Resolving the Challenge of Aerodynamic Damping

Researchers in the early supersonic era also faced the challenges posed by the lack of aerodynamic damping. Aerodynamic damping is the nat­ural resistance of an airplane to rotational movement about its center of gravity while flying in the atmosphere. In its simplest form, it consists of forces created on aerodynamic surfaces that are some distance from the center of gravity (cg). For example, when an airplane rotates about the cg in the pitch axis, the horizontal tail, being some distance aft of the cg, will translate up or down. This translational motion produces a vertical lift force on the tail surface and a moment (force times dis­tance) that tends to resist the rotational motion. This lift force opposes the rotation regardless of the direction of the motion. The resisting force will be proportional to the rate of rotation, or pitch rate. The faster the rotational rate, the larger will be the resisting force. The magnitude of

the resisting tail lift force is dependent on the change in angle of attack created by the rotation. This change in angle of attack is the vector sum of the rotational velocity and the forward velocity of the airplane. For low forward velocities, the angle of attack change is quite large and the natural damping is also large. The high aerodynamic damping associ­ated with the low speeds of the Wright brothers flights contributed a great deal to the brothers’ ability to control the static longitudinal insta­bility of their early vehicles.

At very high forward speed, the same pitch rate will produce a much smaller change in angle of attack and thus lower damping. For practical purposes, all aerodynamic damping can be considered to be inversely proportional to true velocity. The significance of this is that an airplane’s natural resistance to oscillatory motion, in all axes, disappears as the true speed increases. At hypersonic speeds (above Mach 5), any rota­tional disturbance will create an oscillation that will essentially not damp out by itself.

As airplanes flew ever faster, this lightly damped, oscillatory ten­dency became more obvious and was a hindrance to accurate weap­ons delivery for military aircraft, and pilot and passenger comfort for commercial aircraft. Evaluating the seriousness of the damping chal­lenge in an era when aircraft design was changing markedly (from the straight-wing propeller-driven airplane to the swept and delta wing jet and beyond). It occupied a great amount of attention from the NACA and early NASA researchers, who recognized that it would pose a con­tinuing hindrance to the exploitation of the transonic and supersonic region, and the hypersonic beyond.[678]

In general, aerodynamic damping has a positive influence on han­dling qualities, because it tends to suppress the oscillatory tendencies of a naturally stable airplane. Unfortunately, it gradually disappears as the speed increases, indicating the need for some artificial method of suppressing these oscillations during high-speed flight. In the preelec­tronic flight control era, the solution was the modification of flight con­trol systems to incorporate electronic damper systems, often referred to as Stability Augmentation Systems (SAS). A damper system for one axis con­sisted of a rate gyro measuring rotational rate in that axis, a gain-chang­ing circuit that adjusted the size of the needed control command, and a

servo mechanism that added additional control surface commands to the commands from the pilot’s stick. Control surface commands were generated that were proportional to the measured rotational rate (feed­back) but opposite in sign, thus driving the rotational rate toward zero.

Damper systems were installed in at least one axis of all of the Century – series fighters (F-100 through F-107), and all were successful in stabilizing the aircraft in high-speed flight.[679] Development of stability augmentation systems—and their refinement through contractor, Air Force-Navy, and NACA-NASA testing—was crucial to meeting the challenge of developing Cold War airpower forces, made yet more demanding because the United States and the larger NATO alliance chose a conscious strategy of using advanced technology to generate high-leverage aircraft systems that could offset larger numbers of less-individually capable Soviet-bloc designs.[680]

Early, simple damper systems were so-called single-string systems and were designed to be "fail-safe.” A single gyro, servo, and wiring system were installed for each axis. The feedback gains were quite low, tailored to the damping requirements at high speed, at which very little control surface travel was necessary. The servo travel was limited to a very small value, usually less than 2 degrees of control surface movement. A failure in the system could drive the servo to its maximum travel, but the transient motion was small and easily compensated by the pilot. Loss of a damper at high speed thus reduced the comfort level, or weapons delivery accu­racy, but was tolerable, and, at lower speeds associated with takeoff and landing, the natural aerodynamic damping was adequate.

One of the first airplanes to utilize electronic redundancy in the design of its flight control system was the X-15 rocket-powered research air­plane, which, at the time of its design, faced numerous unknowns. Because of the extreme flight conditions (Mach 6 and 250,000-foot alti­tude), the servo travel needed for damping was quite large, and the pilot could not compensate if the servo received a hard-over signal.

The solution was the incorporation of an independent, but identical, feedback "monitoring” channel in addition to the "working” channel in each axis. The servo commands from the monitor and working channel were continuously compared, and when a disagreement was detected, the system was automatically disengaged and the servo centered. This provided the equivalent level of protection to the limited-authority fail­safe damper systems incorporated in the Century series fighters. Two of the three X-15s retained this fail-safe damper system throughout the 9-year NASA-Air Force-Navy test program, although a backup roll rate gyro was added to provide fail-operational, fail-safe capability in the roll axis.[681] Refining the X-15’s SAS system necessitated a great amount of analysis and simulator work before the pilots deemed it acceptable, particularly as the X-15’s stability deteriorated markedly at higher angles of attack above Mach 2. Indeed, one of the major aspects of the X-15’s research program was refining understanding of the complexities of hypersonic stability and control, particularly during reentry at high angles of attack.[682]

The electronic revolution dramatically reshaped design approaches to damping and stability. Once it was recognized that electronic assis­tance was beneficial to a pilot’s ability to control an airplane, the con­cept evolved rapidly. By adding a third independent channel, and some electronic voting logic, a failed channel could be identified and its sig­nal "voted out,” while retaining the remaining two channels active. If a second failure occurred (that is, the two remaining channels did not agree), the system would be disconnected and the damper would become inoperable. Damper systems of this type were referred to as fail – operational, fail-safe (FOFS) systems. Further enhancement was provided by comparing the pilot’s stick commands with the measured airplane response and using analog computer circuits to tailor servo commands so that the airplane response was nearly the same for all flight con­ditions. These systems were referred to as Command Augmentation Systems (CAS). The next step in the evolution was the incorporation of a mathematical model of the desired aircraft response into the ana­log computer circuitry. An error signal was generated by comparing

the instantaneous measured aircraft response with the desired mathe­matical-model response, and the servo commands forced the airplane to fly per the mathematical model, regardless of the airplane’s inherent aerodynamic tendencies. These systems were called "model-following.”

Even higher levels of redundancy were necessary for safe operation of these advanced control concepts after multiple failures, and the fail­ure logic became increasingly more complex. Establishing the proper "trip” levels, where an erroneous comparison would result in the exclu­sion of one channel, was an especially challenging task. If the trip levels were too tight, a small difference between the outputs of two perfectly good gyros would result in nuisance trips, while trip levels that were too loose could result in a failed gyro not being recognized in a timely manner. Trip levels were usually adjusted during flight test to provide the safest settings.

NASA’s Space Shuttle orbiter utilized five independent control system computers. Four had identical software. This provided fail-operational, fail-operational, fail-safe (FOFOFS) capability. The fifth computer used a different software program with a "get-me-home” capability as a last resort (often referred to as the "freeze-dried” control system computer).