-ALC R:1.439 (SS-3 SHYSTER)
---------------------------------------------------------------------
3. LOX/RP-1    R:2.7 (RD-180) | 4. LOX/RP-1    R:2.3 (F-1)
5. LOX/Syntin  R:2.7 (RD-180) | 6. LOX/Syntin  R:2.3 (F-1)
7. LOX/Boctane R:2.7 (RD-180) | 8. LOX/Boctane R:2.3 (F-1)
                    9. LF2/Boctane R:2.4
---------------------------------------------------------------------
10. LOX/Methane R:2.7%        | 11. LOX/Methane    R:3.5 (RD-192)
---------------------------------------------------------------------
12.  NTO/MMH R:2.0            | 13. NTO/MMH R:1.9
14. NTO/MMH R:1.6% (STS OMS)  | 15. NTO/MMH R:1.3
---------------------------------------------------------------------
                     16. NTO/UDMH R:1.83%
17. NTO/UDMH R:2.2 (YF-20)    | 18. NTO/UDMH R:2.7 (RD-253/RD-270)
---------------------------------------------------------------------
19. NTO/A-50 R:2.0 | 20. NTO/A-50 R:1.9 (TII SI)| 21% NTO/A-50 R1.6 (LMDE/SPS)
---------------------------------------------------------------------
22. LOX/LH2 R:6.0 (SSME)      | 23. LOX/LH2 R:5.5 (J-2 P/U #1)
24. LOX/LH2 R:5.0 (RL-10)     | 25. LOX/LH2 R:4.5 (J-2 P/U #2)
26. LOX/LH2 R:4.0
---------------------------------------------------------------------
27. LF2/LH2 R:12     | 28. LF2/LH2 R:10    | 29. LF2/LH2 R:8
---------------------------------------------------------------------
30. IRFNA/UDMH R:1.87%        | 31. IRFNA/UDMH R:2.6 (Agena)
---------------------------------------------------------------------
Input Propellant you wish to use for calculations: 11

Select chamber pressure delta increase per cycle if that cycle hits nozzle
constraints. Smaller deltas require more computational cycles. Some crude
examples:

1  PSI Delta : 18~    million iterations.
5  PSI Delta : 3.625~ million iterations.
10 PSI Delta : 1.819~ million iterations.

Input Pressure Delta Per Cycle: 1

Iterations now beginning. This may take some time, particularly if you chose
a low chamber pressure for your starting point and/or a small Pc Delta.

Nozzle Exit Diameter found that matches specifications asked for.

A total of 1,171.000 combinations were evaluated before finding a match.

The Engine you've sized has 4.000 thrust chambers, each with
a throat diameter of 4.144 inches and a exit diameter of 11.720 inches.

Each chamber operates at 700.000psia and generates 13,800.954 lbf of
thrust, for a total of 55,203.817 lbf.

Finished computing your stuff.

Holding for acknowledgement to pass 0 and end.

[MKSheppard]

Oh yes; source for both programs is included in the ZIP.

Yes; you could use this if you were so inclined to begin work on a semi-procedurally generated rocket engine (HINT HINT).

[Justin Kerbice]

What does it do?

Simple Command Line Interface (CLI) Program that writes output to files or to a console window.

and so what ?

Better maybe to put really what is all about on top of your post .

As usual, some kind of people say the stupid "no pic no click", here it could be "no summary no click".

By the way, you're "back" MkSheppard, I have really enjoyed your tutorial series on making parts and the adapter you've made, it have helped me a lot when I began.

[NathanKell]

Welp, guess it's Procedural Engines time.