W

Each

* VCHP = Variable Conductance Heat Pipe

* VCHP = Variable Conductance Heat Pipe

Power

Thermal control power estimates normally only consist of heater power, unless electronic controllers are used. The power required for operation of dedicated electronics used for thermal control should be allocated to the thermal system estimates.

• Operational: for worst case energy balance determination, these estimates should be made for the coldest operational scenario. Mechanical thermostats or electronic controllers with a combination of proportional, integral, and differential functionality may provide control of these heaters. Onboard software may also be utilized if sufficient temperature sensors are provided.

• Survival: these estimates should be based on a cold survival scenario. Circuits should be controlled in the most reliable (and redundant) manner, utilizing the minimum resources that would likely exist in this mode, e.g., reduced onboard software capability may not allow for heater control algorithms or fault detection and isolation routines. For example, hydrazine propulsion systems require protection from freezing (2 °C) at all times and should be considered as a survival circuit.

Telemetry and Commands

Temperature sensors that are located inside components are used to ensure the correct temperature information is monitored and a determination of the thermal health and safety status of the spacecraft or instrument can be made. Every effort should be made to provide for sufficient temperature sensors early in the program, to ensure that sufficient interfaces are available to read these sensors. Thermal commands are usually necessary only when electronic or software heater control is used. While commands can usually be added late in the design flow, the associated hardware (relays to enable or disable the circuit) must be identified early.

11.6 Structures and Mechanisms

Thomas P. Sarafin, Instar Engineering Peter G. Doukas, Lockheed Martin Astronautics James R. McCandless and William R. Britton, Lockheed Martin Astronautics

The structures and mechanisms subsystem mechanically supports all other spacecraft subsystems, attaches the spacecraft to the launch vehicle, and provides for ordnance-activated separation. The design must satisfy all strength and stiffness requirements of the spacecraft and of its interface to the booster. Primary structure carries the spacecraft's major loads; secondary structure supports wire bundles, pro-pellant lines, nonstructural doors, and brackets for components typically under 5 kg.

In this section, we describe how to develop a preliminary design for a structures subsystem. We begin by considering the spacecraft's operating environments and designing the structure with overall spacecraft packaging in mind. After conducting numerous design trades, we then assess each structural member for its most likely failure modes, possible weight savings, and need for reinforcement See Fig. 11 -25 for details.

Section 11.6.1 Section 11A2 Section 11.6.3 Section 11.6.4 Section 11.8.5

Section 11.6.1 Section 11A2 Section 11.6.3 Section 11.6.4 Section 11.8.5

Fig. 11-25. The Preliminary Design Process for Structures and Mechanisms. We move from left to right, iterating as needed, when designing the spacecraft structure.

11.6.1 Structural Requirements

Structures must endure environments from manufacture to the end of the mission. Team members should contribute from all disciplines: engineering, manufacturing, integration, test and mission operations. This interdisciplinary approach ensures coverage of all critical requirements—even those which seem minor. The following discussion of the Space Shuttle's external tank structure show why we should not overlook any event in the structure's lifetime.

The aluminum skin of the external tank must have a very tight manufacturing tolerance. Adding just 0.0254 mm (0.001 in) thickness to the entire shell of its forward tank for liquid oxygen adds 220 kg to tank mass. Special handling fixtures must cradle the tank's wall sections to keep them from collapsing during welding, as they cannot support their own weight. The nose section of the completed external tank experiences its most severe loads before launch from winds occasionally gusting against an empty and unpressurized tank. Table 11-50 lists typical mission phases and possible sources for structural requirements.

TABLE 11-50. Typical Sources for Structural Requirements by Mission Phase. The structural design must account for specific loads in every phase.

Mission Phase

Source of Requirements

Manufacture and Assembly

• Handling fixture or container reactions

• Stresses induced by manufacturing processes (welding)

Transport and Himdllng

• Crane or dolly reactions

• Land, sea, or air transport environments

Testing

• Environments from vibration or acoustic tests

• Test fixture reaction loads

Preiaunch

• Handling during stacking sequence and pre-flight checks

Launch and Ascent

• Steady-state booster accelerations

• Vibro-acoustic noise during launch and transonic phase

• Propulsion system engine vibrations

• Transient loads during booster ignition and bum-out, stage separations, vehicle maneuvers, propellant slosh, and payload fairing separation

• Pyrotechnic shock from separation events

Mission Operations

• Steady-state thruster accelerations

• Transient loads during pointing maneuvers and attitude control bums or docking events

• Pyrotechnic shock from separation events, deployments

• Thermal environments

Reentry and Landing (if applicable)

• Aerodynamic heating

• Transient wind and landing loads

The launch vehicle is the most obvious source of structural requirements, dictating the spacecraft's weight, geometry, rigidity, and strength. The launch vehicle, selected orbit, and upper stage determine the spacecraft's allowable weight See Table 18-4 for launch-vehicle data.

The core body structure and spacecraft adapter typically account for 10% to 20% of a spacecraft's dry weight Appendages, component boxes, and'most secondary structures apply to the weight of other subsystems. On the structures and mechanisms subsystem, we normally increase the estimated weight by 10% for fasteners and fittings. We should also add approximately 25% for weight growth to account for program additions, underestimating, and inadequate understanding of requirements. The spacecraft item most often underestimated or neglected is electronic wiring, sometimes approaching 10% of a spacecraft's dry weight Of course, allowances for weight growth may vary for a component or subsystem, based on its design maturity and schedule risk. As subsystem designs mature, known weights replace growth estimates.

A spacecraft's size depends on choosing the payload fairing compatible with the launch vehicle. These protective shrouds shield the spacecraft from direct air loading and contamination. The spacecraft and its fairing have a prescribed dynamic envelope, or space allocation, that takes into account expected deflection and the possible addition of thermal protection blankets. The spacecraft must be rigid enough so the fairing and spacecraft do not encroach on each other's envelope. Although the Space Shuttle does not have a traditional payload fairing, its cargo bay requires a similar envelope. See Fig. 18-8 for an example.

The spacecraft's rigidity requirements specify more than maximum deflection. A launch-vehicle structure has certain natural frequencies that respond to forces from both internal (engine oscillations) and external (aerodynamic effects) sources. The launch vehicle contractor lists known natural frequencies for each launch vehicle (see Table 18-9) and describes associated axial, bending, or torsional (twisting) modes. The spacecraft structure tailored to avoid the launch vehicle's natural frequencies will experience much lower loads. Typical resonance sources to avoid include interaction between the spacecraft and the launch vehicle's control system, oscillations in the propulsion system (pogo), aerodynamic buffeting during ascent, and bending of the solid rocket motors.

Engine thrust during launch and ascent exposes the spacecraft to steady-state acceleration along its axis. This acceleration steadily increases as a booster depletes fuel (less mass to propel), but comes to an abrupt end, or transient, at burn-out. The acceleration resumes suddenly, with another transient, as the next stage ignites. Wind gusts and vehicle maneuvers can induce lateral transients. Transients and steady-state accelerations cause inertial loads, commonly specified as load factors, or multiples of weight at sea level. Table 18-8 shows typical load factors for several launch-vehicle events.

Random vibration from engines and other sources is a critical source of load. At lift-off, the major source of random vibration is acoustic noise, which radiates from the engines to engulf the vehicle. Acoustics develop from aerodynamic turbulence when the vehicle passes through the transonic portion of its flight. Structures with high surface area and low mass, including skin sections and solar array panels, respond strongly to acoustic noise.

Load factors do not express random vibration correctly. Three parameters that help describe random vibration are distribution, frequency content, and magnitude. Typically we assume that a random spectrum has a Gaussian distribution, which determines the percentage of time the vibration is within certain limits. The frequency content is most commonly expressed as power spectral density {PSD) even though "acceleration" is more precise than "power" in this application. Vibrational power in a signal is proportional to acceleration-squared. This is divided by the frequency bandwidth over which the signal was integrated, to make the quantity independent of bandwidth. Thus, PSD is in units of g2/Hz. To illustrate the power spectral density, we use a log-log plot of g2/Hz against frequency. The square root of the area under the curve is the time history's rms value. This value equals one standard deviation, a, of the random acceleration. Figure 11-26 shows a random signal, its normal distribution, and a typical PSD plot

Pyrotechnic shock, another source of load, comes from explosive separation events involving the boosters, payload fairing, and spacecraft, as well as release mechanisms for solar panels and other deployable appendages. This shock causes high acceleration and high frequency over a very short time (see Fig. 18-11). Because shock loads attenuate quickly, they seldom damage structures removed from the immediate impulse, but they may seriously harm nearby electronic components.

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