Staging and Control of Rockets and Missiles:

Multi staging of Rockets:

A multistage (or multi-stage) rocket is a rocket that uses two or more stages, each of which contains its own engines and propellant. A tandem or serial stage is mounted on top of another stage; a parallel stage is attached alongside another stage. The result is effectively two or more rockets stacked on top of or attached next to each other. Taken together these are sometimes called a launch vehicle. Two stage rockets are quite common, but rockets with as many as five separate stages have been successfully launched. By jettisoning stages when they run out of propellant, the mass of the remaining rocket is decreased. This staging allows the thrust of the remaining stages to more easily accelerate the rocket to its final speed and height.

In serial or tandem staging schemes, the first stage is at the bottom and is usually the largest, the second stage and subsequent upper stages are above it, usually decreasing in size. In parallel staging schemes solid or liquid rocket boosters are used to assist with lift-off. These are sometimes referred to as ‘stage 0’. In the typical case, the first stage and booster engines fire to propel the entire rocket upwards. When the boosters run out of fuel, they are detached from the rest of the rocket (usually with some kind of small explosive charge) and fall away. The first stage then burns to completion and falls off. This leaves a smaller rocket, with the second stage on the bottom, which then fires. Known in rocketry circles as staging, this process is repeated until the final stage’s motor burns to completion.

In some cases with serial staging, the upper stage ignites before the separation- the interstate ring is designed with this in mind, and the thrust is used to help positively separate the two vehicles.

The Taurus rocket is unusual in that its ‘stage 1’ ignites in flight; this designation is used because its upper three stages are identical to those of the Pegasus rocket, with the ‘stage 0′ booster replacing the Pegasus’ carrier aircraft.


The main reason for multi-stage rockets and boosters is that once the fuel is exhausted, the space and structure which contained it and the motors themselves are useless and only add weight to the vehicle which slows down its future acceleration. By dropping the stages which are no longer useful, the rocket lightens itself. The thrust of future stages is able to provide more acceleration than if the earlier stage were still attached, or a single, large rocket would be capable of. When a stage drops off, the rest of the rocket is still traveling near the speed that the whole assembly reached at burn-out time. This means that it needs less total fuel to reach a given velocity and/or altitude.

A further advantage is that each stage can use a different type of rocket motor each tuned for its particular operating conditions. Thus the lower stage motors are designed for use at atmospheric pressure, while the upper stages can use motors suited to near vacuum conditions. Lower stages tend to require more structure than upper as they need to bear their own weight plus that of the stages above them, optimizing the structure of each stage decreases the weight of the total vehicle and provides further advantage.


On the downside, staging requires the vehicle to lift motors which are not being used until later, as well as making the entire rocket more complex and harder to build. In addition, each staging event is a significant point of failure during a launch, with the possibility of separation failure, ignition failure, and stage collision. Nevertheless the savings are so great that every rocket ever used to deliver a payload into orbit has had staging of some sort.

In more recent times the usefulness of the technique has come into question due to developments in technology. In the case of the Space Shuttle the costs of space launches appear to be mostly composed of the operational costs of the people involved, as opposed to fuel or equipment. Reducing these costs appears to be the best way to lower the overall launch costs. New technology that is mainly in the theoretical and developmental stages is being looked at to lower the costs of launch vehicles. More information can be found on single stage to orbit designs that do not have separate stages.

Stage Separation Dynamics:

A missile stage separation event occurs when the booster motor separates from the upper stage vehicle components. Active stage separation involves the firing of the upper stage motor to initiate separation, and this flight event can be effectively modeled using CFD. In some cases, stage separation can be treated using a quasi-steady assumption to decouple the relative motion of the two bodies from the fluid dynamics (this is valid if the fluid dynamic time scale is short compared to the motion of the bodies). While this approach can produce a large amount of CFD data for use in Monte Carlo flight simulations, it fails to capture the dynamics of the separation event. The Loci/CHEM team has developed modules to simulate the full six degrees of freedom (6 DOF) relative motion of the separating bodies. The methodology couples the aerodynamics and propulsion to the integrated body motion along the flight trajectory. Overset unstructured grids with advanced hole cutting and interpolation procedures provide a robust tool that can be used for the analysis of stage and store separation systems, launch simulations, or other types of multiple bodies in relative motion.

Rocket Thrust Vector Control Methods:

In addition to providing a propulsive force to a flying vehicle, a rocket propulsion system can provide moments to rotate the flying vehicle and thus provide control of the vehicle’s attitude and flight path. By controlling the direction of the thrust vectors through the mechanisms described later in the chapter, it is possible to control a vehicle’s pitch, yaw, and roll motions.

All chemical propulsion systems can be provided with one of several types of thrust vector control (TVC) mechanisms. Some of these apply either to solid, hybrid, or to liquid propellant rocket propulsion systems, but most are specific to only one of these propulsion categories. We will describe two types of thrust vector control concept: (1) for an engine or a motor with a single nozzle; and

(2) For those that have two or more nozzles.

Thrust vector control is effective only while the propulsion system is operating and creating an exhaust jet. For the flight period, when a rocket propulsion system is not firing and therefore its TVC is inoperative, a separate mechanism needs to be provided to the flying vehicle for achieving control over its attitude or flight path.

Aerodynamic fins (fixed and movable) continue to be very effective for controlling vehicle flight within the earth’s atmosphere, and almost all weather rockets, antiaircraft missiles, and air-to-surface missiles use them. Even though aerodynamic control surfaces provide some additional drag, their effectiveness in terms of vehicle weight, turning moment, and actuating power consumption is difficult to surpass with any other flight control method. Vehicle flight control can also be achieved by a separate attitude control propulsion system as described in Sections 4.6, 6.8, and 11.3. Here six or more small liquid propellant thrusters (with a separate feed system and a separate control) provide small moments to the vehicle in flight during, before, or after the operation of

The main rocket propulsion system.

The reasons for TVC are: (1) to willfully change a flight path or trajectory

(e.g., changing the direction of the flight path of a target-seeking missile); (2) to rotate the vehicle or change its attitude during powered flight; (3) to correct for deviation from the intended trajectory or the attitude during powered flight; or

(4) To correct for thrust misalignment of a fixed nozzle in the main propulsion system during its operation, when the main thrust vector misses the vehicle’s center of gravity.

Pitch moments are those that raise or lower the nose of a vehicle; yaw moments turn the nose sideways; and roll moments are applied about the main axis of the flying vehicle (Fig. 16-1). Usually, the thrust vector of the main rocket nozzle is in the direction of the vehicle axis and goes through the vehicle’s center of gravity. Thus it is possible to obtain pitch and yaw control moments by the simple deflection of the main rocket thrust vector; however, roll control usually requires the use of two or more rotary vanes or two or more separately hinged propulsion system nozzles. Figure 16-2 explains the pitch moment obtained by a hinged thrust chamber or nozzle. The side force and the pitch moment vary as the sine of the effective angle of thrust vector deflection.


Many different mechanisms have been used successfully. Several are illustrated in Refs. 16-1 and 16-2. They can be classified into four categories:

1. Mechanical deflection of the nozzle or thrust chamber.

2. Insertion of heat-resistant movable bodies into the exhaust jet; these experience aerodynamic forces and cause a deflection of a part of the exhaust gas flow.

3. Injection of fluid into the side of the diverging nozzle section, causing an asymmetrical distortion of the supersonic exhaust flow.



4. Separate thrust-producing devices that are not part of the main flow through the nozzle.

Each category is described briefly below and in Table 16-1, where the four categories are separated by horizontal lines. Figure 16-3 illustrates several TVC mechanisms. All of the TVC schemes shown here have been used in production vehicles.

In the hinge or gimbal scheme (a hinge permits rotation about one axis only, whereas a gimbal is essentially a universal joint), the whole engine is pivoted on a bearing and thus the thrust vector is rotated. For small angles this scheme has negligible losses in specific impulse and is used in many vehicles. It requires a flexible set of propellant piping (bellows) to allow the propellant to flow from the tanks of the vehicle to the movable engine. The Space Shuttle (Fig. 1-13) has two gimballed orbit maneuver engines, and three gimballed main engines.

Figures 6-1, 6-3, and 8-19 show gimballed engines. Some Soviet launch vehicles use multiple thrusters and hinges, while many U.S. vehicles use gimbals.

Jet vanes are pairs of heat-resistant, aerodynamic wing-shaped surfaces submerged in the exhaust jet of a fixed rocket nozzle. They were first used about 55 years ago. They cause extra drag (2 to 5% less is; drag increases with larger vane deflections) and erosion of the vane material. Graphite jet vanes were used in the German V-2 missile in World War II and in the Scud missiles fired by Iraq in 1991. The advantage of having roll control with a single nozzle often outweighs the performance penalties.

Small auxiliary thrust chambers were used in the Thor and early version of Atlas missiles. They provide roll control while the principal rocket engine operates. They are fed from the same feed system as the main rocket engine. This scheme is still used on some Russian booster rocket vehicles.

The injection of secondary fluid through the wall of the nozzle into the main gas stream has the effect of forming oblique shocks in the nozzle diverging


L, used with liquid propellant engines; S, used with solid propellant motors.



FIGURE 16–3. Simple schematic diagrams of eight different TVC mechanisms. Actuators and structural details are not shown. The letter L means it is used with liquid propellant rocket engines and S means it is used with solid propellant motors.

Section, thus causing an unsymmetrical distribution of the main gas flow, which produces a side force. The secondary fluid can be stored liquid or gas from a separate hot gas generator (the gas would then still be sufficiently cool to be piped), a direct bleed from the chamber, or the injection of a catalyzed monopropellant.

When the deflections are small, this is a low-loss scheme, but for large moments (large side forces) the amount of secondary fluid becomes excessive. This scheme has found application in a few large solid propellant rockets, such as Titan IIIC and one version of Minuteman.

Of all the mechanical deflection types, the movable nozzles are the most efficient. They do not significantly reduce the thrust or the specific impulse and are weight-competitive with the other mechanical types. The flexible nozzle, shown in Figs. 16-3 and 16–4, is a common type of TVC used with solid propellant motors. The molded, multilayer bearing pack acts as a seal, a load transfers bearing, and a viscoelastic flexure. It uses the deformation of a stacked set of doubly curved elastomeric (rubbery) layers between spherical metal sheets to carry the loads and allow an angular deflection of the nozzle axis. The flexible seal nozzle has been used in launch vehicles and large strategic missiles, where the environmental temperature extremes are modest. At low temperature the elastomer becomes stiff and the actuation torques increase substantially, requiring a much larger actuation system. Figure 1 6-5 describes a different type of flexible nozzle. It uses a movable joint with a toroidal hydraulic bag to transfer loads. There are double seals to prevent leaks of hot gas and various insulators to keep the structure below 200°F or 93°C.


FIGURE 16-4. Two methods of using flexible nozzle bearings with different locations for the center of rotation. The bearing support ring is made of metal or plastic sheet shims formed into rings with spherical contours (white) bonded together by layers of molded elastomer or rubber (black stripes). Although only five elastomeric layers are shown for clarity, many flexible bearings have 10 to 20 layers.


FIGURE 16-5. Simplified cross section of an upper-stage solid propellant rocket motor (IUS) using an insulated carbon-fiber/carbon-matrix nozzle, an insulated Kevlar filament- wound case, a pyrogen igniter, forward and aft stress-relieving boots, a fluid-filled bearing, and an elastomeric seal assembly in the nozzle to allow 4½ ° of thrust vector deflection. This motor has a loaded weight of 22,874 lbf, a propellant with hydroxyl terminated polybutadiene binder, a weight of 21,400 lbf, a burnout weight of 1360 lbf, a motor mass fraction of 0.941, a nozzle throat diameter of 6.48 in., and a nozzle exit area ratio of 63.8. The motor burns 146 sec at an average pressure of 651 psi (886 psi maximum) and an average thrust of 44,000 lbf (60,200 lbf maximum), with an effective altitude specific impulse of 295 sec. Top drawing is cross section of motor; bottom drawing is enlarged cross section of nozzle package assembly.


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