When selecting protection devices, consideration must be given to the prospective short-circuit current (PSC) at the location the device is to be installed in (AS/NZS 3000:2018 clause 2.5.4.1). The PSC can be determined by measurement or by calculation. Many multi-function testers are able to measure the PSC.

PSC can also be calculated by undertaking a Live Fault Loop Impedance measurement (in ohms) and dividing the voltage by this measurement.

Example

Voltage tests as 240 volts and the measured fault loop impedance between incoming line and neutral is 0.08.

Max PSC = E/R

240/0.08 = 3000 A or 3kA

Three Phase Supplies
Where there is a three phase supply, the PSC is likely to be between line conductors. In this case, the PSC can be calculated by multiplying the single phase reading by the square root of 3 or 1.73. Therefore in the example above, the PSC would be 3000 x 1.73 = 5190 A or 5.19kA.

At a pressure of 1 bar, the temperature of liquid boiling oxygen stabilizes at 90 K. For sub-cooling of LOX, the temperature should be lower. It is possible to cool LOX by forced evaporation by a pressure lower than 1 bar. But the LOX tank in a rocket should be as light as possible. If the pressure inside the tank is substantially lower than outside, extra strength and weight is necessary. But according to these papers: (1) (2) and (3) there is another method.

Cold helium gas is injected at the bottom of the tank and the bubbles raise in the LOX. At the surface of the bubbles, LOX evaporates into the bubble and cools the remaining LOX. But extra space is needed for the bubbles in the LOX and for the gas mixture of helium and oxygen above the liquid level. For topping off, the injection of helium is stopped and the remaining space is filled with LOX. Figure 8 of the first paper shows the effect of different helium gas temperatures. The cooling works best with helium at 85 K, but even helium at 150 K cools the LOX.

A bubble injected into the LOX consists of 100 % helium and 0 % oxygen at first. The LOX around this bubble would boil just like in a vaccum because the partial pressure of oxygen in this bubble is zero. Even a bubble consisting of 50 % helium and 50 % oxygen is able to cool LOX at 90 K. Without sub cooling in a tank with boiling LOX at 90 K, the gas above the liquid is 100 % oxygen and the partial pressure of oxygen is 1 bar. If the partial pressure of oxygen is lower than 1 bar in the gas above the liquid or inside the bubbles, the LOX is cooled by evaporation.

At the launch pad the LOX may be precooled using a heat exchanger with ground suplied liquid nitrogen boiling at 77.355 K. To save weight of the rocket, this heat exchanger should be outside the rocket but close to it. Liquid nitrogen and oxygen should not be mixed to avoid solving of nitrogen within the LOX. Cooling with helium bubbles may be used within the rocket LOX tank.

TL;DR: Technically, the SO and YouTube licenses are (probably) incompatible, so if you want to use copyrightable elements from SO in a YouTube video, you'll have to get permission from the original author. Certainly it never hurts to ask, even if you might not strictly need it.

But there was another “incident” during the interview that CNN remarked on that truly caused a meltdown on the part of the left. At the beginning of the interview, Sinema’s ringtone on her phone went off. “Her ringtone is the refrain from a song in the musical ‘Hamilton’ that includes the lyrics “you don’t have the votes,” CNN noted. “It’s been her ringtone since 2015, the year the musical was originally released, her spokesman told CNN.” That caused the folks on the left to melt down, declaring their hate for her, while people on the right declared it a master troll. //

Dan McLaughlin
@baseballcrank
Troll level: 1790.

The same question could well be asked of the LM's descent engine and the main engine on the Apollo service module, however, which did both need to fire in free-fall. In those cases, the smaller RCS thrusters on the LM or CSM were fired first, to "settle" the tankage and separate the fuel from the helium. In the LM case, this "ullage burn" was about 7.5 seconds. The first couple of service module burns -- typically for mid-course correction while en route to the moon -- generally didn't need an ullage burn prior, as the tanks would be full of propellant with little or no volume of helium. SPS burns later in the mission did require ullage burns. The RCS thrusters produced about 100 lbs of thrust each, and four would be used for the ullage burn, yielding roughly 1/200 g acceleration.

The same RCS ullage burn technique would also apply to a situation where the descent engine failed and the ascent engine needed to be used for abort from free-fall, or in flight testing of the ascent engine.

That, in turn, raises the question of how helium ingestion was avoided in the RCS thrusters, since they were also helium pressurized. In those cases, the helium was separated from the propellants by a teflon bladder, so the helium didn't mix with the propellants. This was more practical to do on the smaller scale of the RCS propellant tanks than it would have been for the larger engines.

One might imagine that 60+ years of development must have produced large gains, but chemical rocket performance is fundamentally limited by the amount of energy in the chemical fuels, and the 1960s engines were already getting at least 2/3 of the maximum theoretically possible performance (see comparison table below). //

The usual primary metric is specific impulse.

But specific impulse is a somewhat unintuitive quantity to understand, so let's start with effective exhaust velocity, which is the average speed of an exhaust particle (in the backward direction). For example, the Rocketdyne F-1 engines used in the first stage of the Saturn V (the Apollo rocket) have an effective exhaust velocity of 2.58 km/s at sea level.

What does 2.58 km/s mean in terms of rocket performance? It means if you build a rocket whose weight is about 63% fuel, and you fire the engine in deep space until the fuel runs out, the rocket will now be going 2.58 km/s faster in whatever direction it was pointing: //

So, what is change in velocity, Δv, good for? In the solar system there are two main uses for Δv: launching from the surface to achieve orbit, and transferring from one orbit to another. The article Delta-v budget has some examples, but the most relevant to Apollo is the Δv to get into low Earth orbit from a sea level launch, which is (very roughly) around 10 km/s. That breaks down as about 8 km/s of required velocity to stay in orbit (any slower and you'll come back down) and 2 km/s spent lifting the rocket against gravity and pushing through the air on the way up. //

So let's take a quick comparison of ve for the F-1 and the SpaceX Merlin engine. This is a relatively fair comparison because both burn RP-1 (refined kerosene) and liquid oxygen in a gas-generator cycle. These characteristics are good for a first stage due to high energy density per unit volume and high thrust, although other fuels have better ve

F-1 2.58 km/s (sea level)
Merlin 2.77 km/s (sea level)
F-1 2.98 km/s (vacuum) 65% of max
Merlin 3.05 km/s (vacuum) 66% of max
Theoretical max 4.61 km/s (vacuum)

The theoretical maximum is based on the total chemical energy in the fuel. //

Finally then, what is specific impulse? It's obtained from ve

by dividing by the gravitational acceleration on Earth:

Isp=veg

where g is usually standard gravity, or about 9.81ms2. The resulting quantity has units of seconds. For example, for the F-1 at sea level, Isp=263s

What is the physical significance of Isp?

Well, consider our rocket from before with 63% fuel by mass. Suppose we start the rocket while it is sitting on the pad, let it just barely lift off, then hover just off the pad until it runs out of fuel (this assumes we can arbitrarily throttle the engine without affecting its performance, which is not realistic, but ignore that). Isp is how long it will hover. That is because, for every second of hovering, we consume 9.81 m/s of Δv in order to overcome gravitational acceleration accumulated during that second. After Isp seconds, all of our Δv is gone.

Example: Calculate Size of Bus bar having Following Details

Bus bar Current Details:
Rated Voltage = 415V,50Hz ,
Desire Maximum Current Rating of Bus bar =630Amp.
Fault Current (Isc)= 50KA ,Fault Duration (t) =1sec.
Bus bar Temperature details:

Schneider Electric's Fault Current Calculator -- Single phase or three phase

To acheive full thrust of the first stage F-1 engines of the Saturn V only liquid oxygen should be pumped into the combustion chambers. A mix of gaseous and liquid oxygen would reduce the desired mass flow of oxygen to the engines and may damage or even destroy the oxygen pumps. A reduced mass flow of oxygen would reduce thrust and endanger a successful liftoff.

But the LOX lines through the fuel tank had a bad ratio of surface to volume, much worse than the large oxygen tank. The heat flow to the LOX lines would cause the LOX inside to boil heavily.

By bubbling of cold gaseous helium through the LOX lines and tank the LOX was cooled before launch below the boiling temperature. The gaseous helium and oxygen at the top of the tank are exhausted through umbilicals and the evaporated LOX is replaced by toping of more LOX from the large tanks at the launch pad.

Just before ignition the helium bubbling is finished and the very cold LOX within the lines and tank does not boil as long as its temperature is below the boiling point. So the oxygen pumps and combustion chambers could be fed with pure liquid oxygen free from bubbles.

NASA's Space Transportation System (STS) vehicle, better known as the Space Shuttle, used two single engine Solid Rocket Boosters (SRB) as Stage 0, an engineless external tank providing propellant for the three Space Shuttle Main Engines (SSME) on the orbiter as stage 1, and additional two Orbital Maneuvering System (OMS) hypergolic liquid-propellant rocket engines on the Space Shuttle orbiter as stage 2.

The two solid rocket boosters used roughly 500,000 kg (1.1 Mlb) of a 11-star perforated solid propellant cake of Ammonium Perchlorate Composite Propellant (APCP - a mixture of of ammonium perchlorate, aluminium, iron oxide, PBAN or HTPB polymers, and an epoxy curing agent) each, that provided 124 seconds of burn time with a specific impulse (Isp) of 269 s that provided 12.5 MN of thrust per SRB and the external tank that came in three different configurations (mostly progressively reducing tank's own weight) capacity was 629,340 kg (1,387,457 lb) of cryogenic liquid oxygen (LOX) as the oxidizer and 106,261 kg (234,265 lb) of cryogenic liquid hydrogen (LH2) as the fuel components of the bipropellant LOX/LH2 that provided 480 seconds of burn time with specific impulse of 455 seconds, resulting in 5.45 MN of thrust at sea-level (for the Super Lightweight Tank or SLWT, the last and most advanced of the three versions used with STS).

So to answer your question directly, not counting the OMS propellant as per the specifics of your question, the total mass of all propellants of the SRBs (stage 0) and the external tank (stage 1) was at launch of the STS 1,735,601 kg (3,821,722 lb). The solid rocket boosters provided roughly 83% of liftoff thrust for the Space Shuttle and were the largest, most powerful solid-propellant motors flown to date.