Analytical Investigation of a Flicker-Type Roll Control for a Mach Number 6 Missile with Aerodynamic Controls Over An Altitude Range of 82,000 to 282,000 feet

An analytical investigation has been carried out to determine the responses of a flicker-type roll control incorporated in a missile which traverses a range of Mach number of 6.3 at an altitude of 82,000 feet to 5.26 at an altitude of 282,000 feet. The missile has 80 deg delta wings in a cruciform arrangement with aerodynamic controls attached to the fuselage near the wing trailing edge and indexed 450 to the wings. Most of the investigation was carried out on an analog computer. Results showed that roll stabilization that may be adequate for many cases can be obtained over the altitude range considered, providing that the rate factor can be changed with altitude. The response would be improved if the control deflection were made larger at the higher altitudes. lag times less than 0.04 second improve the response appreciably. Asymmetries that produce steady rolling moments can be very detrimental to the response in some cases. The wing damping made a negligible contribution to the response

Some Effects of Roll Rate on the Longitudinal Stability Characteristics of a Cruciform Missile Configuration as Determined from Flight Test for a Mach Number Range of 1.1. to 1.8

A model of a cruciform missile configuration having a low-aspect-ratio wing equipped with flap-type controls was flight tested in order to determine stability and control characteristics while rolling at about 5 radians per second. Comparison is made with results from a similar model which rolled at a much lower rate. Results showed that, if the ratio of roll rate to natural circular frequency in pitch is not greater than about 0.3, the motion following a step disturbance in pitch essentially remains in a plane in space. The slope of normal- force coefficient against angle of attack C(sub N(sub alpha)) was the same as for the slowly rolling model at 0 degrees control deflection but C(sub N(sub alpha)) was much higher for the faster rolling model at about 5 degrees control deflection. The slope of pitching-moment coefficient against angle of attack C(sub m(sub alpha)) as determined from the model period of oscillation was the same for both models at 0 degrees control deflection but was lower for the faster rolling model at about 5 degrees control deflection. Damping data for the faster rolling model showed considerably more scatter than for the slowly rolling model

Transonic-flutter Investigation of Wings Attached to Two Low-acceleration Rocket-propelled Vehicles

Two low-acceleration transonic-flutter vehicles were launched and flown. The first carried two test wings, one of which fluttered at M = 0.92 at a frequency of 61.4 cycles per second. The reference flutter speed determined from two-dimensional theory for an unswept wing in incompressible flow is conservative when compared to the experimental flutter speed. The second vehicle carried two test wings, one of which failed at M = 0.71 because of low-frequency divergent oscillation. Since this failure was not caused by conventional flexure-torsion flutter, no comparison with a reference flutter speed can be made

NACA Research Memorandums

Report presenting a free-flight investigation over a range of Mach numbers near zero lift to determine the aileron effectiveness and damping in roll of the full-scale Hughes Falcon missile, D configuration