The starting of a thrust chamber has to be controlled so that a timely and even ignition of propellants is achieved and the flow and thrust are built up smoothly and quickly to their rated value (see Ref. 6-1). The initial propellant flow is less than full flow, and the starting mixture ratio is usually different from the operating mixture ratio. A low initial flow prevents an excessive accumulation of unignited propellants in the chamber.
The starting injection velocity is low, the initial vaporization, atomization, and mixing of propellants in a cold combustion chamber is incomplete, and there are local regions of lean and rich mixtures. With cryogenic propellants the initial chamber temperature can be below ambient. The optimum starting mixture is therefore only an average of a range of mixture ratios, all of which should be readily ignited. Mixture ratios near the stoichiometric mixture ratio have a high heat release per unit of propellant mass and therefore permit bringing the chamber and the gases up to equilibrium faster than would be possible with other mixtures. The operating mixture ratio is usually fuel rich and is selected for optimum specific impulse. One method of analytical modeling of the ignition of cryogenic propellants is given in Ref. 8-16.
The time delay for starting a thrust chamber ideally consists of the following time periods:
(1) time needed to fully open the propellant valves (typically 0.002 to more than 1.00 sec, depending on valve type and its size and upstream pressure);
(2) time needed to fill the liquid passage volume between the valve seat and the injector face (piping, internal injector feed holes, and cavities);
(3) time for forming discrete streams or jets of liquid propellant (sometimes gaseous propellant, if cryogenic liquid is preheated by heat of ambient temperature cooling jacket) and for initial atomization into small droplets and for mixing these droplets;
(4) time needed for droplets to vaporize and ignite (laboratory tests show this to be very short, 0.02 to 0.05 sec, but this depends on the propellants and the available heat);
(5) once ignition is achieved at a particular location in the chamber, it takes time to spread the flame or to heat all the mixed propellant that has entered into the chamber, to vaporize it, and to raise it to ignition temperature;
(6) time needed to raise the chamber to the point where combustion will be self sustaining, and then to its full pressure.
There are overlaps in these delays and several of them can occur simultaneously. The delays [items (1), (2), (3), (5), and (6) above] are longer with large injectors or large diameter chambers. Small thrusters can usually be started very quickly, in a few milliseconds, while larger units require 1 sec or more.
In starting a thrust chamber one propellant always reaches the chamber a short time ahead of the other; it is almost impossible to synchronize exactly the fuel and oxidizer feed systems so that the propellants reach the chamber simultaneously at all injection holes. Frequently, a more reliable ignition is assured when one of the propellants is intentionally made to reach the chamber first. For example, for a fuel-rich starting mixture the fuel is admitted first. Reference 8-17 describes the control of the propellant lead.
Other factors influencing the starting flows, the propellant lead or lag, and some of the delays mentioned above are the liquid pressures supplied to the injector (e.g., regulated pressure), the temperature of the propellant (some can be close to their vapor point), and the amount of insoluble gas (air bubbles) mixed with the initial quantity of propellants.
The propellant valves (and the flow passages betwen them and the injector face) are often so designed that they operate in a definite sequence, thereby assuring an intentional lead of one of the propellants and a controlled buildup of flow and mixture ratio. Often the valves are only partially opened, avoiding an accumulation of hazardous unburned propellant mixture in the chamber. Once combustion is established, the valves are fully opened and full flow may reach the thrust chamber assembly. The initial reduced flow burning period is called the preliminary stage. Section 10.5 describes the starting controls.
Full flow in the larger thrust chambers is not initiated with non-self-igniting propellants until the controller received a signal of successful ignition. The verification of ignition or initial burning is often built into engine controls using visual detection (photocell), heat detection (pyrometer), a fusible wire link, or sensing of a pressure rise. If the starting controls are not designed properly, unburnt propellant may accumulate in the chamber; upon ignition it may then explode, causing sometimes severe damage to the rocket engine. Starting controls and engine flow calibrations are discussed in Section 10.5
Non-spontaneously ignitable propellants need to be activated by absorbing energy prior to combustion initiation. This energy is supplied by the ignition system. Once ignition has begun the flame is self-supporting. The igniter has to be located near the injector in such a manner that a satisfactory starting mixture at low initial flow is present at the time of igniter activation, yet it should not hinder or obstruct the steady-state combustion process. At least five different types of successful propellant ignition systems have been used.
Spark plug ignition has been used successfully on liquid oxygen-gasoline and on oxygen-hydrogen thrust chambers, particularly for multiple starts during flight. The spark splug is often built into the injector, as shown in Fig. 9-6.
Ignition by electrically heated wires has been accomplished, but at times has proven to be less reliable than spark ignition for liquid propellants.
Pyrotechnic ignition uses a solid propellant squib or grain of a few seconds' burning duration. The solid propellant charge is electrically ignited and burns with a hot flame within the combustion chamber. Almost all solid propellant rockets and many liquid rocket chambers are ignited in this fashion. The igniter container may be designed to fit directly onto the injector or the chamber (see Fig. 8-1), or may be held in the chamber from outside through the nozzle. This ignition method can only be used once; thereafter the charge has to be replaced.
In precombustion chamber ignition a small chamber is built next to the main combustion chamber and connected through an orifice; this is similar to the precombustion chamber used in some internal combustion engines. A small amount of fuel and oxidizer is injected into the precombustion chamber and ignited. The burning mixture enters the main combustion chamber in a torchlike fashion and ignites the larger main propellant flow which is injected into the main chamber. This ignition procedure permits repeated starting of variable-thrust engines and has proved successful with the liquid oxygen-gasoline and oxygen-hydrogen thrust chambers.
Auxiliary fluid ignition is a method whereby some liquid or gas, in addition to the regular fuel and oxidizer, is injected into the combustion chamber for a very short period during the starting operation. This fluid is hypergolic, which means it produces spontaneous combustion with either the fuel or the oxidizer. The combustion of nitric acid and some organic fuels can, for instance, be initiated by the introduction of a small quantity of hydrazine or aniline at the beginning of the rocket operation. Liquids that ignite with air (zinc diethyl or aluminum triethyl), when preloaded in the fuel piping, can accomplish a hypergolic ignition. The flow diagram of the RD 170 Russian rocket engine in Fig. 10-10 shows several cylindrical containers prefilled with a hypergolic liquid, one for each of the high pressure fuel supply lines; this hypergolic liquid is pushed out (by the initial fuel) into the thrust chambers and into the pre-burners to start their ignitions.
In vehicles with multiple engines or thrust chambers it is required to start two or more together. It is often difficult to get exactly simultaneous starts. Usually the passage or manifold volumes of each thrust chamber and their respective values are designed to be the same. The temperature of the initial propellant fed to each thrust chamber and the lead time of the first quantity of propellant entering into the chambers have to be controlled. This is needed, for example, in two small thrusters when used to apply roll torques to a vehicle. It is also one of the reasons why large space launch vehicles are not released from their launch facility until there is assurance that all the thrust chambers are started and operating.
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