http://wiki.newmars.com/api.php?action=feedcontributions&user=John+Creighton&feedformat=atomNewMarsWiki - User contributions [en]2024-03-28T14:08:23ZUser contributionsMediaWiki 1.29.1http://wiki.newmars.com/index.php?title=Talk:Heat_transfer_coefficient&diff=799Talk:Heat transfer coefficient2009-09-05T06:35:12Z<p>John Creighton: </p>
<hr />
<div>Wikipedia gives some typical numbers for the heat transfer coefficients for air:<br />
<br />
Air - h = 10 to 100 W/(m2K)<br />
http://en.wikipedia.org/wiki/Heat_transfer_coefficient<br />
<br />
Since the martian atmosphere is thin then these numbers should be even smaller for mars. If the temperature difference is small (less then a degree) then the heat transfer is small. I wonder how much insulation it would take so that the outside wall temperature is near that of the martian atmosphere. In such cases perhaps most of the heat losses will be radiative. However, radiative heat losses could somewhat be balanced by the back radiation in the atmosphere. I wonder if the atmosphere is thick enough to have significant back radiation. <br />
<br />
[[User:John Creighton|John Creighton]] 06:34, 5 September 2009 (UTC)</div>John Creightonhttp://wiki.newmars.com/index.php?title=Talk:Heat_transfer_coefficient&diff=798Talk:Heat transfer coefficient2009-09-05T06:34:48Z<p>John Creighton: New page: Wikipedia gives some typical numbers for the heat transfer coefficients for air: Air - h = 10 to 100 W/(m2K) http://en.wikipedia.org/wiki/Heat_transfer_coefficient Since the martian atmo...</p>
<hr />
<div>Wikipedia gives some typical numbers for the heat transfer coefficients for air:<br />
<br />
Air - h = 10 to 100 W/(m2K)<br />
http://en.wikipedia.org/wiki/Heat_transfer_coefficient<br />
<br />
Since the martian atmosphere is thin then these numbers should be even smaller for mars. If the temperature difference is small (less then a degree) then the heat transfer is small. I wonder how much insulation it would take so that the outside wall temperature is near that of the martian atmosphere. In such cases perhaps most of the heat losses will be radiative. However, radiative heat losses could somewhat be balanced by the back radiation in the atmosphere. I wonder if the atmosphere is thick enough to have significant back radiation. <br />
<br />
[[User:John Creighton|John Creighton]] 06:34, 5 September 2009 (UTC)John Creighton</div>John Creightonhttp://wiki.newmars.com/index.php?title=Tank_geometries&diff=794Tank geometries2009-08-15T19:57:37Z<p>John Creighton: /* Sphere m */</p>
<hr />
<div>== Intro ==<br />
As mentioned in [[RocketDesignSpreadsheet#Section_2_Oxidizer_Tank_Structure_Mass]], there are some important dimensionless quantities which determine how tanks weight scales with volume:<br />
<br />
== (Maxium Circumfrance)/(Volume^(1/3)) ==<br />
<br />
The point of the tank with maximum circumference will be near the weakest cross section of the tank with regards to pressure containment (see [[Pressure_Vessel]]). Strictly speaking for reasons of ease of computation, we take a cross section, near the weakest section, and account for any possible error by including a factor of safety. For instance, in the case of a cylinder, the plane of the cross section is taken to be parallel to the axis of the cylinder. If this is not the weakest cross section with regards to pressure containment it will be close. <br />
<br />
==== Sphere ====<br />
<br />
In the case of a sphere all cross sections, though the center of the sphere have the same circumference.<br />
<br />
<math>C=2*\pi*r</math><br />
<br />
The volume of a sphere is given by:<br />
<br />
<math>V=(4/3)\pi*r^3</math><br />
<br />
Therefore:<br />
<br />
<math>\frac{C}{V^(1/3)}=\frac{2*\pi}{((4/3)\pi)^{1/3}} \approx 3.9</math><br />
<br />
==== Cylinder ====<br />
<br />
As mentioned above the maximum circumference of a cylinder will occur near the cross section, which is parallel to the axis of the cylinder. <br />
<br />
If h is the height of the cylinder and r is the radius then, the circumfrance of this cross section, will go around half a sphere twice (at the ends) and along the tank height twice. Therefore:<br />
<br />
<math>C=2h+2*\pi*r</math><br />
<br />
The volume of this cylinder, will be equal to the cross sectional area multiplied by the hieght plus the area at the ends (two half spheres), which gives.<br />
<br />
<math>V=\pi*r^2*h+4/3*pi*r^3</math><br />
<br />
Now if the hight is some multiple (n) of the radius we get:<br />
<br />
<math>C=2(n+\pi)r</math><br />
<br />
<math>V=\pi*r^2*n*r+4/3*pi*r^3=\pi(n+(4/3))r^3(/math><br />
<br />
Therefore:<br />
<br />
<math>(C/V^(1/3))=\frac{2(n+\pi)}{(\pi(n+(4/3)))^(1/3)}</math> <br />
<br />
== (Maximum Cross Sectional Area)/(Volume^(2/3)) ==<br />
The point of the tank with maximum circumference will be near the weakest cross section of the tank with regards to pressure containment (see [[Pressure_Vessel]]). Strictly speaking for reasons of ease of computation, we take a cross section, near the weakest section, and account for any possible error by including a factor of safety. For instance, in the case of a cylinder, the plane of the cross section is taken to be parallel to the axis of the cylinder. If this is not the weakest cross section with regards to pressure containment it will be close.<br />
<br />
==== Sphere ====<br />
<br />
The cross sectional area of any cross section though the center of the sphere is the same regardless of the cross section and is given by:<br />
<br />
<math>C=\pi r^2</math><br />
<br />
The volume of a sphere is:<br />
<br />
<math>V=(4/3)\pi*r^3</math><br />
<br />
Therefore<br />
<br />
<math>(C/V^(2/3))=\frac{\pi}{((4/3)\pi)^{2/3}}</math><br />
<br />
==== Cylinder ====<br />
<br />
As mentioned above the maximum cross sectional area of a cylinder will occur near the cross section, which is parallel to the axis of the cylinder. The cross sectional area will consist of two parts, the main part of the cylinder which will have an area of 2*r*h, and the two half spheres on the end will have an area of pi*r^2 This gives:<br />
<br />
<math>A_c=2 \ r \ h+\pi r^2</math><br />
<br />
and if h=nr then we get:<br />
<br />
<math>A_c=2 \ r \ h+\pi r^2</math><br />
<br />
<math>A_c=2 \ r \ n \ r+\pi r^2=(2n+pi)r^2</math><br />
<br />
And, therefore<br />
<br />
<math>(C/V^(2/3))=\frac{(2n+pi)}{((4/3)\pi)^{2/3)}</math><br />
<br />
== (Surface Area)/(Volume^(2/3)) ==<br />
<br />
The total mass of the tank, is the surface area multiplied by the wall thickness, multiplied by the density of the material.<br />
<br />
==== Sphere ====<br />
<br />
The surface area of a sphere is given by:<br />
<br />
<math>A_s=4 \pi r^2</math><br />
<br />
Therefore:<br />
<br />
<math>(A_s)/V^(2/3)=\frac{4 \pi}{((4/3)\pi)^{2/3)}</math><br />
<br />
==== Cylinder ====<br />
<br />
The surface area for the main part of the cylinder will be the circumference (2 pi r) multiplied by the height (h). The ends are two half spheres and will have a surface area of 4 pi r^2. Therefore<br />
<br />
<math>A_s=2 \pi r \ h + 4 \pi r^2</math><br />
<br />
and if h=n r then:<br />
<br />
<math>A_s=2 \pi r \ n \ r + 4 \pi r^2=\pi(n +r)r^2</math><br />
<br />
which gives:<br />
<br />
<math>(A_s)/(V^2/3)=\frac{\pi(n +r)}{((4/3)\pi)^{2/3)}</math></div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=793RocketDesignSpreadsheet2009-08-15T19:56:59Z<p>John Creighton: /* Section 2 Oxidizer Tank Structure Mass */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the [[Tank_geometries|geometry of the oxidizer tank]]. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[Pressure_Vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
(For more details on these ratios see: [[Tank_geometries]])<br />
<br />
== Sec-3 Fuel Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the fuel tank and the [[Tank_geometries|geometry of the fuel]] tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each fuel tank and the total structural mass of all the fuel tanks combined. <br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]]<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
The final section of the spread sheet gives some overall performance of the rocket. In this part you specify the delta V you want to travel, the thrust per weight of the engine and the overall thrust per weight. The spread sheet then calculates the mass ratio and the approximate payload. The exit velocity of the fuel is chosen base on the above choice of fuel. Exit velocity is the product of gravity multiplied by the ISP.[4]<br />
<br />
This section also attempts to calculate gravity losses (Also known as Gravity Drag) [5]. The [[optimal ground launch flight path]] is such that the vertical component of the thrust cancels out gravity and the rest of the thrust is direct to horizontal acceleration.<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.<br />
<br />
[4] http://en.wikipedia.org/wiki/Specific_impulse#Specific_impulse_as_a_speed_.28effective_exhaust_velocity.29 - Relationship between specific impulse and ISP.<br />
<br />
[5] http://en.wikipedia.org/wiki/Gravity_drag - Wikipedia article on gravity drag.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=792RocketDesignSpreadsheet2009-08-15T19:56:19Z<p>John Creighton: /* Sec-3 Fuel Tank Structure Mass */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the geometry of the oxidizer tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[Pressure_Vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
(For more details on these ratios see: [[Tank_geometries]])<br />
<br />
== Sec-3 Fuel Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the fuel tank and the [[Tank_geometries|geometry of the fuel]] tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each fuel tank and the total structural mass of all the fuel tanks combined. <br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]]<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
The final section of the spread sheet gives some overall performance of the rocket. In this part you specify the delta V you want to travel, the thrust per weight of the engine and the overall thrust per weight. The spread sheet then calculates the mass ratio and the approximate payload. The exit velocity of the fuel is chosen base on the above choice of fuel. Exit velocity is the product of gravity multiplied by the ISP.[4]<br />
<br />
This section also attempts to calculate gravity losses (Also known as Gravity Drag) [5]. The [[optimal ground launch flight path]] is such that the vertical component of the thrust cancels out gravity and the rest of the thrust is direct to horizontal acceleration.<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.<br />
<br />
[4] http://en.wikipedia.org/wiki/Specific_impulse#Specific_impulse_as_a_speed_.28effective_exhaust_velocity.29 - Relationship between specific impulse and ISP.<br />
<br />
[5] http://en.wikipedia.org/wiki/Gravity_drag - Wikipedia article on gravity drag.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=791RocketDesignSpreadsheet2009-08-15T19:50:09Z<p>John Creighton: /* Sec-3 Fuel Tank Structure Mass */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the geometry of the oxidizer tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[Pressure_Vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
(For more details on these ratios see: [[Tank_geometries]])<br />
<br />
== Sec-3 Fuel Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the fuel tank and the [[Tank_geometries][geometry of the fuel]] tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each fuel tank and the total structural mass of all the fuel tanks combined. <br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]]<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
The final section of the spread sheet gives some overall performance of the rocket. In this part you specify the delta V you want to travel, the thrust per weight of the engine and the overall thrust per weight. The spread sheet then calculates the mass ratio and the approximate payload. The exit velocity of the fuel is chosen base on the above choice of fuel. Exit velocity is the product of gravity multiplied by the ISP.[4]<br />
<br />
This section also attempts to calculate gravity losses (Also known as Gravity Drag) [5]. The [[optimal ground launch flight path]] is such that the vertical component of the thrust cancels out gravity and the rest of the thrust is direct to horizontal acceleration.<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.<br />
<br />
[4] http://en.wikipedia.org/wiki/Specific_impulse#Specific_impulse_as_a_speed_.28effective_exhaust_velocity.29 - Relationship between specific impulse and ISP.<br />
<br />
[5] http://en.wikipedia.org/wiki/Gravity_drag - Wikipedia article on gravity drag.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=790RocketDesignSpreadsheet2009-08-15T19:48:48Z<p>John Creighton: /* Section 2 Oxidizer Tank Structure Mass */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the geometry of the oxidizer tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[Pressure_Vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
(For more details on these ratios see: [[Tank_geometries]])<br />
<br />
== Sec-3 Fuel Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the fuel tank and the geometry of the fuel tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each fuel tank and the total structural mass of all the fuel tanks combined. <br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]]<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
The final section of the spread sheet gives some overall performance of the rocket. In this part you specify the delta V you want to travel, the thrust per weight of the engine and the overall thrust per weight. The spread sheet then calculates the mass ratio and the approximate payload. The exit velocity of the fuel is chosen base on the above choice of fuel. Exit velocity is the product of gravity multiplied by the ISP.[4]<br />
<br />
This section also attempts to calculate gravity losses (Also known as Gravity Drag) [5]. The [[optimal ground launch flight path]] is such that the vertical component of the thrust cancels out gravity and the rest of the thrust is direct to horizontal acceleration.<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.<br />
<br />
[4] http://en.wikipedia.org/wiki/Specific_impulse#Specific_impulse_as_a_speed_.28effective_exhaust_velocity.29 - Relationship between specific impulse and ISP.<br />
<br />
[5] http://en.wikipedia.org/wiki/Gravity_drag - Wikipedia article on gravity drag.</div>John Creightonhttp://wiki.newmars.com/index.php?title=Tank_geometries&diff=789Tank geometries2009-08-15T19:47:25Z<p>John Creighton: /* Intro */</p>
<hr />
<div>== Intro ==<br />
As mentioned in [[RocketDesignSpreadsheet#Section_2_Oxidizer_Tank_Structure_Mass]], there are some important dimensionless quantities which determine how tanks weight scales with volume:<br />
<br />
== (Maxium Circumfrance)/(Volume^(1/3)) ==<br />
<br />
The point of the tank with maximum circumference will be near the weakest cross section of the tank with regards to pressure containment (see [[Pressure_Vessel]]). Strictly speaking for reasons of ease of computation, we take a cross section, near the weakest section, and account for any possible error by including a factor of safety. For instance, in the case of a cylinder, the plane of the cross section is taken to be parallel to the axis of the cylinder. If this is not the weakest cross section with regards to pressure containment it will be close. <br />
<br />
==== Sphere ====<br />
<br />
In the case of a sphere all cross sections, though the center of the sphere have the same circumference.<br />
<br />
<math>C=2*\pi*r</math><br />
<br />
The volume of a sphere is given by:<br />
<br />
<math>V=(4/3)\pi*r^3</math><br />
<br />
Therefore:<br />
<br />
<math>\frac{C}{V^(1/3)}=\frac{2*\pi}{((4/3)\pi)^{1/3}} \approx 3.9</math><br />
<br />
==== Cylinder ====<br />
<br />
As mentioned above the maximum circumference of a cylinder will occur near the cross section, which is parallel to the axis of the cylinder. <br />
<br />
If h is the height of the cylinder and r is the radius then, the circumfrance of this cross section, will go around half a sphere twice (at the ends) and along the tank height twice. Therefore:<br />
<br />
<math>C=2h+2*\pi*r</math><br />
<br />
The volume of this cylinder, will be equal to the cross sectional area multiplied by the hieght plus the area at the ends (two half spheres), which gives.<br />
<br />
<math>V=\pi*r^2*h+4/3*pi*r^3</math><br />
<br />
Now if the hight is some multiple (n) of the radius we get:<br />
<br />
<math>C=2(n+\pi)r</math><br />
<br />
<math>V=\pi*r^2*n*r+4/3*pi*r^3=\pi(n+(4/3))r^3(/math><br />
<br />
Therefore:<br />
<br />
<math>(C/V^(1/3))=\frac{2(n+\pi)}{(\pi(n+(4/3)))^(1/3)}</math> <br />
<br />
== (Maximum Cross Sectional Area)/(Volume^(2/3)) ==<br />
The point of the tank with maximum circumference will be near the weakest cross section of the tank with regards to pressure containment (see [[Pressure_Vessel]]). Strictly speaking for reasons of ease of computation, we take a cross section, near the weakest section, and account for any possible error by including a factor of safety. For instance, in the case of a cylinder, the plane of the cross section is taken to be parallel to the axis of the cylinder. If this is not the weakest cross section with regards to pressure containment it will be close.<br />
<br />
==== Sphere ====<br />
<br />
The cross sectional area of any cross section though the center of the sphere is the same regardless of the cross section and is given by:<br />
<br />
<math>C=\pi r^2</math><br />
<br />
The volume of a sphere is:<br />
<br />
<math>V=(4/3)\pi*r^3</math><br />
<br />
Therefore<br />
<br />
<math>(C/V^(2/3))=\frac{\pi}{((4/3)\pi)^{2/3}}</math><br />
<br />
==== Cylinder ====<br />
<br />
As mentioned above the maximum cross sectional area of a cylinder will occur near the cross section, which is parallel to the axis of the cylinder. The cross sectional area will consist of two parts, the main part of the cylinder which will have an area of 2*r*h, and the two half spheres on the end will have an area of pi*r^2 This gives:<br />
<br />
<math>A_c=2 \ r \ h+\pi r^2</math><br />
<br />
and if h=nr then we get:<br />
<br />
<math>A_c=2 \ r \ h+\pi r^2</math><br />
<br />
<math>A_c=2 \ r \ n \ r+\pi r^2=(2n+pi)r^2</math><br />
<br />
And, therefore<br />
<br />
<math>(C/V^(2/3))=\frac{(2n+pi)}{((4/3)\pi)^{2/3)}</math><br />
<br />
== (Surface Area)/(Volume^(2/3)) ==<br />
<br />
The total mass of the tank, is the surface area multiplied by the wall thickness, multiplied by the density of the material.<br />
<br />
==== Sphere m ====<br />
<br />
The surface area of a sphere is given by:<br />
<br />
<math>A_s=4 \pi r^2</math><br />
<br />
Therefore:<br />
<br />
<math>(A_s)/V^(2/3)=\frac{4 \pi}{((4/3)\pi)^{2/3)}</math><br />
<br />
==== Cylinder ====<br />
<br />
The surface area for the main part of the cylinder will be the circumference (2 pi r) multiplied by the height (h). The ends are two half spheres and will have a surface area of 4 pi r^2. Therefore<br />
<br />
<math>A_s=2 \pi r \ h + 4 \pi r^2</math><br />
<br />
and if h=n r then:<br />
<br />
<math>A_s=2 \pi r \ n \ r + 4 \pi r^2=\pi(n +r)r^2</math><br />
<br />
which gives:<br />
<br />
<math>(A_s)/(V^2/3)=\frac{\pi(n +r)}{((4/3)\pi)^{2/3)}</math></div>John Creightonhttp://wiki.newmars.com/index.php?title=Tank_geometries&diff=788Tank geometries2009-08-15T19:46:40Z<p>John Creighton: /* (Surface Area)/(Volume^(2/3) */</p>
<hr />
<div>== Intro ==<br />
As mentioned in [[RocketDesignSpreadsheet#Section_2_Oxidizer_Tank_Structure_Mass]] stated here, there are some important dimensionless quantities which determine how tanks scale with size:<br />
<br />
== (Maxium Circumfrance)/(Volume^(1/3)) ==<br />
<br />
The point of the tank with maximum circumference will be near the weakest cross section of the tank with regards to pressure containment (see [[Pressure_Vessel]]). Strictly speaking for reasons of ease of computation, we take a cross section, near the weakest section, and account for any possible error by including a factor of safety. For instance, in the case of a cylinder, the plane of the cross section is taken to be parallel to the axis of the cylinder. If this is not the weakest cross section with regards to pressure containment it will be close. <br />
<br />
==== Sphere ====<br />
<br />
In the case of a sphere all cross sections, though the center of the sphere have the same circumference.<br />
<br />
<math>C=2*\pi*r</math><br />
<br />
The volume of a sphere is given by:<br />
<br />
<math>V=(4/3)\pi*r^3</math><br />
<br />
Therefore:<br />
<br />
<math>\frac{C}{V^(1/3)}=\frac{2*\pi}{((4/3)\pi)^{1/3}} \approx 3.9</math><br />
<br />
==== Cylinder ====<br />
<br />
As mentioned above the maximum circumference of a cylinder will occur near the cross section, which is parallel to the axis of the cylinder. <br />
<br />
If h is the height of the cylinder and r is the radius then, the circumfrance of this cross section, will go around half a sphere twice (at the ends) and along the tank height twice. Therefore:<br />
<br />
<math>C=2h+2*\pi*r</math><br />
<br />
The volume of this cylinder, will be equal to the cross sectional area multiplied by the hieght plus the area at the ends (two half spheres), which gives.<br />
<br />
<math>V=\pi*r^2*h+4/3*pi*r^3</math><br />
<br />
Now if the hight is some multiple (n) of the radius we get:<br />
<br />
<math>C=2(n+\pi)r</math><br />
<br />
<math>V=\pi*r^2*n*r+4/3*pi*r^3=\pi(n+(4/3))r^3(/math><br />
<br />
Therefore:<br />
<br />
<math>(C/V^(1/3))=\frac{2(n+\pi)}{(\pi(n+(4/3)))^(1/3)}</math> <br />
<br />
== (Maximum Cross Sectional Area)/(Volume^(2/3)) ==<br />
The point of the tank with maximum circumference will be near the weakest cross section of the tank with regards to pressure containment (see [[Pressure_Vessel]]). Strictly speaking for reasons of ease of computation, we take a cross section, near the weakest section, and account for any possible error by including a factor of safety. For instance, in the case of a cylinder, the plane of the cross section is taken to be parallel to the axis of the cylinder. If this is not the weakest cross section with regards to pressure containment it will be close.<br />
<br />
==== Sphere ====<br />
<br />
The cross sectional area of any cross section though the center of the sphere is the same regardless of the cross section and is given by:<br />
<br />
<math>C=\pi r^2</math><br />
<br />
The volume of a sphere is:<br />
<br />
<math>V=(4/3)\pi*r^3</math><br />
<br />
Therefore<br />
<br />
<math>(C/V^(2/3))=\frac{\pi}{((4/3)\pi)^{2/3}}</math><br />
<br />
==== Cylinder ====<br />
<br />
As mentioned above the maximum cross sectional area of a cylinder will occur near the cross section, which is parallel to the axis of the cylinder. The cross sectional area will consist of two parts, the main part of the cylinder which will have an area of 2*r*h, and the two half spheres on the end will have an area of pi*r^2 This gives:<br />
<br />
<math>A_c=2 \ r \ h+\pi r^2</math><br />
<br />
and if h=nr then we get:<br />
<br />
<math>A_c=2 \ r \ h+\pi r^2</math><br />
<br />
<math>A_c=2 \ r \ n \ r+\pi r^2=(2n+pi)r^2</math><br />
<br />
And, therefore<br />
<br />
<math>(C/V^(2/3))=\frac{(2n+pi)}{((4/3)\pi)^{2/3)}</math><br />
<br />
== (Surface Area)/(Volume^(2/3)) ==<br />
<br />
The total mass of the tank, is the surface area multiplied by the wall thickness, multiplied by the density of the material.<br />
<br />
==== Sphere m ====<br />
<br />
The surface area of a sphere is given by:<br />
<br />
<math>A_s=4 \pi r^2</math><br />
<br />
Therefore:<br />
<br />
<math>(A_s)/V^(2/3)=\frac{4 \pi}{((4/3)\pi)^{2/3)}</math><br />
<br />
==== Cylinder ====<br />
<br />
The surface area for the main part of the cylinder will be the circumference (2 pi r) multiplied by the height (h). The ends are two half spheres and will have a surface area of 4 pi r^2. Therefore<br />
<br />
<math>A_s=2 \pi r \ h + 4 \pi r^2</math><br />
<br />
and if h=n r then:<br />
<br />
<math>A_s=2 \pi r \ n \ r + 4 \pi r^2=\pi(n +r)r^2</math><br />
<br />
which gives:<br />
<br />
<math>(A_s)/(V^2/3)=\frac{\pi(n +r)}{((4/3)\pi)^{2/3)}</math></div>John Creightonhttp://wiki.newmars.com/index.php?title=Tank_geometries&diff=787Tank geometries2009-08-15T19:45:50Z<p>John Creighton: New page: == Intro == As mentioned in RocketDesignSpreadsheet#Section_2_Oxidizer_Tank_Structure_Mass stated here, there are some important dimensionless quantities which determine how tanks scal...</p>
<hr />
<div>== Intro ==<br />
As mentioned in [[RocketDesignSpreadsheet#Section_2_Oxidizer_Tank_Structure_Mass]] stated here, there are some important dimensionless quantities which determine how tanks scale with size:<br />
<br />
== (Maxium Circumfrance)/(Volume^(1/3)) ==<br />
<br />
The point of the tank with maximum circumference will be near the weakest cross section of the tank with regards to pressure containment (see [[Pressure_Vessel]]). Strictly speaking for reasons of ease of computation, we take a cross section, near the weakest section, and account for any possible error by including a factor of safety. For instance, in the case of a cylinder, the plane of the cross section is taken to be parallel to the axis of the cylinder. If this is not the weakest cross section with regards to pressure containment it will be close. <br />
<br />
==== Sphere ====<br />
<br />
In the case of a sphere all cross sections, though the center of the sphere have the same circumference.<br />
<br />
<math>C=2*\pi*r</math><br />
<br />
The volume of a sphere is given by:<br />
<br />
<math>V=(4/3)\pi*r^3</math><br />
<br />
Therefore:<br />
<br />
<math>\frac{C}{V^(1/3)}=\frac{2*\pi}{((4/3)\pi)^{1/3}} \approx 3.9</math><br />
<br />
==== Cylinder ====<br />
<br />
As mentioned above the maximum circumference of a cylinder will occur near the cross section, which is parallel to the axis of the cylinder. <br />
<br />
If h is the height of the cylinder and r is the radius then, the circumfrance of this cross section, will go around half a sphere twice (at the ends) and along the tank height twice. Therefore:<br />
<br />
<math>C=2h+2*\pi*r</math><br />
<br />
The volume of this cylinder, will be equal to the cross sectional area multiplied by the hieght plus the area at the ends (two half spheres), which gives.<br />
<br />
<math>V=\pi*r^2*h+4/3*pi*r^3</math><br />
<br />
Now if the hight is some multiple (n) of the radius we get:<br />
<br />
<math>C=2(n+\pi)r</math><br />
<br />
<math>V=\pi*r^2*n*r+4/3*pi*r^3=\pi(n+(4/3))r^3(/math><br />
<br />
Therefore:<br />
<br />
<math>(C/V^(1/3))=\frac{2(n+\pi)}{(\pi(n+(4/3)))^(1/3)}</math> <br />
<br />
== (Maximum Cross Sectional Area)/(Volume^(2/3)) ==<br />
The point of the tank with maximum circumference will be near the weakest cross section of the tank with regards to pressure containment (see [[Pressure_Vessel]]). Strictly speaking for reasons of ease of computation, we take a cross section, near the weakest section, and account for any possible error by including a factor of safety. For instance, in the case of a cylinder, the plane of the cross section is taken to be parallel to the axis of the cylinder. If this is not the weakest cross section with regards to pressure containment it will be close.<br />
<br />
==== Sphere ====<br />
<br />
The cross sectional area of any cross section though the center of the sphere is the same regardless of the cross section and is given by:<br />
<br />
<math>C=\pi r^2</math><br />
<br />
The volume of a sphere is:<br />
<br />
<math>V=(4/3)\pi*r^3</math><br />
<br />
Therefore<br />
<br />
<math>(C/V^(2/3))=\frac{\pi}{((4/3)\pi)^{2/3}}</math><br />
<br />
==== Cylinder ====<br />
<br />
As mentioned above the maximum cross sectional area of a cylinder will occur near the cross section, which is parallel to the axis of the cylinder. The cross sectional area will consist of two parts, the main part of the cylinder which will have an area of 2*r*h, and the two half spheres on the end will have an area of pi*r^2 This gives:<br />
<br />
<math>A_c=2 \ r \ h+\pi r^2</math><br />
<br />
and if h=nr then we get:<br />
<br />
<math>A_c=2 \ r \ h+\pi r^2</math><br />
<br />
<math>A_c=2 \ r \ n \ r+\pi r^2=(2n+pi)r^2</math><br />
<br />
And, therefore<br />
<br />
<math>(C/V^(2/3))=\frac{(2n+pi)}{((4/3)\pi)^{2/3)}</math><br />
<br />
== (Surface Area)/(Volume^(2/3) ==<br />
<br />
The total mass of the tank, is the surface area multiplied by the wall thickness, multiplied by the density of the material.<br />
<br />
==== Sphere m ====<br />
<br />
The surface area of a sphere is given by:<br />
<br />
<math>A_s=4 \pi r^2</math><br />
<br />
Therefore:<br />
<br />
<math>(A_s)/V^(2/3)=\frac{4 \pi}{((4/3)\pi)^{2/3)}</math><br />
<br />
==== Cylinder ====<br />
<br />
The surface area for the main part of the cylinder will be the circumference (2 pi r) multiplied by the height (h). The ends are two half spheres and will have a surface area of 4 pi r^2. Therefore<br />
<br />
<math>A_s=2 \pi r \ h + 4 \pi r^2</math><br />
<br />
and if h=n r then:<br />
<br />
<math>A_s=2 \pi r \ n \ r + 4 \pi r^2=\pi(n +r)r^2</math><br />
<br />
which gives:<br />
<br />
<math>(A_s)/(V^2/3)=\frac{\pi(n +r)}{((4/3)\pi)^{2/3)}</math></div>John Creightonhttp://wiki.newmars.com/index.php?title=SIL&diff=786SIL2009-08-14T21:41:54Z<p>John Creighton: /* LOPA (Layers of Protection Analysis */</p>
<hr />
<div>== Introduction ==<br />
Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple terms, SIL is a measurement of performance required for a Safety Instrumented Function (SIF). Bellow shows a table [1] which gives, the SILL level, PDF (probability of failure on demand) and the RDF (risk reduction factor). <br />
<br />
<br />
{| class="wikitable"<br />
|-<br />
! SIL<br />
! PFD<br />
! RRF<br />
|-<br />
| 1<br />
| 0.1-0.01<br />
| 10-100<br />
|-<br />
| 2<br />
| 0.01-0.001<br />
| 100-1000<br />
|-<br />
| 3<br />
| 0.001-0.0001<br />
| 1000-10,000<br />
|-<br />
| 4<br />
| 0.0001-0.00001<br />
| 10,000-100,000<br />
|}<br />
<br />
== Selection of SIL ==<br />
<br />
There are many methods to select the safety Integrity Level, these include, Risk Matrix, Risk Graph, Layers of Protection Analysis (LOPA). With regards to LOC NASA has considerably lower levels of safety then in comparison to industry because, the crew are not considered civilians, and the gains of space exploration to be with the risk at least in terms of LOC. <br />
<br />
Bellow is what is a graph of what is considered acceptable and intolerable in terms of fatalities by (HSE Books 2001)[2]. Of course these graphs would very depending on the size of the population and the significance of the endeavor. <br />
<br />
[[Image:Fatality_graph.JPG]]<br />
<br />
== Sill Selection Matrices ==<br />
<br />
A SIL matrix tells us how reliable a safety function must be given the likely hood and the severity of an event. Bellow a 3D sill selection matrix is shown [3]. If you can achieve the desired safety levels independently of the other layers then a 3D sill section matrix may not be necessary. However, there may be a maximum amount of reliability we can achieve from a given safety function and therefore we must consider how each layer of protection contributes to the reliability of the overall system. For instance, it was suggested in the Augastine commission that an abort system only reduces your LOC (Loss of crew) by about a factor of 10. If this does not give the required reliability in terms of loss of crew, then we must consider the reliability of the other layers of the system. <br />
<br />
[[Image:SILL_Matrix.JPG]]<br />
<br />
Generally a separate SIL selection matrix is done for each type of consequence, these can include, Loss of mission, Loss of Crew, Environmental impact and Damage to the reputation of the organization. Generally each type of consequence is considered separately and it is the type of consequence which requires the greatest level of safety which drives the design.<br />
<br />
== LOPA (Layers of Protection Analysis) ==<br />
<br />
== Risk Graphs ==<br />
<br />
<br />
[[Image:Risk_Graph.JPG]]<br />
<br />
Figure from: <br />
<br />
Different SIL (Safety Integrity Level) Selection Techniques<br />
Can Yield Significantly Different Answers<br />
By Paul Gruhn, PE, CFSE<br />
President<br />
L&M Engineering<br />
Houston, TX<br />
[3]<br />
<br />
== Nuclear Power and Space ==<br />
<br />
In order to reduce the consequences of a failure, NASA uses hard ceramics, to minimize the environmental impact of launch failure. With regards to nuclear propulsion, it is general considered a much greater environmental risk if the reactor is turned on before the rocket reaches a stable orbit, then if it is turned on after it reaches a stable orbit. The required reliability necessary for such a consequence is a matter of debate but no doubt the necessary safeguards to prevent this incident will likely add significantly to the weight and reliability of the overall rocket. <br />
<br />
== References ==<br />
<br />
[1] http://en.wikipedia.org/wiki/Safety_Integrity_Level<br />
<br />
[2] - http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf - Methods of Determining Safety Integrity Level (SIL)<br />
Requirements - Pros and Cons<br />
by W G Gulland (4-sight Consulting)<br />
<br />
[3] http://www.isa.org/Content/Microsites838/Safety_Division/Home818/ISA_2004_Safety_Papers/Different_SIL_Selection_Techniques_Can_Yield_Different_Answers.pdf <br />
<br />
<br />
[4] http://www.iceweb.com.au/sis/target_sis.htm - Techniques for Assigning A Target Safety Integrity Level Angela E. Summers, Ph.D. This paper was published in ISA Transactions 37 (1998) 95-104.</div>John Creightonhttp://wiki.newmars.com/index.php?title=SIL&diff=785SIL2009-08-14T21:41:04Z<p>John Creighton: </p>
<hr />
<div>== Introduction ==<br />
Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple terms, SIL is a measurement of performance required for a Safety Instrumented Function (SIF). Bellow shows a table [1] which gives, the SILL level, PDF (probability of failure on demand) and the RDF (risk reduction factor). <br />
<br />
<br />
{| class="wikitable"<br />
|-<br />
! SIL<br />
! PFD<br />
! RRF<br />
|-<br />
| 1<br />
| 0.1-0.01<br />
| 10-100<br />
|-<br />
| 2<br />
| 0.01-0.001<br />
| 100-1000<br />
|-<br />
| 3<br />
| 0.001-0.0001<br />
| 1000-10,000<br />
|-<br />
| 4<br />
| 0.0001-0.00001<br />
| 10,000-100,000<br />
|}<br />
<br />
== Selection of SIL ==<br />
<br />
There are many methods to select the safety Integrity Level, these include, Risk Matrix, Risk Graph, Layers of Protection Analysis (LOPA). With regards to LOC NASA has considerably lower levels of safety then in comparison to industry because, the crew are not considered civilians, and the gains of space exploration to be with the risk at least in terms of LOC. <br />
<br />
Bellow is what is a graph of what is considered acceptable and intolerable in terms of fatalities by (HSE Books 2001)[2]. Of course these graphs would very depending on the size of the population and the significance of the endeavor. <br />
<br />
[[Image:Fatality_graph.JPG]]<br />
<br />
== Sill Selection Matrices ==<br />
<br />
A SIL matrix tells us how reliable a safety function must be given the likely hood and the severity of an event. Bellow a 3D sill selection matrix is shown [3]. If you can achieve the desired safety levels independently of the other layers then a 3D sill section matrix may not be necessary. However, there may be a maximum amount of reliability we can achieve from a given safety function and therefore we must consider how each layer of protection contributes to the reliability of the overall system. For instance, it was suggested in the Augastine commission that an abort system only reduces your LOC (Loss of crew) by about a factor of 10. If this does not give the required reliability in terms of loss of crew, then we must consider the reliability of the other layers of the system. <br />
<br />
[[Image:SILL_Matrix.JPG]]<br />
<br />
Generally a separate SIL selection matrix is done for each type of consequence, these can include, Loss of mission, Loss of Crew, Environmental impact and Damage to the reputation of the organization. Generally each type of consequence is considered separately and it is the type of consequence which requires the greatest level of safety which drives the design.<br />
<br />
== LOPA (Layers of Protection Analysis ==<br />
<br />
== Risk Graphs ==<br />
<br />
<br />
[[Image:Risk_Graph.JPG]]<br />
<br />
Figure from: <br />
<br />
Different SIL (Safety Integrity Level) Selection Techniques<br />
Can Yield Significantly Different Answers<br />
By Paul Gruhn, PE, CFSE<br />
President<br />
L&M Engineering<br />
Houston, TX<br />
[3]<br />
<br />
== Nuclear Power and Space ==<br />
<br />
In order to reduce the consequences of a failure, NASA uses hard ceramics, to minimize the environmental impact of launch failure. With regards to nuclear propulsion, it is general considered a much greater environmental risk if the reactor is turned on before the rocket reaches a stable orbit, then if it is turned on after it reaches a stable orbit. The required reliability necessary for such a consequence is a matter of debate but no doubt the necessary safeguards to prevent this incident will likely add significantly to the weight and reliability of the overall rocket. <br />
<br />
== References ==<br />
<br />
[1] http://en.wikipedia.org/wiki/Safety_Integrity_Level<br />
<br />
[2] - http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf - Methods of Determining Safety Integrity Level (SIL)<br />
Requirements - Pros and Cons<br />
by W G Gulland (4-sight Consulting)<br />
<br />
[3] http://www.isa.org/Content/Microsites838/Safety_Division/Home818/ISA_2004_Safety_Papers/Different_SIL_Selection_Techniques_Can_Yield_Different_Answers.pdf <br />
<br />
<br />
[4] http://www.iceweb.com.au/sis/target_sis.htm - Techniques for Assigning A Target Safety Integrity Level Angela E. Summers, Ph.D. This paper was published in ISA Transactions 37 (1998) 95-104.</div>John Creightonhttp://wiki.newmars.com/index.php?title=SIL&diff=784SIL2009-08-14T21:35:47Z<p>John Creighton: /* Introduction */</p>
<hr />
<div>== Introduction ==<br />
Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple terms, SIL is a measurement of performance required for a Safety Instrumented Function (SIF). Bellow shows a table [1] which gives, the SILL level, PDF (probability of failure on demand) and the RDF (risk reduction factor). <br />
<br />
<br />
{| class="wikitable"<br />
|-<br />
! SIL<br />
! PFD<br />
! RRF<br />
|-<br />
| 1<br />
| 0.1-0.01<br />
| 10-100<br />
|-<br />
| 2<br />
| 0.01-0.001<br />
| 100-1000<br />
|-<br />
| 3<br />
| 0.001-0.0001<br />
| 1000-10,000<br />
|-<br />
| 4<br />
| 0.0001-0.00001<br />
| 10,000-100,000<br />
|}<br />
<br />
== Selection of SIL ==<br />
<br />
There are many methods to select the safety Integrity Level, these include, Risk Matrix, Risk Graph, Layers of Protection Analysis (LOPA). With regards to LOC NASA has considerably lower levels of safety then in comparison to industry because, the crew are not considered civilians, and the gains of space exploration to be with the risk at least in terms of LOC. <br />
<br />
Bellow is what is a graph of what is considered acceptable and intolerable in terms of fatalities by (HSE Books 2001)[2]. Of course these graphs would very depending on the size of the population and the significance of the endeavor. <br />
<br />
[[Image:Fatality_graph.JPG]]<br />
<br />
== Sill Selection Matrices ==<br />
<br />
A SIL matrix tells us how reliable a safety function must be given the likely hood and the severity of an event. Bellow a 3D sill selection matrix is shown [3]. If you can achieve the desired safety levels independently of the other layers then a 3D sill section matrix may not be necessary. However, there may be a maximum amount of reliability we can achieve from a given safety function and therefore we must consider how each layer of protection contributes to the reliability of the overall system. For instance, it was suggested in the Augastine commission that an abort system only reduces your LOC (Loss of crew) by about a factor of 10. If this does not give the required reliability in terms of loss of crew, then we must consider the reliability of the other layers of the system. <br />
<br />
[[Image:SILL_Matrix.JPG]]<br />
<br />
Generally a separate SIL selection matrix is done for each type of consequence, these can include, Loss of mission, Loss of Crew, Environmental impact and Damage to the reputation of the organization. Generally each type of consequence is considered separately and it is the type of consequence which requires the greatest level of safety which drives the design.<br />
<br />
== LOPA (Layers of Protection Analysis ==<br />
<br />
== Risk Graphs ==<br />
<br />
<br />
[[Image:Risk_Graph.JPG]]<br />
<br />
Figure from: <br />
<br />
Different SIL (Safety Integrity Level) Selection Techniques<br />
Can Yield Significantly Different Answers<br />
By Paul Gruhn, PE, CFSE<br />
President<br />
L&M Engineering<br />
Houston, TX<br />
[3]<br />
<br />
== Refferences ==<br />
<br />
[1] http://en.wikipedia.org/wiki/Safety_Integrity_Level<br />
<br />
[2] - http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf - Methods of Determining Safety Integrity Level (SIL)<br />
Requirements - Pros and Cons<br />
by W G Gulland (4-sight Consulting)<br />
<br />
[3] http://www.isa.org/Content/Microsites838/Safety_Division/Home818/ISA_2004_Safety_Papers/Different_SIL_Selection_Techniques_Can_Yield_Different_Answers.pdf <br />
<br />
<br />
[4] http://www.iceweb.com.au/sis/target_sis.htm - Techniques for Assigning A Target Safety Integrity Level Angela E. Summers, Ph.D. This paper was published in ISA Transactions 37 (1998) 95-104.</div>John Creightonhttp://wiki.newmars.com/index.php?title=SIL&diff=783SIL2009-08-14T21:34:33Z<p>John Creighton: /* Selection of SIL */</p>
<hr />
<div>== Introduction ==<br />
Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple terms, SIL is a measurement of performance required for a Safety Instrumented Function (SIF).<br />
<br />
<br />
{| class="wikitable"<br />
|-<br />
! SIL<br />
! PFD<br />
! RRF<br />
|-<br />
| 1<br />
| 0.1-0.01<br />
| 10-100<br />
|-<br />
| 2<br />
| 0.01-0.001<br />
| 100-1000<br />
|-<br />
| 3<br />
| 0.001-0.0001<br />
| 1000-10,000<br />
|-<br />
| 4<br />
| 0.0001-0.00001<br />
| 10,000-100,000<br />
|}<br />
<br />
== Selection of SIL ==<br />
<br />
There are many methods to select the safety Integrity Level, these include, Risk Matrix, Risk Graph, Layers of Protection Analysis (LOPA). With regards to LOC NASA has considerably lower levels of safety then in comparison to industry because, the crew are not considered civilians, and the gains of space exploration to be with the risk at least in terms of LOC. <br />
<br />
Bellow is what is a graph of what is considered acceptable and intolerable in terms of fatalities by (HSE Books 2001)[2]. Of course these graphs would very depending on the size of the population and the significance of the endeavor. <br />
<br />
[[Image:Fatality_graph.JPG]]<br />
<br />
== Sill Selection Matrices ==<br />
<br />
A SIL matrix tells us how reliable a safety function must be given the likely hood and the severity of an event. Bellow a 3D sill selection matrix is shown [3]. If you can achieve the desired safety levels independently of the other layers then a 3D sill section matrix may not be necessary. However, there may be a maximum amount of reliability we can achieve from a given safety function and therefore we must consider how each layer of protection contributes to the reliability of the overall system. For instance, it was suggested in the Augastine commission that an abort system only reduces your LOC (Loss of crew) by about a factor of 10. If this does not give the required reliability in terms of loss of crew, then we must consider the reliability of the other layers of the system. <br />
<br />
[[Image:SILL_Matrix.JPG]]<br />
<br />
Generally a separate SIL selection matrix is done for each type of consequence, these can include, Loss of mission, Loss of Crew, Environmental impact and Damage to the reputation of the organization. Generally each type of consequence is considered separately and it is the type of consequence which requires the greatest level of safety which drives the design.<br />
<br />
== LOPA (Layers of Protection Analysis ==<br />
<br />
== Risk Graphs ==<br />
<br />
<br />
[[Image:Risk_Graph.JPG]]<br />
<br />
Figure from: <br />
<br />
Different SIL (Safety Integrity Level) Selection Techniques<br />
Can Yield Significantly Different Answers<br />
By Paul Gruhn, PE, CFSE<br />
President<br />
L&M Engineering<br />
Houston, TX<br />
[3]<br />
<br />
== Refferences ==<br />
<br />
[1] http://en.wikipedia.org/wiki/Safety_Integrity_Level<br />
<br />
[2] - http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf - Methods of Determining Safety Integrity Level (SIL)<br />
Requirements - Pros and Cons<br />
by W G Gulland (4-sight Consulting)<br />
<br />
[3] http://www.isa.org/Content/Microsites838/Safety_Division/Home818/ISA_2004_Safety_Papers/Different_SIL_Selection_Techniques_Can_Yield_Different_Answers.pdf <br />
<br />
<br />
[4] http://www.iceweb.com.au/sis/target_sis.htm - Techniques for Assigning A Target Safety Integrity Level Angela E. Summers, Ph.D. This paper was published in ISA Transactions 37 (1998) 95-104.</div>John Creightonhttp://wiki.newmars.com/index.php?title=SIL&diff=782SIL2009-08-14T21:34:12Z<p>John Creighton: /* Sill Selection Matrices */</p>
<hr />
<div>== Introduction ==<br />
Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple terms, SIL is a measurement of performance required for a Safety Instrumented Function (SIF).<br />
<br />
<br />
{| class="wikitable"<br />
|-<br />
! SIL<br />
! PFD<br />
! RRF<br />
|-<br />
| 1<br />
| 0.1-0.01<br />
| 10-100<br />
|-<br />
| 2<br />
| 0.01-0.001<br />
| 100-1000<br />
|-<br />
| 3<br />
| 0.001-0.0001<br />
| 1000-10,000<br />
|-<br />
| 4<br />
| 0.0001-0.00001<br />
| 10,000-100,000<br />
|}<br />
<br />
== Selection of SIL ==<br />
<br />
There are many methods to select the safety Integrity Level, these include, Risk Matrix, Risk Graph, Layers of Protection Analysis (LOPA). With regards to LOC NASA has considerably lower levels of safety then in comparison to industry because, the crew are not considered civilians, and the gains of space exploration to be with the risk at least in terms of LOC. <br />
<br />
Bellow is what is a graph of what is considered acceptable and intolerable in terms of fatalities by (HSE Books 2001)[1]. Of course these graphs would very depending on the size of the population and the significance of the endeavor. <br />
<br />
[[Image:Fatality_graph.JPG]] <br />
<br />
== Sill Selection Matrices ==<br />
<br />
A SIL matrix tells us how reliable a safety function must be given the likely hood and the severity of an event. Bellow a 3D sill selection matrix is shown [3]. If you can achieve the desired safety levels independently of the other layers then a 3D sill section matrix may not be necessary. However, there may be a maximum amount of reliability we can achieve from a given safety function and therefore we must consider how each layer of protection contributes to the reliability of the overall system. For instance, it was suggested in the Augastine commission that an abort system only reduces your LOC (Loss of crew) by about a factor of 10. If this does not give the required reliability in terms of loss of crew, then we must consider the reliability of the other layers of the system. <br />
<br />
[[Image:SILL_Matrix.JPG]]<br />
<br />
Generally a separate SIL selection matrix is done for each type of consequence, these can include, Loss of mission, Loss of Crew, Environmental impact and Damage to the reputation of the organization. Generally each type of consequence is considered separately and it is the type of consequence which requires the greatest level of safety which drives the design.<br />
<br />
== LOPA (Layers of Protection Analysis ==<br />
<br />
== Risk Graphs ==<br />
<br />
<br />
[[Image:Risk_Graph.JPG]]<br />
<br />
Figure from: <br />
<br />
Different SIL (Safety Integrity Level) Selection Techniques<br />
Can Yield Significantly Different Answers<br />
By Paul Gruhn, PE, CFSE<br />
President<br />
L&M Engineering<br />
Houston, TX<br />
[3]<br />
<br />
== Refferences ==<br />
<br />
[1] http://en.wikipedia.org/wiki/Safety_Integrity_Level<br />
<br />
[2] - http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf - Methods of Determining Safety Integrity Level (SIL)<br />
Requirements - Pros and Cons<br />
by W G Gulland (4-sight Consulting)<br />
<br />
[3] http://www.isa.org/Content/Microsites838/Safety_Division/Home818/ISA_2004_Safety_Papers/Different_SIL_Selection_Techniques_Can_Yield_Different_Answers.pdf <br />
<br />
<br />
[4] http://www.iceweb.com.au/sis/target_sis.htm - Techniques for Assigning A Target Safety Integrity Level Angela E. Summers, Ph.D. This paper was published in ISA Transactions 37 (1998) 95-104.</div>John Creightonhttp://wiki.newmars.com/index.php?title=SIL&diff=781SIL2009-08-14T21:32:49Z<p>John Creighton: /* Risk Graphs */</p>
<hr />
<div>== Introduction ==<br />
Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple terms, SIL is a measurement of performance required for a Safety Instrumented Function (SIF).<br />
<br />
<br />
{| class="wikitable"<br />
|-<br />
! SIL<br />
! PFD<br />
! RRF<br />
|-<br />
| 1<br />
| 0.1-0.01<br />
| 10-100<br />
|-<br />
| 2<br />
| 0.01-0.001<br />
| 100-1000<br />
|-<br />
| 3<br />
| 0.001-0.0001<br />
| 1000-10,000<br />
|-<br />
| 4<br />
| 0.0001-0.00001<br />
| 10,000-100,000<br />
|}<br />
<br />
== Selection of SIL ==<br />
<br />
There are many methods to select the safety Integrity Level, these include, Risk Matrix, Risk Graph, Layers of Protection Analysis (LOPA). With regards to LOC NASA has considerably lower levels of safety then in comparison to industry because, the crew are not considered civilians, and the gains of space exploration to be with the risk at least in terms of LOC. <br />
<br />
Bellow is what is a graph of what is considered acceptable and intolerable in terms of fatalities by (HSE Books 2001)[1]. Of course these graphs would very depending on the size of the population and the significance of the endeavor. <br />
<br />
[[Image:Fatality_graph.JPG]] <br />
<br />
== Sill Selection Matrices ==<br />
<br />
A SIL matrix tells us how reliable a safety function must be given the likely hood and the severity of an event. Bellow a 3D sill selection matrix is shown. If you can achieve the desired safety levels independently of the other layers then a 3D sill section matrix may not be necessary. However, there may be a maximum amount of reliability we can achieve from a given safety function and therefore we must consider how each layer of protection contributes to the reliability of the overall system. For instance, it was suggested in the Augastine commission that an abort system only reduces your LOC (Loss of crew) by about a factor of 10. If this does not give the required reliability in terms of loss of crew, then we must consider the reliability of the other layers of the system. <br />
<br />
[[Image:SILL_Matrix.JPG]]<br />
<br />
Generally a separate SIL selection matrix is done for each type of consequence, these can include, Loss of mission, Loss of Crew, Environmental impact and Damage to the reputation of the organization. Generally each type of consequence is considered separately and it is the type of consequence which requires the greatest level of safety which drives the design.<br />
<br />
== LOPA (Layers of Protection Analysis ==<br />
<br />
== Risk Graphs ==<br />
<br />
<br />
[[Image:Risk_Graph.JPG]]<br />
<br />
Figure from: <br />
<br />
Different SIL (Safety Integrity Level) Selection Techniques<br />
Can Yield Significantly Different Answers<br />
By Paul Gruhn, PE, CFSE<br />
President<br />
L&M Engineering<br />
Houston, TX<br />
[3]<br />
<br />
== Refferences ==<br />
<br />
[1] http://en.wikipedia.org/wiki/Safety_Integrity_Level<br />
<br />
[2] - http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf - Methods of Determining Safety Integrity Level (SIL)<br />
Requirements - Pros and Cons<br />
by W G Gulland (4-sight Consulting)<br />
<br />
[3] http://www.isa.org/Content/Microsites838/Safety_Division/Home818/ISA_2004_Safety_Papers/Different_SIL_Selection_Techniques_Can_Yield_Different_Answers.pdf <br />
<br />
<br />
[4] http://www.iceweb.com.au/sis/target_sis.htm - Techniques for Assigning A Target Safety Integrity Level Angela E. Summers, Ph.D. This paper was published in ISA Transactions 37 (1998) 95-104.</div>John Creightonhttp://wiki.newmars.com/index.php?title=SIL&diff=780SIL2009-08-14T21:31:57Z<p>John Creighton: /* Refferences */</p>
<hr />
<div>== Introduction ==<br />
Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple terms, SIL is a measurement of performance required for a Safety Instrumented Function (SIF).<br />
<br />
<br />
{| class="wikitable"<br />
|-<br />
! SIL<br />
! PFD<br />
! RRF<br />
|-<br />
| 1<br />
| 0.1-0.01<br />
| 10-100<br />
|-<br />
| 2<br />
| 0.01-0.001<br />
| 100-1000<br />
|-<br />
| 3<br />
| 0.001-0.0001<br />
| 1000-10,000<br />
|-<br />
| 4<br />
| 0.0001-0.00001<br />
| 10,000-100,000<br />
|}<br />
<br />
== Selection of SIL ==<br />
<br />
There are many methods to select the safety Integrity Level, these include, Risk Matrix, Risk Graph, Layers of Protection Analysis (LOPA). With regards to LOC NASA has considerably lower levels of safety then in comparison to industry because, the crew are not considered civilians, and the gains of space exploration to be with the risk at least in terms of LOC. <br />
<br />
Bellow is what is a graph of what is considered acceptable and intolerable in terms of fatalities by (HSE Books 2001)[1]. Of course these graphs would very depending on the size of the population and the significance of the endeavor. <br />
<br />
[[Image:Fatality_graph.JPG]] <br />
<br />
== Sill Selection Matrices ==<br />
<br />
A SIL matrix tells us how reliable a safety function must be given the likely hood and the severity of an event. Bellow a 3D sill selection matrix is shown. If you can achieve the desired safety levels independently of the other layers then a 3D sill section matrix may not be necessary. However, there may be a maximum amount of reliability we can achieve from a given safety function and therefore we must consider how each layer of protection contributes to the reliability of the overall system. For instance, it was suggested in the Augastine commission that an abort system only reduces your LOC (Loss of crew) by about a factor of 10. If this does not give the required reliability in terms of loss of crew, then we must consider the reliability of the other layers of the system. <br />
<br />
[[Image:SILL_Matrix.JPG]]<br />
<br />
Generally a separate SIL selection matrix is done for each type of consequence, these can include, Loss of mission, Loss of Crew, Environmental impact and Damage to the reputation of the organization. Generally each type of consequence is considered separately and it is the type of consequence which requires the greatest level of safety which drives the design.<br />
<br />
== LOPA (Layers of Protection Analysis ==<br />
<br />
== Risk Graphs ==<br />
<br />
<br />
[[Image:Risk_Graph.JPG]]<br />
<br />
Figure from: <br />
<br />
Different SIL (Safety Integrity Level) Selection Techniques<br />
Can Yield Significantly Different Answers<br />
By Paul Gruhn, PE, CFSE<br />
President<br />
L&M Engineering<br />
Houston, TX<br />
pgruhn@landmengineering.com<br />
<br />
== Refferences ==<br />
<br />
[1] http://en.wikipedia.org/wiki/Safety_Integrity_Level<br />
<br />
[2] - http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf - Methods of Determining Safety Integrity Level (SIL)<br />
Requirements - Pros and Cons<br />
by W G Gulland (4-sight Consulting)<br />
<br />
[3] http://www.isa.org/Content/Microsites838/Safety_Division/Home818/ISA_2004_Safety_Papers/Different_SIL_Selection_Techniques_Can_Yield_Different_Answers.pdf <br />
<br />
<br />
[4] http://www.iceweb.com.au/sis/target_sis.htm - Techniques for Assigning A Target Safety Integrity Level Angela E. Summers, Ph.D. This paper was published in ISA Transactions 37 (1998) 95-104.</div>John Creightonhttp://wiki.newmars.com/index.php?title=SIL&diff=779SIL2009-08-14T21:31:03Z<p>John Creighton: /* Risk Graphs */</p>
<hr />
<div>== Introduction ==<br />
Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple terms, SIL is a measurement of performance required for a Safety Instrumented Function (SIF).<br />
<br />
<br />
{| class="wikitable"<br />
|-<br />
! SIL<br />
! PFD<br />
! RRF<br />
|-<br />
| 1<br />
| 0.1-0.01<br />
| 10-100<br />
|-<br />
| 2<br />
| 0.01-0.001<br />
| 100-1000<br />
|-<br />
| 3<br />
| 0.001-0.0001<br />
| 1000-10,000<br />
|-<br />
| 4<br />
| 0.0001-0.00001<br />
| 10,000-100,000<br />
|}<br />
<br />
== Selection of SIL ==<br />
<br />
There are many methods to select the safety Integrity Level, these include, Risk Matrix, Risk Graph, Layers of Protection Analysis (LOPA). With regards to LOC NASA has considerably lower levels of safety then in comparison to industry because, the crew are not considered civilians, and the gains of space exploration to be with the risk at least in terms of LOC. <br />
<br />
Bellow is what is a graph of what is considered acceptable and intolerable in terms of fatalities by (HSE Books 2001)[1]. Of course these graphs would very depending on the size of the population and the significance of the endeavor. <br />
<br />
[[Image:Fatality_graph.JPG]] <br />
<br />
== Sill Selection Matrices ==<br />
<br />
A SIL matrix tells us how reliable a safety function must be given the likely hood and the severity of an event. Bellow a 3D sill selection matrix is shown. If you can achieve the desired safety levels independently of the other layers then a 3D sill section matrix may not be necessary. However, there may be a maximum amount of reliability we can achieve from a given safety function and therefore we must consider how each layer of protection contributes to the reliability of the overall system. For instance, it was suggested in the Augastine commission that an abort system only reduces your LOC (Loss of crew) by about a factor of 10. If this does not give the required reliability in terms of loss of crew, then we must consider the reliability of the other layers of the system. <br />
<br />
[[Image:SILL_Matrix.JPG]]<br />
<br />
Generally a separate SIL selection matrix is done for each type of consequence, these can include, Loss of mission, Loss of Crew, Environmental impact and Damage to the reputation of the organization. Generally each type of consequence is considered separately and it is the type of consequence which requires the greatest level of safety which drives the design.<br />
<br />
== LOPA (Layers of Protection Analysis ==<br />
<br />
== Risk Graphs ==<br />
<br />
<br />
[[Image:Risk_Graph.JPG]]<br />
<br />
Figure from: <br />
<br />
Different SIL (Safety Integrity Level) Selection Techniques<br />
Can Yield Significantly Different Answers<br />
By Paul Gruhn, PE, CFSE<br />
President<br />
L&M Engineering<br />
Houston, TX<br />
pgruhn@landmengineering.com<br />
<br />
== Refferences ==<br />
<br />
[1] - http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf - Methods of Determining Safety Integrity Level (SIL)<br />
Requirements - Pros and Cons<br />
by W G Gulland (4-sight Consulting)<br />
<br />
[2] http://www.iceweb.com.au/sis/target_sis.htm - Techniques for Assigning A Target Safety Integrity Level Angela E. Summers, Ph.D. This paper was published in ISA Transactions 37 (1998) 95-104.</div>John Creightonhttp://wiki.newmars.com/index.php?title=SIL&diff=778SIL2009-08-14T21:28:54Z<p>John Creighton: /* Risk Graphs */</p>
<hr />
<div>== Introduction ==<br />
Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple terms, SIL is a measurement of performance required for a Safety Instrumented Function (SIF).<br />
<br />
<br />
{| class="wikitable"<br />
|-<br />
! SIL<br />
! PFD<br />
! RRF<br />
|-<br />
| 1<br />
| 0.1-0.01<br />
| 10-100<br />
|-<br />
| 2<br />
| 0.01-0.001<br />
| 100-1000<br />
|-<br />
| 3<br />
| 0.001-0.0001<br />
| 1000-10,000<br />
|-<br />
| 4<br />
| 0.0001-0.00001<br />
| 10,000-100,000<br />
|}<br />
<br />
== Selection of SIL ==<br />
<br />
There are many methods to select the safety Integrity Level, these include, Risk Matrix, Risk Graph, Layers of Protection Analysis (LOPA). With regards to LOC NASA has considerably lower levels of safety then in comparison to industry because, the crew are not considered civilians, and the gains of space exploration to be with the risk at least in terms of LOC. <br />
<br />
Bellow is what is a graph of what is considered acceptable and intolerable in terms of fatalities by (HSE Books 2001)[1]. Of course these graphs would very depending on the size of the population and the significance of the endeavor. <br />
<br />
[[Image:Fatality_graph.JPG]] <br />
<br />
== Sill Selection Matrices ==<br />
<br />
A SIL matrix tells us how reliable a safety function must be given the likely hood and the severity of an event. Bellow a 3D sill selection matrix is shown. If you can achieve the desired safety levels independently of the other layers then a 3D sill section matrix may not be necessary. However, there may be a maximum amount of reliability we can achieve from a given safety function and therefore we must consider how each layer of protection contributes to the reliability of the overall system. For instance, it was suggested in the Augastine commission that an abort system only reduces your LOC (Loss of crew) by about a factor of 10. If this does not give the required reliability in terms of loss of crew, then we must consider the reliability of the other layers of the system. <br />
<br />
[[Image:SILL_Matrix.JPG]]<br />
<br />
Generally a separate SIL selection matrix is done for each type of consequence, these can include, Loss of mission, Loss of Crew, Environmental impact and Damage to the reputation of the organization. Generally each type of consequence is considered separately and it is the type of consequence which requires the greatest level of safety which drives the design.<br />
<br />
== LOPA (Layers of Protection Analysis ==<br />
<br />
== Risk Graphs ==<br />
<br />
<br />
[[Image:Risk_Graph.JPG]]<br />
<br />
== Refferences ==<br />
<br />
[1] - http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf - Methods of Determining Safety Integrity Level (SIL)<br />
Requirements - Pros and Cons<br />
by W G Gulland (4-sight Consulting)<br />
<br />
[2] http://www.iceweb.com.au/sis/target_sis.htm - Techniques for Assigning A Target Safety Integrity Level Angela E. Summers, Ph.D. This paper was published in ISA Transactions 37 (1998) 95-104.</div>John Creightonhttp://wiki.newmars.com/index.php?title=File:Risk_Graph.JPG&diff=777File:Risk Graph.JPG2009-08-14T21:28:14Z<p>John Creighton: http://www.isa.org/Content/Microsites838/Safety_Division/Home818/ISA_2004_Safety_Papers/Different_SIL_Selection_Techniques_Can_Yield_Different_Answers.pdf</p>
<hr />
<div>http://www.isa.org/Content/Microsites838/Safety_Division/Home818/ISA_2004_Safety_Papers/Different_SIL_Selection_Techniques_Can_Yield_Different_Answers.pdf</div>John Creightonhttp://wiki.newmars.com/index.php?title=SIL&diff=776SIL2009-08-14T21:24:48Z<p>John Creighton: /* Sill Selection Matrices */</p>
<hr />
<div>== Introduction ==<br />
Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple terms, SIL is a measurement of performance required for a Safety Instrumented Function (SIF).<br />
<br />
<br />
{| class="wikitable"<br />
|-<br />
! SIL<br />
! PFD<br />
! RRF<br />
|-<br />
| 1<br />
| 0.1-0.01<br />
| 10-100<br />
|-<br />
| 2<br />
| 0.01-0.001<br />
| 100-1000<br />
|-<br />
| 3<br />
| 0.001-0.0001<br />
| 1000-10,000<br />
|-<br />
| 4<br />
| 0.0001-0.00001<br />
| 10,000-100,000<br />
|}<br />
<br />
== Selection of SIL ==<br />
<br />
There are many methods to select the safety Integrity Level, these include, Risk Matrix, Risk Graph, Layers of Protection Analysis (LOPA). With regards to LOC NASA has considerably lower levels of safety then in comparison to industry because, the crew are not considered civilians, and the gains of space exploration to be with the risk at least in terms of LOC. <br />
<br />
Bellow is what is a graph of what is considered acceptable and intolerable in terms of fatalities by (HSE Books 2001)[1]. Of course these graphs would very depending on the size of the population and the significance of the endeavor. <br />
<br />
[[Image:Fatality_graph.JPG]] <br />
<br />
== Sill Selection Matrices ==<br />
<br />
A SIL matrix tells us how reliable a safety function must be given the likely hood and the severity of an event. Bellow a 3D sill selection matrix is shown. If you can achieve the desired safety levels independently of the other layers then a 3D sill section matrix may not be necessary. However, there may be a maximum amount of reliability we can achieve from a given safety function and therefore we must consider how each layer of protection contributes to the reliability of the overall system. For instance, it was suggested in the Augastine commission that an abort system only reduces your LOC (Loss of crew) by about a factor of 10. If this does not give the required reliability in terms of loss of crew, then we must consider the reliability of the other layers of the system. <br />
<br />
[[Image:SILL_Matrix.JPG]]<br />
<br />
Generally a separate SIL selection matrix is done for each type of consequence, these can include, Loss of mission, Loss of Crew, Environmental impact and Damage to the reputation of the organization. Generally each type of consequence is considered separately and it is the type of consequence which requires the greatest level of safety which drives the design.<br />
<br />
== LOPA (Layers of Protection Analysis ==<br />
<br />
== Risk Graphs ==<br />
<br />
<br />
== Refferences ==<br />
<br />
[1] - http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf - Methods of Determining Safety Integrity Level (SIL)<br />
Requirements - Pros and Cons<br />
by W G Gulland (4-sight Consulting)<br />
<br />
[2] http://www.iceweb.com.au/sis/target_sis.htm - Techniques for Assigning A Target Safety Integrity Level Angela E. Summers, Ph.D. This paper was published in ISA Transactions 37 (1998) 95-104.</div>John Creightonhttp://wiki.newmars.com/index.php?title=File:SILL_Matrix.JPG&diff=775File:SILL Matrix.JPG2009-08-14T21:14:35Z<p>John Creighton: The above shows a 3D SILL Section matrix. The required reliability of a safety function is dependent both on the severity of an event and the likelihood of an event.
http://www.isa.org/Content/Microsites838/Safety_Division/Home818/ISA_2004_Safety_Pape</p>
<hr />
<div>The above shows a 3D SILL Section matrix. The required reliability of a safety function is dependent both on the severity of an event and the likelihood of an event. <br />
<br />
<br />
http://www.isa.org/Content/Microsites838/Safety_Division/Home818/ISA_2004_Safety_Papers/Different_SIL_Selection_Techniques_Can_Yield_Different_Answers.pdf</div>John Creightonhttp://wiki.newmars.com/index.php?title=SIL&diff=774SIL2009-08-14T21:07:50Z<p>John Creighton: New page: == Introduction == Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple te...</p>
<hr />
<div>== Introduction ==<br />
Safety Integrity Level (SIL) is defined as a relative level of risk-reduction provided by a safety function, or to specify a target level of risk reduction. In simple terms, SIL is a measurement of performance required for a Safety Instrumented Function (SIF).<br />
<br />
<br />
{| class="wikitable"<br />
|-<br />
! SIL<br />
! PFD<br />
! RRF<br />
|-<br />
| 1<br />
| 0.1-0.01<br />
| 10-100<br />
|-<br />
| 2<br />
| 0.01-0.001<br />
| 100-1000<br />
|-<br />
| 3<br />
| 0.001-0.0001<br />
| 1000-10,000<br />
|-<br />
| 4<br />
| 0.0001-0.00001<br />
| 10,000-100,000<br />
|}<br />
<br />
== Selection of SIL ==<br />
<br />
There are many methods to select the safety Integrity Level, these include, Risk Matrix, Risk Graph, Layers of Protection Analysis (LOPA). With regards to LOC NASA has considerably lower levels of safety then in comparison to industry because, the crew are not considered civilians, and the gains of space exploration to be with the risk at least in terms of LOC. <br />
<br />
Bellow is what is a graph of what is considered acceptable and intolerable in terms of fatalities by (HSE Books 2001)[1]. Of course these graphs would very depending on the size of the population and the significance of the endeavor. <br />
<br />
[[Image:Fatality_graph.JPG]] <br />
<br />
== Sill Selection Matrices ==<br />
<br />
== LOPA (Layers of Protection Analysis ==<br />
<br />
== Risk Graphs ==<br />
<br />
<br />
== Refferences ==<br />
<br />
[1] - http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf - Methods of Determining Safety Integrity Level (SIL)<br />
Requirements - Pros and Cons<br />
by W G Gulland (4-sight Consulting)<br />
<br />
[2] http://www.iceweb.com.au/sis/target_sis.htm - Techniques for Assigning A Target Safety Integrity Level Angela E. Summers, Ph.D. This paper was published in ISA Transactions 37 (1998) 95-104.</div>John Creightonhttp://wiki.newmars.com/index.php?title=File:Fatality_graph.JPG&diff=773File:Fatality graph.JPG2009-08-14T21:00:25Z<p>John Creighton: http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf
The HSE have
suggested that a risk of one 50 fatality event in 5,000 years is intolerable (HSE Books 2001).</p>
<hr />
<div>http://4-sightconsulting.co.uk/Current_Papers/Determining_SILs/Methods_of_Determining_Safety_Integrity_Level.pdf<br />
<br />
The HSE have<br />
suggested that a risk of one 50 fatality event in 5,000 years is intolerable (HSE Books 2001).</div>John Creightonhttp://wiki.newmars.com/index.php?title=Optimal_ground_launch_flight_path&diff=771Optimal ground launch flight path2009-08-14T07:55:30Z<p>John Creighton: /* A Visual Basic Function to Calculate the Optimal Flight Path */</p>
<hr />
<div>== Introduction ==<br />
The current derivation of the rocket equation on wikipedia (Aug 13 2009) seems to be overly cumbersome:<br />
http://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation<br />
<br />
The rocket equation basically comes down to the basic law of Newtonian dynamics. That is the sum of the forces equals the mass multiplied by the acceleration.<br />
<br />
<math>\sum F=ma</math><br />
<br />
The product of the rocket exhaust velocity multiplied the mass flow is the rate momentum is leaving the rocket. Because momentum must be conserved the rocket gains an equal and opposite amount of momentum as the exhaust. A more convenient form of newtons law is:<br />
<br />
<math>\sum F=ma=m\frac{dv}{dt}=\frac{dmv}{dt}=\frac{dp}{dt}</math><br />
<br />
That is the force is the rate momentum changes with time. Consequently, the force exerted on the rocket by the exhaust is equal to the velocity multiplied by the mass flow. Taking into gravity the fores acting in the direction of flight and dividing both sides by the mass we obtain:<br />
<br />
<math>a=V_e\frac{dm}{dt}+g sin(\theta)</math><br />
<br />
== Optimal Flight Path ==<br />
<br />
The optimal flight path occurs when the vertical component of the thrust cancels out gravity and the remaining part of the thrust is due to horizontal acceleration. This can be written as:<br />
<br />
<math>g-\frac{v^2}{R}cos(\theta)-T sin (\theta)=0</math><br />
<br />
where:<br />
g is the gravitational acceleration<br />
R is the distance from the center of the earth<br />
T is the thrust per mass.<br />
<br />
Using trig identities:<br />
<br />
<math>g+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}-T sin (\theta)=0</math><br />
<br />
Rearranging:<br />
<br />
<math>+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}=T sin (\theta)-g</math><br />
<br />
Squaring both sides:<br />
<br />
<math>\frac{v^4}{R^2}(1-sin^2(\theta))=T^2 sin^2 (\theta)-2gTsin(\theta)+g^2</math><br />
<br />
Rearranging:<br />
<br />
<math>(T^2+\frac{v^4}{R^2}) sin^2 (\theta)-2gTsin(\theta)+g^2-\frac{v^4}{R^2}=0</math><br />
<br />
This equation can be solved using the Quadratic Equation.<br />
<br />
== A Visual Basic Function to Calculate the Optimal Thrust Angle ==<br />
<br />
Bellow is a visual basic function used in [[RocketDesignSpreadsheet]] to find the optimal thrust angle.<br />
<br />
Function thrustAngle(V, T, Optional h, Optional g)<br />
Re = 6378.137 * 1000 ' Radious of The Earth<br />
If IsMissing(h) Then<br />
h = 0<br />
End If<br />
If IsMissing(g) Then<br />
Gc = 0.00000000006674 ' N(m/kg)^2Gravational Constant<br />
Mearth = 5.9736E+24 ' Mass of the earth<br />
g = Gc * Mearth / (h + Re) ^ 2<br />
End If<br />
R = h + Re<br />
a = (T ^ 2 + V ^ 4 / R ^ 2)<br />
b = -2 * g * T<br />
c = g ^ 2 - V ^ 4 / R ^ 2<br />
des = b ^ 2 - 4 * a * c<br />
If des < 0 Then<br />
thrustAngle = 0<br />
Else<br />
r1 = (-b + Sqr(des)) / (2 * a)<br />
r2 = (-b - Sqr(des)) / (2 * a)<br />
theata1 = WorksheetFunction.Asin(r1)<br />
theata2 = WorksheetFunction.Asin(r2)<br />
resid1 = g - V ^ 2 / R * Cos(theata1) - T * Sin(theata1)<br />
resid2 = g - V ^ 2 / R * Cos(theata2) - T * Sin(theata2)<br />
If Abs(resid1) < Abs(resid2) Then<br />
thrustAngle = theata1<br />
Else<br />
thrustAngle = theata2<br />
End If<br />
End If<br />
If thrustAngle < 0 Then<br />
thrustAngle = 0<br />
End If<br />
End Function<br />
<br />
== Total Gravity Losses ==<br />
<br />
Returning to the equation mentioned in the introduction:<br />
<br />
<math>a=V_e\frac{dm}{dt}-g sin(\theta)</math><br />
<br />
If we integrate this equation over time, we get the change the change in velocity we achieve. The term g sin(\theta) represents the gravity losses, and if we integrate this term over the flight path, it will gives us the additional delta V we need to overcome because of gravity losses. This term does not need to be calculated for the whole flight path because these losses will taper off as the thrust angle approaches horizontal.</div>John Creightonhttp://wiki.newmars.com/index.php?title=Optimal_ground_launch_flight_path&diff=770Optimal ground launch flight path2009-08-14T06:07:13Z<p>John Creighton: /* A Visual Basic Function to Calculate the Optimal Flight Path */</p>
<hr />
<div>== Introduction ==<br />
The current derivation of the rocket equation on wikipedia (Aug 13 2009) seems to be overly cumbersome:<br />
http://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation<br />
<br />
The rocket equation basically comes down to the basic law of Newtonian dynamics. That is the sum of the forces equals the mass multiplied by the acceleration.<br />
<br />
<math>\sum F=ma</math><br />
<br />
The product of the rocket exhaust velocity multiplied the mass flow is the rate momentum is leaving the rocket. Because momentum must be conserved the rocket gains an equal and opposite amount of momentum as the exhaust. A more convenient form of newtons law is:<br />
<br />
<math>\sum F=ma=m\frac{dv}{dt}=\frac{dmv}{dt}=\frac{dp}{dt}</math><br />
<br />
That is the force is the rate momentum changes with time. Consequently, the force exerted on the rocket by the exhaust is equal to the velocity multiplied by the mass flow. Taking into gravity the fores acting in the direction of flight and dividing both sides by the mass we obtain:<br />
<br />
<math>a=V_e\frac{dm}{dt}+g sin(\theta)</math><br />
<br />
== Optimal Flight Path ==<br />
<br />
The optimal flight path occurs when the vertical component of the thrust cancels out gravity and the remaining part of the thrust is due to horizontal acceleration. This can be written as:<br />
<br />
<math>g-\frac{v^2}{R}cos(\theta)-T sin (\theta)=0</math><br />
<br />
where:<br />
g is the gravitational acceleration<br />
R is the distance from the center of the earth<br />
T is the thrust per mass.<br />
<br />
Using trig identities:<br />
<br />
<math>g+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}-T sin (\theta)=0</math><br />
<br />
Rearranging:<br />
<br />
<math>+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}=T sin (\theta)-g</math><br />
<br />
Squaring both sides:<br />
<br />
<math>\frac{v^4}{R^2}(1-sin^2(\theta))=T^2 sin^2 (\theta)-2gTsin(\theta)+g^2</math><br />
<br />
Rearranging:<br />
<br />
<math>(T^2+\frac{v^4}{R^2}) sin^2 (\theta)-2gTsin(\theta)+g^2-\frac{v^4}{R^2}=0</math><br />
<br />
This equation can be solved using the Quadratic Equation.<br />
<br />
== A Visual Basic Function to Calculate the Optimal Flight Path ==<br />
<br />
Bellow is a visual basic function used in [[RocketDesignSpreadsheet]] to find the optimal thrust angle.<br />
<br />
Function thrustAngle(V, T, Optional h, Optional g)<br />
Re = 6378.137 * 1000 ' Radious of The Earth<br />
If IsMissing(h) Then<br />
h = 0<br />
End If<br />
If IsMissing(g) Then<br />
Gc = 0.00000000006674 ' N(m/kg)^2Gravational Constant<br />
Mearth = 5.9736E+24 ' Mass of the earth<br />
g = Gc * Mearth / (h + Re) ^ 2<br />
End If<br />
R = h + Re<br />
a = (T ^ 2 + V ^ 4 / R ^ 2)<br />
b = -2 * g * T<br />
c = g ^ 2 - V ^ 4 / R ^ 2<br />
des = b ^ 2 - 4 * a * c<br />
If des < 0 Then<br />
thrustAngle = 0<br />
Else<br />
r1 = (-b + Sqr(des)) / (2 * a)<br />
r2 = (-b - Sqr(des)) / (2 * a)<br />
theata1 = WorksheetFunction.Asin(r1)<br />
theata2 = WorksheetFunction.Asin(r2)<br />
resid1 = g - V ^ 2 / R * Cos(theata1) - T * Sin(theata1)<br />
resid2 = g - V ^ 2 / R * Cos(theata2) - T * Sin(theata2)<br />
If Abs(resid1) < Abs(resid2) Then<br />
thrustAngle = theata1<br />
Else<br />
thrustAngle = theata2<br />
End If<br />
End If<br />
If thrustAngle < 0 Then<br />
thrustAngle = 0<br />
End If<br />
End Function<br />
<br />
== Total Gravity Losses ==<br />
<br />
Returning to the equation mentioned in the introduction:<br />
<br />
<math>a=V_e\frac{dm}{dt}-g sin(\theta)</math><br />
<br />
If we integrate this equation over time, we get the change the change in velocity we achieve. The term g sin(\theta) represents the gravity losses, and if we integrate this term over the flight path, it will gives us the additional delta V we need to overcome because of gravity losses. This term does not need to be calculated for the whole flight path because these losses will taper off as the thrust angle approaches horizontal.</div>John Creightonhttp://wiki.newmars.com/index.php?title=Optimal_ground_launch_flight_path&diff=768Optimal ground launch flight path2009-08-14T00:23:14Z<p>John Creighton: /* A Visual Basic Function to Caculate the Optimal Flight Path */</p>
<hr />
<div>== Introduction ==<br />
The current derivation of the rocket equation on wikipedia (Aug 13 2009) seems to be overly cumbersome:<br />
http://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation<br />
<br />
The rocket equation basically comes down to the basic law of Newtonian dynamics. That is the sum of the forces equals the mass multiplied by the acceleration.<br />
<br />
<math>\sum F=ma</math><br />
<br />
The product of the rocket exhaust velocity multiplied the mass flow is the rate momentum is leaving the rocket. Because momentum must be conserved the rocket gains an equal and opposite amount of momentum as the exhaust. A more convenient form of newtons law is:<br />
<br />
<math>\sum F=ma=m\frac{dv}{dt}=\frac{dmv}{dt}=\frac{dp}{dt}</math><br />
<br />
That is the force is the rate momentum changes with time. Consequently, the force exerted on the rocket by the exhaust is equal to the velocity multiplied by the mass flow. Taking into gravity the fores acting in the direction of flight and dividing both sides by the mass we obtain:<br />
<br />
<math>a=V_e\frac{dm}{dt}+g sin(\theta)</math><br />
<br />
== Optimal Flight Path ==<br />
<br />
The optimal flight path occurs when the vertical component of the thrust cancels out gravity and the remaining part of the thrust is due to horizontal acceleration. This can be written as:<br />
<br />
<math>g-\frac{v^2}{R}cos(\theta)-T sin (\theta)=0</math><br />
<br />
where:<br />
g is the gravitational acceleration<br />
R is the distance from the center of the earth<br />
T is the thrust per mass.<br />
<br />
Using trig identities:<br />
<br />
<math>g+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}-T sin (\theta)=0</math><br />
<br />
Rearranging:<br />
<br />
<math>+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}=T sin (\theta)-g</math><br />
<br />
Squaring both sides:<br />
<br />
<math>\frac{v^4}{R^2}(1-sin^2(\theta))=T^2 sin^2 (\theta)-2gTsin(\theta)+g^2</math><br />
<br />
Rearranging:<br />
<br />
<math>(T^2+\frac{v^4}{R^2}) sin^2 (\theta)-2gTsin(\theta)+g^2-\frac{v^4}{R^2}=0</math><br />
<br />
This equation can be solved using the Quadratic Equation.<br />
<br />
== A Visual Basic Function to Calculate the Optimal Flight Path ==<br />
<br />
Bellow is a visual basic function used in [[RocketDesignSpreadsheet]] the optimal thrust angle.<br />
<br />
<br />
<nowiki>Function thrustAngle(V, T, Optional h, Optional g)<br />
Re = 6378.137 * 1000 ' Radious of The Earth<br />
If IsMissing(h) Then<br />
h = 0<br />
End If<br />
If IsMissing(g) Then<br />
Gc = 0.00000000006674 ' N(m/kg)^2Gravational Constant<br />
Mearth = 5.9736E+24 ' Mass of the earth<br />
g = Gc * Mearth / (h + Re) ^ 2<br />
End If<br />
R = h + Re<br />
a = (T ^ 2 + V ^ 4 / R ^ 2)<br />
b = -2 * g * T<br />
c = g ^ 2 - V ^ 4 / R ^ 2<br />
des = b ^ 2 - 4 * a * c<br />
If des < 0 Then<br />
thrustAngle = 0<br />
Else<br />
r1 = (-b + Sqr(des)) / (2 * a)<br />
r2 = (-b - Sqr(des)) / (2 * a)<br />
theata1 = WorksheetFunction.Asin(r1)<br />
theata2 = WorksheetFunction.Asin(r2)<br />
resid1 = g - V ^ 2 / R * Cos(theata1) - T * Sin(theata1)<br />
resid2 = g - V ^ 2 / R * Cos(theata2) - T * Sin(theata2)<br />
If Abs(resid1) < Abs(resid2) Then<br />
thrustAngle = theata1<br />
Else<br />
thrustAngle = theata2<br />
End If<br />
End If<br />
If thrustAngle < 0 Then<br />
thrustAngle = 0<br />
End If<br />
<br />
<br />
End Function<br />
</nowiki><br />
<br />
== Total Gravity Losses ==<br />
<br />
Returning to the equation mentioned in the introduction:<br />
<br />
<math>a=V_e\frac{dm}{dt}-g sin(\theta)</math><br />
<br />
If we integrate this equation over time, we get the change the change in velocity we achieve. The term g sin(\theta) represents the gravity losses, and if we integrate this term over the flight path, it will gives us the additional delta V we need to overcome because of gravity losses. This term does not need to be calculated for the whole flight path because these losses will taper off as the thrust angle approaches horizontal.</div>John Creightonhttp://wiki.newmars.com/index.php?title=Optimal_ground_launch_flight_path&diff=767Optimal ground launch flight path2009-08-13T10:01:49Z<p>John Creighton: /* Total Gravity Losses */</p>
<hr />
<div>== Introduction ==<br />
The current derivation of the rocket equation on wikipedia (Aug 13 2009) seems to be overly cumbersome:<br />
http://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation<br />
<br />
The rocket equation basically comes down to the basic law of Newtonian dynamics. That is the sum of the forces equals the mass multiplied by the acceleration.<br />
<br />
<math>\sum F=ma</math><br />
<br />
The product of the rocket exhaust velocity multiplied the mass flow is the rate momentum is leaving the rocket. Because momentum must be conserved the rocket gains an equal and opposite amount of momentum as the exhaust. A more convenient form of newtons law is:<br />
<br />
<math>\sum F=ma=m\frac{dv}{dt}=\frac{dmv}{dt}=\frac{dp}{dt}</math><br />
<br />
That is the force is the rate momentum changes with time. Consequently, the force exerted on the rocket by the exhaust is equal to the velocity multiplied by the mass flow. Taking into gravity the fores acting in the direction of flight and dividing both sides by the mass we obtain:<br />
<br />
<math>a=V_e\frac{dm}{dt}+g sin(\theta)</math><br />
<br />
== Optimal Flight Path ==<br />
<br />
The optimal flight path occurs when the vertical component of the thrust cancels out gravity and the remaining part of the thrust is due to horizontal acceleration. This can be written as:<br />
<br />
<math>g-\frac{v^2}{R}cos(\theta)-T sin (\theta)=0</math><br />
<br />
where:<br />
g is the gravitational acceleration<br />
R is the distance from the center of the earth<br />
T is the thrust per mass.<br />
<br />
Using trig identities:<br />
<br />
<math>g+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}-T sin (\theta)=0</math><br />
<br />
Rearranging:<br />
<br />
<math>+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}=T sin (\theta)-g</math><br />
<br />
Squaring both sides:<br />
<br />
<math>\frac{v^4}{R^2}(1-sin^2(\theta))=T^2 sin^2 (\theta)-2gTsin(\theta)+g^2</math><br />
<br />
Rearranging:<br />
<br />
<math>(T^2+\frac{v^4}{R^2}) sin^2 (\theta)-2gTsin(\theta)+g^2-\frac{v^4}{R^2}=0</math><br />
<br />
This equation can be solved using the Quadratic Equation.<br />
<br />
== A Visual Basic Function to Caculate the Optimal Flight Path ==<br />
<br />
Bellow is a visual basic function used in [[RocketDesignSpreadsheet]] the optimal thrust angle.<br />
<br />
<br />
<nowiki>Function thrustAngle(V, T, Optional h, Optional g)<br />
Re = 6378.137 * 1000 ' Radious of The Earth<br />
If IsMissing(h) Then<br />
h = 0<br />
End If<br />
If IsMissing(g) Then<br />
Gc = 0.00000000006674 ' N(m/kg)^2Gravational Constant<br />
Mearth = 5.9736E+24 ' Mass of the earth<br />
g = Gc * Mearth / (h + Re) ^ 2<br />
End If<br />
R = h + Re<br />
a = (T ^ 2 + V ^ 4 / R ^ 2)<br />
b = -2 * g * T<br />
c = g ^ 2 - b ^ 4 / R ^ 2<br />
des = b ^ 2 - 4 * a * c<br />
If des < 0 Then<br />
thrustAngle = 0<br />
Else<br />
r1 = (-b + Sqr(des)) / (2 * a)<br />
r2 = (-b - Sqr(des)) / (2 * a)<br />
theata1 = WorksheetFunction.Asin(r1)<br />
theata2 = WorksheetFunction.Asin(r2)<br />
resid1 = g - V ^ 2 / R * Cos(theata1) - T * Sin(theata1)<br />
resid2 = g - V ^ 2 / R * Cos(theata2) - T * Sin(theata2)<br />
If Abs(resid1) < Abs(resid2) Then<br />
thrustAngle = theata1<br />
Else<br />
thrustAngle = theata2<br />
End If<br />
End If<br />
End Function<br />
</nowiki><br />
<br />
== Total Gravity Losses ==<br />
<br />
Returning to the equation mentioned in the introduction:<br />
<br />
<math>a=V_e\frac{dm}{dt}-g sin(\theta)</math><br />
<br />
If we integrate this equation over time, we get the change the change in velocity we achieve. The term g sin(\theta) represents the gravity losses, and if we integrate this term over the flight path, it will gives us the additional delta V we need to overcome because of gravity losses. This term does not need to be calculated for the whole flight path because these losses will taper off as the thrust angle approaches horizontal.</div>John Creightonhttp://wiki.newmars.com/index.php?title=Optimal_ground_launch_flight_path&diff=766Optimal ground launch flight path2009-08-13T09:59:06Z<p>John Creighton: /* Introduction */</p>
<hr />
<div>== Introduction ==<br />
The current derivation of the rocket equation on wikipedia (Aug 13 2009) seems to be overly cumbersome:<br />
http://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation<br />
<br />
The rocket equation basically comes down to the basic law of Newtonian dynamics. That is the sum of the forces equals the mass multiplied by the acceleration.<br />
<br />
<math>\sum F=ma</math><br />
<br />
The product of the rocket exhaust velocity multiplied the mass flow is the rate momentum is leaving the rocket. Because momentum must be conserved the rocket gains an equal and opposite amount of momentum as the exhaust. A more convenient form of newtons law is:<br />
<br />
<math>\sum F=ma=m\frac{dv}{dt}=\frac{dmv}{dt}=\frac{dp}{dt}</math><br />
<br />
That is the force is the rate momentum changes with time. Consequently, the force exerted on the rocket by the exhaust is equal to the velocity multiplied by the mass flow. Taking into gravity the fores acting in the direction of flight and dividing both sides by the mass we obtain:<br />
<br />
<math>a=V_e\frac{dm}{dt}+g sin(\theta)</math><br />
<br />
== Optimal Flight Path ==<br />
<br />
The optimal flight path occurs when the vertical component of the thrust cancels out gravity and the remaining part of the thrust is due to horizontal acceleration. This can be written as:<br />
<br />
<math>g-\frac{v^2}{R}cos(\theta)-T sin (\theta)=0</math><br />
<br />
where:<br />
g is the gravitational acceleration<br />
R is the distance from the center of the earth<br />
T is the thrust per mass.<br />
<br />
Using trig identities:<br />
<br />
<math>g+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}-T sin (\theta)=0</math><br />
<br />
Rearranging:<br />
<br />
<math>+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}=T sin (\theta)-g</math><br />
<br />
Squaring both sides:<br />
<br />
<math>\frac{v^4}{R^2}(1-sin^2(\theta))=T^2 sin^2 (\theta)-2gTsin(\theta)+g^2</math><br />
<br />
Rearranging:<br />
<br />
<math>(T^2+\frac{v^4}{R^2}) sin^2 (\theta)-2gTsin(\theta)+g^2-\frac{v^4}{R^2}=0</math><br />
<br />
This equation can be solved using the Quadratic Equation.<br />
<br />
== A Visual Basic Function to Caculate the Optimal Flight Path ==<br />
<br />
Bellow is a visual basic function used in [[RocketDesignSpreadsheet]] the optimal thrust angle.<br />
<br />
<br />
<nowiki>Function thrustAngle(V, T, Optional h, Optional g)<br />
Re = 6378.137 * 1000 ' Radious of The Earth<br />
If IsMissing(h) Then<br />
h = 0<br />
End If<br />
If IsMissing(g) Then<br />
Gc = 0.00000000006674 ' N(m/kg)^2Gravational Constant<br />
Mearth = 5.9736E+24 ' Mass of the earth<br />
g = Gc * Mearth / (h + Re) ^ 2<br />
End If<br />
R = h + Re<br />
a = (T ^ 2 + V ^ 4 / R ^ 2)<br />
b = -2 * g * T<br />
c = g ^ 2 - b ^ 4 / R ^ 2<br />
des = b ^ 2 - 4 * a * c<br />
If des < 0 Then<br />
thrustAngle = 0<br />
Else<br />
r1 = (-b + Sqr(des)) / (2 * a)<br />
r2 = (-b - Sqr(des)) / (2 * a)<br />
theata1 = WorksheetFunction.Asin(r1)<br />
theata2 = WorksheetFunction.Asin(r2)<br />
resid1 = g - V ^ 2 / R * Cos(theata1) - T * Sin(theata1)<br />
resid2 = g - V ^ 2 / R * Cos(theata2) - T * Sin(theata2)<br />
If Abs(resid1) < Abs(resid2) Then<br />
thrustAngle = theata1<br />
Else<br />
thrustAngle = theata2<br />
End If<br />
End If<br />
End Function<br />
</nowiki><br />
<br />
== Total Gravity Losses ==</div>John Creightonhttp://wiki.newmars.com/index.php?title=Optimal_ground_launch_flight_path&diff=765Optimal ground launch flight path2009-08-13T09:57:07Z<p>John Creighton: </p>
<hr />
<div>== Introduction ==<br />
The current derivation of the rocket equation on wikipedia (Aug 13 2009) seems to be overly cumbersome:<br />
http://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation<br />
<br />
The rocket equation basically comes down to the basic law of Newtonian dynamics. That is the sum of the forces equals the mass multiplied by the acceleration.<br />
<br />
<math>\sum F=ma</math><br />
<br />
The product of the rocket exhaust velocity multiplied the mass flow is the rate momentum is leaving the rocket. Because momentum must be conserved the rocket gains an equal and opposite amount of momentum as the exhaust. A more convenient form of newtons law is:<br />
<br />
<math>\sum F=ma=m\frac{dv}{dt}=\frac{dmv}{dt}=\frac{dp}{dt}</math><br />
<br />
That is the force is the rate momentum changes with time. Consequently, the force exerted on the rocket by the exhaust is equal to the velocity multiplied by the mass flow. Taking into gravity the fores acting in the direction of flight are given by:<br />
<br />
<math>ma=V_e\frac{dm}{dt}+g sin(\theta)</math><br />
<br />
== Optimal Flight Path ==<br />
<br />
The optimal flight path occurs when the vertical component of the thrust cancels out gravity and the remaining part of the thrust is due to horizontal acceleration. This can be written as:<br />
<br />
<math>g-\frac{v^2}{R}cos(\theta)-T sin (\theta)=0</math><br />
<br />
where:<br />
g is the gravitational acceleration<br />
R is the distance from the center of the earth<br />
T is the thrust per mass.<br />
<br />
Using trig identities:<br />
<br />
<math>g+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}-T sin (\theta)=0</math><br />
<br />
Rearranging:<br />
<br />
<math>+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}=T sin (\theta)-g</math><br />
<br />
Squaring both sides:<br />
<br />
<math>\frac{v^4}{R^2}(1-sin^2(\theta))=T^2 sin^2 (\theta)-2gTsin(\theta)+g^2</math><br />
<br />
Rearranging:<br />
<br />
<math>(T^2+\frac{v^4}{R^2}) sin^2 (\theta)-2gTsin(\theta)+g^2-\frac{v^4}{R^2}=0</math><br />
<br />
This equation can be solved using the Quadratic Equation.<br />
<br />
== A Visual Basic Function to Caculate the Optimal Flight Path ==<br />
<br />
Bellow is a visual basic function used in [[RocketDesignSpreadsheet]] the optimal thrust angle.<br />
<br />
<br />
<nowiki>Function thrustAngle(V, T, Optional h, Optional g)<br />
Re = 6378.137 * 1000 ' Radious of The Earth<br />
If IsMissing(h) Then<br />
h = 0<br />
End If<br />
If IsMissing(g) Then<br />
Gc = 0.00000000006674 ' N(m/kg)^2Gravational Constant<br />
Mearth = 5.9736E+24 ' Mass of the earth<br />
g = Gc * Mearth / (h + Re) ^ 2<br />
End If<br />
R = h + Re<br />
a = (T ^ 2 + V ^ 4 / R ^ 2)<br />
b = -2 * g * T<br />
c = g ^ 2 - b ^ 4 / R ^ 2<br />
des = b ^ 2 - 4 * a * c<br />
If des < 0 Then<br />
thrustAngle = 0<br />
Else<br />
r1 = (-b + Sqr(des)) / (2 * a)<br />
r2 = (-b - Sqr(des)) / (2 * a)<br />
theata1 = WorksheetFunction.Asin(r1)<br />
theata2 = WorksheetFunction.Asin(r2)<br />
resid1 = g - V ^ 2 / R * Cos(theata1) - T * Sin(theata1)<br />
resid2 = g - V ^ 2 / R * Cos(theata2) - T * Sin(theata2)<br />
If Abs(resid1) < Abs(resid2) Then<br />
thrustAngle = theata1<br />
Else<br />
thrustAngle = theata2<br />
End If<br />
End If<br />
End Function<br />
</nowiki><br />
<br />
== Total Gravity Losses ==</div>John Creightonhttp://wiki.newmars.com/index.php?title=Optimal_ground_launch_flight_path&diff=764Optimal ground launch flight path2009-08-13T09:51:20Z<p>John Creighton: New page: The current derivation of the rocket equation on wikipedia (Aug 13 2009) seems to be overly cumbersome: http://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation Relay, it comes down to th...</p>
<hr />
<div>The current derivation of the rocket equation on wikipedia (Aug 13 2009) seems to be overly cumbersome:<br />
http://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation<br />
<br />
Relay, it comes down to the basic law of Newtonian dynamics. That is the sum of the forces equals the mass multiplied by the acceleration.<br />
<br />
<math>\sum F=ma</math><br />
<br />
The product of the rocket exaust velocity multiplied the mass flow is the rate momentum is leaving the rocket. Because momentum must be conserved the rocket gains an equal and opposite amount of momentum as the exhaust. A more convenient form of newtons law is:<br />
<br />
<math>\sum F=ma=m\frac{dv}{dt}=\frac{dmv}{dt}=\frac{dp}{dt}</math><br />
<br />
That is the force is the rate momentum changes with time. Consequently, the force exerted on the rocket by the exhaust is equal to the velocity multiplied by the mass flow. Taking into gravity the fores allowing the direction of flight are given by:<br />
<br />
<math>ma=V_e\frac{dm}{dt}+g sin(\theta)</math><br />
<br />
The optimal flight path occurs when the vertical component of the thrust cancels out gravity and the remaining part of the thrust is due to horizontal acceleration. This can be written as:<br />
<br />
<math>g-\frac{v^2}{R}cos(\theta)-T sin (\theta)=0</math><br />
<br />
where:<br />
g is the gravitional accleration<br />
R is the distance from the center of the earth<br />
T is the thrust per mass.<br />
<br />
Using trig identities:<br />
<br />
<math>g+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}-T sin (\theta)=0</math><br />
<br />
Rearranging:<br />
<br />
<math>+/-\frac{v^2}{R}\sqrt{1-sin^2(\theta)}=T sin (\theta)-g</math><br />
<br />
Squaring both sides:<br />
<br />
<math>\frac{v^4}{R^2}(1-sin^2(\theta))=T^2 sin^2 (\theta)-2gTsin(\theta)+g^2</math><br />
<br />
Rearranging:<br />
<br />
<math>(T^2+\frac{v^4}{R^2}) sin^2 (\theta)-2gTsin(\theta)+g^2-\frac{v^4}{R^2}=0</math><br />
<br />
This equation can be solved using the Quadratic Equation.<br />
<br />
Bellow is a visual basic function used in [[RocketDesignSpreadsheet]] the optimal thrust angle.<br />
<br />
<nowiki>Function thrustAngle(V, T, Optional h, Optional g)<br />
Re = 6378.137 * 1000 ' Radious of The Earth<br />
If IsMissing(h) Then<br />
h = 0<br />
End If<br />
If IsMissing(g) Then<br />
Gc = 0.00000000006674 ' N(m/kg)^2Gravational Constant<br />
Mearth = 5.9736E+24 ' Mass of the earth<br />
g = Gc * Mearth / (h + Re) ^ 2<br />
End If<br />
R = h + Re<br />
a = (T ^ 2 + V ^ 4 / R ^ 2)<br />
b = -2 * g * T<br />
c = g ^ 2 - b ^ 4 / R ^ 2<br />
des = b ^ 2 - 4 * a * c<br />
If des < 0 Then<br />
thrustAngle = 0<br />
Else<br />
r1 = (-b + Sqr(des)) / (2 * a)<br />
r2 = (-b - Sqr(des)) / (2 * a)<br />
theata1 = WorksheetFunction.Asin(r1)<br />
theata2 = WorksheetFunction.Asin(r2)<br />
resid1 = g - V ^ 2 / R * Cos(theata1) - T * Sin(theata1)<br />
resid2 = g - V ^ 2 / R * Cos(theata2) - T * Sin(theata2)<br />
If Abs(resid1) < Abs(resid2) Then<br />
thrustAngle = theata1<br />
Else<br />
thrustAngle = theata2<br />
End If<br />
End If<br />
End Function<br />
</nowiki></div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=763RocketDesignSpreadsheet2009-08-13T09:29:54Z<p>John Creighton: /* Sec 4 Rocket Performance */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the geometry of the oxidizer tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[Pressure_Vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
== Sec-3 Fuel Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the fuel tank and the geometry of the fuel tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each fuel tank and the total structural mass of all the fuel tanks combined. <br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]]<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
The final section of the spread sheet gives some overall performance of the rocket. In this part you specify the delta V you want to travel, the thrust per weight of the engine and the overall thrust per weight. The spread sheet then calculates the mass ratio and the approximate payload. The exit velocity of the fuel is chosen base on the above choice of fuel. Exit velocity is the product of gravity multiplied by the ISP.[4]<br />
<br />
This section also attempts to calculate gravity losses (Also known as Gravity Drag) [5]. The [[optimal ground launch flight path]] is such that the vertical component of the thrust cancels out gravity and the rest of the thrust is direct to horizontal acceleration.<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.<br />
<br />
[4] http://en.wikipedia.org/wiki/Specific_impulse#Specific_impulse_as_a_speed_.28effective_exhaust_velocity.29 - Relationship between specific impulse and ISP.<br />
<br />
[5] http://en.wikipedia.org/wiki/Gravity_drag - Wikipedia article on gravity drag.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=762RocketDesignSpreadsheet2009-08-13T09:15:41Z<p>John Creighton: /* Refferences */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the geometry of the oxidizer tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[Pressure_Vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
== Sec-3 Fuel Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the fuel tank and the geometry of the fuel tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each fuel tank and the total structural mass of all the fuel tanks combined. <br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]]<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
The final section of the spread sheet gives some overall performance of the rocket. In this part you specify the delta V you want to travel, the thrust per weight of the engine and the overall thrust per weight. The spread sheet then calculates the mass ratio and the approximate payload. The exit velocity of the fuel is chosen base on the above choice of fuel. Exit velocity is the product of gravity multiplied by the ISP.[4]<br />
<br />
This section also attempts to calculate gravity losses (Also known as Gravity Drag) [5]. The optimal flight ground launch flight path is such that the vertical component of the thrust cancels out gravity and the rest of the thrust is direct to horizontal acceleration.<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.<br />
<br />
[4] http://en.wikipedia.org/wiki/Specific_impulse#Specific_impulse_as_a_speed_.28effective_exhaust_velocity.29 - Relationship between specific impulse and ISP.<br />
<br />
[5] http://en.wikipedia.org/wiki/Gravity_drag - Wikipedia article on gravity drag.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=761RocketDesignSpreadsheet2009-08-13T09:15:03Z<p>John Creighton: /* Sec 4 Rocket Performance */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the geometry of the oxidizer tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[Pressure_Vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
== Sec-3 Fuel Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the fuel tank and the geometry of the fuel tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each fuel tank and the total structural mass of all the fuel tanks combined. <br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]]<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
The final section of the spread sheet gives some overall performance of the rocket. In this part you specify the delta V you want to travel, the thrust per weight of the engine and the overall thrust per weight. The spread sheet then calculates the mass ratio and the approximate payload. The exit velocity of the fuel is chosen base on the above choice of fuel. Exit velocity is the product of gravity multiplied by the ISP.[4]<br />
<br />
This section also attempts to calculate gravity losses (Also known as Gravity Drag) [5]. The optimal flight ground launch flight path is such that the vertical component of the thrust cancels out gravity and the rest of the thrust is direct to horizontal acceleration.<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.<br />
<br />
[4] http://en.wikipedia.org/wiki/Specific_impulse#Specific_impulse_as_a_speed_.28effective_exhaust_velocity.29 - Relationship between specific impulse and ISP.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=760RocketDesignSpreadsheet2009-08-13T09:02:11Z<p>John Creighton: /* Refferences */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the geometry of the oxidizer tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[Pressure_Vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
== Sec-3 Fuel Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the fuel tank and the geometry of the fuel tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each fuel tank and the total structural mass of all the fuel tanks combined. <br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]]<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
The final section of the spread sheet gives some overall performance of the rocket. In this part you specify the delta V you want to travel, the thrust per weight of the engine and the overall thrust per weight. The spread sheet then calculates the mass ratio and the approximate payload. The exit velocity of the fuel is chosen base on the above choice of fuel. Exit velocity is the product of gravity multiplied by the ISP.[] <br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.<br />
<br />
[4] http://en.wikipedia.org/wiki/Specific_impulse#Specific_impulse_as_a_speed_.28effective_exhaust_velocity.29 - Relationship between specific impulse and ISP.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=759RocketDesignSpreadsheet2009-08-13T09:01:44Z<p>John Creighton: /* Sec 4 Rocket Performance */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the geometry of the oxidizer tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[Pressure_Vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
== Sec-3 Fuel Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the fuel tank and the geometry of the fuel tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each fuel tank and the total structural mass of all the fuel tanks combined. <br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]]<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
The final section of the spread sheet gives some overall performance of the rocket. In this part you specify the delta V you want to travel, the thrust per weight of the engine and the overall thrust per weight. The spread sheet then calculates the mass ratio and the approximate payload. The exit velocity of the fuel is chosen base on the above choice of fuel. Exit velocity is the product of gravity multiplied by the ISP.[] <br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=758RocketDesignSpreadsheet2009-08-13T08:48:46Z<p>John Creighton: /* Sec_3-Fuel-Tank-Structure-Mass */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the geometry of the oxidizer tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[Pressure_Vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
== Sec-3 Fuel Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the fuel tank and the geometry of the fuel tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each fuel tank and the total structural mass of all the fuel tanks combined. <br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]]<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=757RocketDesignSpreadsheet2009-08-13T08:47:00Z<p>John Creighton: /* Section 2 Oxidizer Tank Structure Mass */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the geometry of the oxidizer tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[Pressure_Vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
== Sec_3-Fuel-Tank-Structure-Mass ==<br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]] ==<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=756RocketDesignSpreadsheet2009-08-13T08:43:14Z<p>John Creighton: /* Section 2 Tank Structure Mass */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Oxidizer Tank Structure Mass ==<br />
<br />
In this section you specify the structural material of the oxidizer tank and the geometry of the oxidizer tank. The spreadsheet then uses the above information to calculate the wall thickness of the tanks, the total mass of the structural material for each oxidizer tank and the total structural mass of all the oxidizer tanks combined.<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
Cmax/ (V)^(1/3) is the ratio of the maximum circumfrance to the cub root of the volume. Because the cross sectional area increases faster then the circumference, the cross section with the maximum cross sectional area, and simmillarly maximum circumfrance will be the point that determines the wall thickness of the [[pressure vessel]].<br />
<br />
Amax/V^(2/3) is the ratio of the maximum area of a cross section of the tank to the volume raised to the power 2/3. This scaling means that these quantities are independent of the volume of the tank. It should also be noted that the maximum area of the cross section for a cylinder is taken as the cross section which is parallel to the axis of the cylinder. Now prof is given to show this is the cross section with the maximum area or that it is the critical cross section for design. However, it seems sufficient for design purposes when the factor of safety is included and the estimates for the tank mass are in good agreement with those of the space shuttle. <br />
<br />
SufArea/V^(2/3) is the ratio of the surface area to the volume raised to the power of 2/3. This quantity is also independent of the volume of the tank for a given geometry. <br />
<br />
The above ratios, let the tank be scaled by volume and thus avoid the need to respecify the tank dimensions each time it is desired to see how a change in volume effects the rocket.<br />
<br />
== Sec_3-Fuel-Tank-Structure-Mass ==<br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]] ==<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=755RocketDesignSpreadsheet2009-08-13T08:25:17Z<p>John Creighton: /* Section 1 Tank Volumes and Fuel Mass */</p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number tanks for both the fuel and oxidizer. The remaining calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Tank Structure Mass ==<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
== Sec_3-Fuel-Tank-Structure-Mass ==<br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]] ==<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=754RocketDesignSpreadsheet2009-08-13T08:23:58Z<p>John Creighton: </p>
<hr />
<div>== Introduction ==<br />
<br />
The RocketDesignSpreadsheet was a spread sheet created by John Creighton with some help form Jumpboy for the purposes of trying to obtain better dry mass estimates, so that rocket performance (i.e. Mass Ratio) can be better calculated. The spreadsheet can be downloaded here:<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
The spreadsheet is described in further detail bellow. It is designed so you can make high level changes such as tank volume and overall thrust to mass ratio and it will scale the details of the rocket accordingly. This will let the user answer questions like:<br />
<br />
-Do we can much efficiency in terms of ratio ratio by increasing the size of the rocket?<br />
-How do low thrust high ISP rockets compare to high thrust low ISP rockets.?<br />
-Are gravity losses a significant factor?<br />
<br />
The Details are described further bellow.<br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
<br />
In the first part of the spread sheet you specify the tank volume and you select the fuel you want to use via a selection box. This value selected outputs an index of the selected value in the cell bellow the selection box. This cell is hidden. Cells further down uses this index to look up the values of the fuel and oxidizer density. You can also specify the number of fuel and the number of oxidizer tanks. The remain calculations, calculate the total volume of each tank, the total fuel/oxidizer mass contained in each tanke and the total mass of the fuel and the oxidizer. <br />
<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Tank Structure Mass ==<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
== Sec_3-Fuel-Tank-Structure-Mass ==<br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]] ==<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.</div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=753RocketDesignSpreadsheet2009-08-13T08:11:58Z<p>John Creighton: </p>
<hr />
<div>I started creating a rocket design spreadsheet.<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
In the current spreadsheet you choose the total tank volume of the rocket and the fuels. The spread sheet will calculate the weight and volume of each tank. There are options to select the material of the tank and the geometry of the tank. Initial tests of the spread sheet show it to be in good agreement with the mass of the shuttles external tank. The first goal of the spree sheet is to give improved dry mass estimates. Given that majority of the rockets weight is in the fuel [1], it is not necessary for a perfect mass breakdown of the rocket. <br />
<br />
== Section 1 Tank Volumes and Fuel Mass ==<br />
[[Image:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG]]<br />
<br />
== Section 2 Tank Structure Mass ==<br />
<br />
[[Image:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG]]<br />
<br />
== Sec_3-Fuel-Tank-Structure-Mass ==<br />
<br />
[[Image:Sec_3-Fuel-Tank-Structure-Mass.JPG]] ==<br />
<br />
== Sec 4 Rocket Performance ==<br />
<br />
[[Image:RDS-Sec_4-Rocket-Performance.JPG]]<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.</div>John Creightonhttp://wiki.newmars.com/index.php?title=File:RDS-Sec_4-Rocket-Performance.JPG&diff=752File:RDS-Sec 4-Rocket-Performance.JPG2009-08-13T08:10:16Z<p>John Creighton: </p>
<hr />
<div></div>John Creightonhttp://wiki.newmars.com/index.php?title=File:Sec_4_-_Performance.JPG&diff=751File:Sec 4 - Performance.JPG2009-08-13T08:08:43Z<p>John Creighton: </p>
<hr />
<div></div>John Creightonhttp://wiki.newmars.com/index.php?title=File:Sec_3-Fuel-Tank-Structure-Mass.JPG&diff=750File:Sec 3-Fuel-Tank-Structure-Mass.JPG2009-08-13T08:07:55Z<p>John Creighton: </p>
<hr />
<div></div>John Creightonhttp://wiki.newmars.com/index.php?title=File:RDS-Sec_2-Oxidizer_Tank_Structure_Mass.JPG&diff=749File:RDS-Sec 2-Oxidizer Tank Structure Mass.JPG2009-08-13T08:04:04Z<p>John Creighton: </p>
<hr />
<div></div>John Creightonhttp://wiki.newmars.com/index.php?title=File:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG&diff=748File:RDS-Sec1-Tank-Volumes-And-Fuel-Mass.JPG2009-08-13T07:59:45Z<p>John Creighton: </p>
<hr />
<div></div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=747RocketDesignSpreadsheet2009-08-10T00:42:01Z<p>John Creighton: </p>
<hr />
<div>I started creating a rocket design spreadsheet.<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
In the current spreadsheet you choose the total tank volume of the rocket and the fuels. The spread sheet will calculate the weight and volume of each tank. There are options to select the material of the tank and the geometry of the tank. Initial tests of the spread sheet show it to be in good agreement with the mass of the shuttles external tank. The first goal of the spree sheet is to give improved dry mass estimates. Given that majority of the rockets weight is in the fuel [1], it is not necessary for a perfect mass breakdown of the rocket. <br />
<br />
[[Image:Rocketdesignspreadsheet.JPG]]<br />
<br />
<br />
<br />
== Mass Breakdown of a rocket ==<br />
<br />
A simple estimate of the rocket mass breakdown was given in [1].<br />
<br />
Mass [kg]<br />
<br />
Total mass 375<br />
<br />
Propellant 300<br />
<br />
Tankage system 57<br />
<br />
Thrusters + plumbing 10.75<br />
<br />
Miscellaneous 7.25<br />
<br />
As you can see the majority of the rocket weight is in fuel. <br />
<br />
<br />
== Other Relevent Spreadsheets ==<br />
<br />
http://www.paul.enutrofal.com/<br />
<br />
== Refferences ==<br />
<br />
[1] http://www.lr.tudelft.nl/live/pagina.jsp?id=3784708c-c434-47f7-ae49-6f1f1d0d7114&lang=en - Gives The Mass Breakdown of a rocket.<br />
<br />
[2] http://dunnspace.com/alternate_ssto_propellants.htm - Information about the rocket propellents was gathered here.<br />
<br />
[3] http://en.wikipedia.org/wiki/Mass_ratio#Examples Contains a nice table for the mass ratio's of several rockets.</div>John Creightonhttp://wiki.newmars.com/index.php?title=Pressure_Vessel&diff=746Pressure Vessel2009-08-09T06:50:07Z<p>John Creighton: New page: == Wall Thickness == One requirement for the pressure vessel is for the walls to be strong enough to withstand the internal pressure. If you take a cross section of the pressure that is p...</p>
<hr />
<div>== Wall Thickness ==<br />
<br />
One requirement for the pressure vessel is for the walls to be strong enough to withstand the internal pressure. If you take a cross section of the pressure that is perpendicular to the walls, then the preasure times the area, must equal the tensile strength of the material multiplied by the circumfrance of the cross section, multiplied by the wall thickness. Including a factor of safety this becomes:<br />
<br />
<math>F_sPA=\sigma C w </math></div>John Creightonhttp://wiki.newmars.com/index.php?title=RocketDesignSpreadsheet&diff=745RocketDesignSpreadsheet2009-08-08T05:46:38Z<p>John Creighton: </p>
<hr />
<div>I started creating a rocket design spreadsheet.<br />
<br />
http://www.geocities.com/s243a/mars/rocketDesignSpreadsheet.xls<br />
<br />
Currently it only has limited functionality as can be seen from the following screen shot:<br />
<br />
[[Image:Rocketdesignspreadsheet.JPG]]</div>John Creighton