Rockets of Today

— stats and specs explained —

For each rocket, I will summarize the rocket’s key specifications and statistics at the bottom, in an abbreviated format. This summary includes:

Here’s an example for an actual rocket. This is a modern rocket without much variation to speak of, and a nice simple set of stats: the Electron by Rocket Lab.

Electron: mass 12.5 t (early ones were lighter), diam 1.2 m, thrust 224 kN, imp 3.4 km/s, type Mk, payload 0.3 t (2.4%), cost $20M/t, record 22/1/2 as of Apr 14, 2022.

And here’s a more complicated rocket, with more options: the Atlas V by ULA. Before the stats, we name the specific configuration of the rocket to which these stats apply, which will usually be the lightest and simplest option. In this case a note is added on the payload capacity to show the gains when optional strap-ons are used. When it comes to the record, this rocket’s history is so long that we add a popup which can be viewed by hovering the mouse over the record text (or touching the spot on mobile devices) to view the records of historical versions. Only a few of the oldest rockets have a popup like this. Some others have a parenthetical note of the record for their legacy versions.

Atlas V 401 (no extra boosters): mass 334 t, diam 3.8 m, thrust 3827 kN, imp 3.3 km/s, type ZOk(+S), payload 9.8 t (2.9%) [18.5 t (3.2%) with 5 boosters], cost $11M/t, record for “V” (as of Mar 19, 2022) 92/0/0! (0 crewed) — for legacy versions about 268/19/47 (4 crewed)
Records of Atlas orbital versions:
early ad-hoc  ’58–’95  45/10/8
Atlas-Able ’59–’60 0/0/3
Mercury-Atlas  ’59–’63 4/4/2 (4 crewed)
Atlas-Agena ’60–’78 90/0/20
Atlas-Centaur  ’63–’97 60/5/14
Atlas II ’91–’04 63/0/0
Atlas III ’00–’05 6/0/0
— so far: —
Atlas V ’02– 92/0/0 (0 crewed)
.

(The exclamation point is just because the Atlas V’s record of 92 launches without ever losing a single payload is a unique accomplishment unmatched by any other rocket. The second most used rocket with a perfect record is the Delta IV, also from ULA, with 44 flights if you include the Heavy version. “0 crewed” is present because though the Atlas V has not flown any people, it was supposed to have done so by now, and hasn’t only because the Starliner capsule is still not ready.)

For historical comparison, a space shuttle could theoretically lift 27.5 metric tons to low orbit, at a cost of around $18 million per ton in today’s money. The Saturn V could lift 135 tons — a number that has never been attempted again since, but probably will be in the next decade. The cost might have been about $14 million per ton incrementally in today’s dollars, but if you prorate the development budget over the small number of launches that were made, each one cost many times that.

For some of the more interesting rockets, this summary of stats will be followed by a link labeled “[show stages]”. If you click this, it will reveal a table of more detailed facts about each stage of the rocket, including optional ones. Depending on how many stages there are and how wide your browser window is, you may have to scroll the table horizontally to see all of them. Once shown, the link changes to “[hide stages]”. For some types which have more than one rocket in the family, there may be a link for each; in other cases there may be a single combined one, if some stages are shared between more than one model. Combined ones are described by a note in parentheses after the link.

The stats shown include many of those above: mass, diameter, thrust, specific impulse, fuel type, and engine type. Several other figures are also included. To clarify everything, it may be best to look at a real example. We will use the Atlas V for that purpose:

Stage name AJ-60A (early) GEM-63 (2020+) Atlas CCB Centaur
Role (pos) count booster (S) ×0-5 booster (S) ×0-5 core (1) upper (2)
Diameter (m)   1.58   1.61   3.81   3.05
Liftoff mass (t) 46.7 49.3 305.1  23.1
Empty mass (t)  2.2  5.2 21.1    2.2 *
Fuel mass (t) ~13    ~13.5  ~76.3  ~3.0
Oxidizer mass (t) ~30    ~30.5  ~208     ~17.7 
Fuel type HTPB HTPB+Al kerosene hydrogen
Engine Aerojet-Rocketdyne
AJ-60A
Northrop-Grumman
GEM-63
Energomash
RD-180
RL-10C ×1-2
Power cycle solid solid staged (ZO) expander (EC)
Chamber pres. (bar) 267    24  
Ox./fuel ratio   2.3?   2.3?   2.72   5.88
Thrust, vac max (kN) 1690     1650     4152      106.3 *
Thrust, SL initial (kN) ~1110      3827    
Spec. imp, vac (km/s)   2.74   2.74   3.31   4.41
Total imp, vac (t·km/s) 117    120    943    ~91   

Starting from the top, the first row gives each stage’s name. For instance, the second stage is named “Centaur” in this case. Sometimes no definite name is known, but most commonly, it’s just some jumble of letters and numbers. There may be parenthetical notes here to indicate stages which are used in only some versions of the rocket.

The second row gives the stage’s role and position. In the main stack, the position is given as a number in parentheses, counting from the bottom, so the main booster is designated as “(1)”. Auxiliary boosters stuck onto the side are designated as “(S)”. Before this is a role description word, which uses terms such as “core” to designate the main booster that supports the rest of the stack, “upper” for a stage above that which is needed to reach orbit, or “kick” for a small stage used only after orbit has already been reached. After the position number there may be a multiplier: in this example, the side booster is followed by “×0-5”, which means that a given launch may use no such boosters, or any number up to five. Or the word “opt” may appear here if the stage is optional, which generally applies to topmost stages.

The next few rows are similar to values already listed in the short summary, but broken down for each separate stage. The diameter may be the same for upper stages as for the core, or may vary; if side boosters are used, they are generally narrower. The liftoff mass gives the total heft when fully fueled and ready. The following row gives the empty or “dry” mass — the weight of the hardware alone with no consumables included — which is an important figure in calculating the overall delta-V capability of the stage. In some cases this may not be known. The next two rows hopefully give the mass of propellant used, with the fuel and oxidizer listed separately. Often these figures are not given out, but if we know the fuel/oxidizer ratio and the dry mass, they can be calculated. Values that are calculated or estimated are designated, as in this example, with a “~” character before the number, to indicate that it may be approximate. The ratio itself is given several rows further down, expressed as how many tons of oxidizer are consumed per ton of fuel.

The next several rows shift from describing the stage as a whole to describing the motor which drives it. First we name what kind of fuel it burns, which may be an abbreviation in cases like the “HTPB” in the booster column (which stands for hydroxyl-terminated polybutadiene, a type of urethane similar to the artificial rubber in car tires). Then we give the name of the engine, which often includes the name of the company that made it, for engines not made in-house by the rocket builder — “Energomash”, for instance, is the Russian company which builds the RD-180. For solid fuel there isn’t usually a distinction between the name of the motor and the name of the stage.

Next comes the engine’s power cycle, which is given as a word instead of a code for the sake of readability. But in cases where the code is more specific, it can be included in parentheses. In our example, the RD-180’s power cycle is given as “staged (ZO)”, and the code clarifies that the exact power cycle is oxygen-rich staged combustion. Likewise, the code for the RL-10C clarifies that it uses a closed expander cycle. The engine name may be followed by a multiplier, as in this case we see “×1-2” after the RL-10C, meaning that the Centaur stage can be equipped with either a single or dual engine.

Next comes the internal pressure level which the engine produces in its combustion chamber when running at full power, in bar (one bar is pretty close to sea level atmospheric pressure). Generally speaking, a higher number here means a higher performance engine. No figure is available for the solid booster because the number is both difficult to measure and rather variable over the duration of the burn. Even a weak rocket engine like the RL-10C produces astonishing pressure, when you consider that one whole side of the chamber is practically wide open, with only a mild restriction at the nozzle. This helps give one an appreciation for the colossal amounts of power and energy which get pushed through such a comparatively small device, and why the hard part of building a rocket engine is not the chamber or nozzle or bell, but the fuel pump that pushes the propellant into it. The only way to keep the pressure so high in such an open chamber is to shove in the fuel and oxidizer at an almost impossibly rapid rate.

The final engine stat is the oxidizer to fuel ratio, as already mentioned above. Sometimes these numbers are calculated, estimated, or outright guessed — for instance, in the case of the solid booster, the exact mix of propellants is a trade secret involving many secondary ingredients, so our ballpark estimate of 2.3 is followed by a question mark. For liquid engines the ratio may also be inexact; sometimes different versions of an engine tweak the mix ratio, and some engines even change it in midflight, in order to balance tradeoffs such as peak power vs manageable temperature.

After this, the last four rows describe the stage’s overall performance. First comes the maximum thrust under vacuum conditions, in kilonewtons. Then comes the initial thrust at sea level, which is lower. This is the type of thrust which was given in the brief summary above the table. The figure there may combine the thrust values of the core and the side boosters, but in this case it does not because the side boosters are entirely optional. Note that the sea level thrust for the solid booster is marked as estimated or approximate. This is common with solid motors as the thrust may vary considerably over the duration of the burn. For upper stages, no figure is given for sea level thrust, as it is not applicable.

The final two rows give specific impulse and total impulse. Specific impulse is as described for the brief summary: the amount of momentum produced per ton of propellant, in vacuum. Total impulse is essentially the specific impulse multiplied by the supply of propellant — the total amount of momentum which a rocket stage is able to produce before running out. This measures the rocket’s overall motive capability. This stat is the best gauge of how much rocket you’re getting for your money, and for solid motors, it’s the one essential stat that cuts through the approximations and equivocations to give you the real picture.

If a value in this table has some nuance which needs a footnote to explain it, you’ll see a “*” after the contents of the field. This indicates that there’s a text note which can be seen by hovering the mouse over the cell. For example, the vacuum thrust figure for the Centaur stage has an asterisk, and if you hover the mouse you see “Doubled if two engines are used.” Likewise, the Empty Mass value has an asterisk, and a note estimating that the value might be about 2.5 tons with dual engines. Sorry, as yet there is no way to view these supplementary notes in most mobile browsers.