A combined experimental and numerical study was completed to compare the effectiveness of canards and nosecone articulation for flight control of a Mach 2 missile. Two missile designs were tested one has a 17 to 1 fineness ratio and the other has a fineness ratio of 12. Both had a 2.7 fineness-ratio, tangent ogive nose cone. The articulation or blending section of the nose is one missile diameter long and starts one diameter aft of the ogive nosecone. The shape of the blending region is based on a unique articulation mechanism surrounded by a flexible skin. The nosecone is not in motion during the test, each articulation is a separate model. The canards used for these test were based on a common interceptor missile canard. They were placed on the constant diameter section of the missile just upstream of the articulation section. Forces were captured on these models using and internal force balance in the Trisonic tunnel facility at the US Air Force Academy. Wind tunnel tests were completed for angles of attack from -10 ° to 10° with nose deflection and canard deflections of 0°, 5° and 10°. The nose cone was also rotated to azimuthal angles of a 0°, 22.5° and 45° in addition to the angle of attack changes. The results indicated the longer model was impacted by shock reflections in the wind tunnel, reducing missile stability. It also showed that nosecone articulation produces results comparable to the canard but with a 20% reduction in forebody drag. A combination of the canards with nose deflection produced synergistic effects producing a 33% larger C_(m_0 ) than the nosecone deflection and canards combined, but combination made missile unstable. Finally, out-of-plane deflection of the nosecone provides a complex flow field that did not produce forces and moments in the direction of the deflection. In some cases the moments were reversed.

HIFiRE 6 is a test flight to validate the use of adaptive flight control on a hypersonic vehicle. The Air Force Academy Department of Aeronautics conducted wind tunnel testing of the 10.5-inch HIFiRE 6 model at the off-design Mach numbers 3.0 (Re=9,765,000), 3.5 (Re=9,975,000), and 4.3 (Re=9,765,000) to support the development of the flight control algorithm for the test flight. By examining schlieren images, and pressure data from both the wind tunnel and inside the scramjet engine inlet and isolator, engine unstart of the three dimensional inlet was characterized for ranges of both sideslip angles and angles of attack: -5.4° to 5.7° and -5.4° to 14.8°, respectively. The pressure data that indicates the start and unstart conditions were captured by five pressure ports along the inlet and four more pressure ports along the isolator of the vehicle. The combined schlieren and pressure data indicated that the HIFiRE 6 scramjet would only start at the highest Mach number tested. At Mach 4.3 the scramjet design was found to start at -5.4° angle of attack and remain started until positive 6.0° degrees with no sideslip. However, once the vehicle became unstarted it remained unstarted until it was lowered again to -5.4° angle of attack. HIFiRE 6 was able to start at 0° angle of attack when a sideslip of -5.4° was applied and stayed started until positive 5.7°. Using the unstart results a performance map was created for Mach 4.3 test to indicate the flight orientations where unstart could be expected. The Kantrowitz limit was applied to the results to show the low Mach number limit is driven by the choking of the inlet and can be predicted for the three dimensional scramjet inlets.

This a combined experimental and computational study to evaluate the aerodynamic forces resulting from the interaction of small pin protuberances and the boundary layer of a flat plate in Mach 3.0 flow. Cylindrical pins with diameters of 2 mm and 6 mm were tested. The pin heights were 2 mm, 4 mm and 6 mm and they were tested on the flat plate with a turbulent boundary layer thickness of approximately 3 mm. The 2 mm diameter pin geometry was also tested in a group of four to assess mutual interference between neighboring pins. Oil flow, schlieren imagery, and pressure data were recorded on the nine pin configurations to capture a three dimensional understanding of the flow structures. The numerical results were calculated using a Reynolds-Averaged Navier Stokes code. The numerical results were compared to the experimental results for validation. Then they were used to predict the aerodynamic forces induced by a group of protuberances. This was a first step in assessing the possibility of using an array of independently deployable pin protuberances on a missile for flight control at supersonic speeds. Under this concept both the height and number of pins deployed could be controlled. The results showed that the scale of the upstream fluid structures is a function of the pin height to diameter ratio similar to previous studies. There were some fundamental changes in the flow downstream as the pin height was increased from within the boundary layer to outside the boundary layer. A periodic pressure of signal between 300 to 400 Hz was recordeded for all cases in the wake of the pin. The side force produced by the smaller pin ranges up to 1.2 N. The side force form the larger pin ranges up to 4.2 N. The drag forces for the smaller and larger pins ranged up to 0.87 N and 2.5 N respectively. To the first order the forces scaled to the product of the pin height and the combined width of the pins in each configuration.

This a combined experimental and computational study to evaluate the aerodynamic forces resulting from the interaction of small cylindrical protuberances and the boundary layer of a cylindrical body in Mach 3.0 flow. It is a continuation of a study on the effects of the same protuberances on the flow over a flat plate model. Cylindrical pin protuberances with diameters of 2 mm and 6 mm were tested. Protuberance heights of 2 mm, 4 mm and 6 mm were tested on a 63.5 mm diameter cylinder with an ogive nosecone. The boundary layer thickness was approximately 2.6mm. The set 2 mm diameter protuberances were also tested in a group of four at all three heights. Oil flow visualization, schlieren photography, and pressure data were recorded with a variety of pin configurations to capture a three dimensional understanding of the flow structures. Both schlieren and pressure data were captured a 10 kHz to capture the unsteady flow behavior. The results were that the there was a decrease of between 5 and 9% in the side forces produced by the pins. The drag was increased in some cases but reduced by up to 31% for the largest pins. Flow visualization showed that the upstream separation point did not move due to the curvature but the lateral spread of the separation front was limited and the peak pressures decreased due to flow relief from the addition on curvature.

As the practical flight regime extends farther into the supersonic and hypersonic regimes, the interaction between shock waves and the boundary layer that forms along a body will become more significant. These interactions lead to the potential for locally separated flows and high temperature regions on the surface. Because of the limitations of high Mach number wind tunnel testing, wind tunnel tests alone will not be able to provide the information needed to design these high Mach number aircraft. This study combines supersonic (M = 3) wind tunnel test of shock-wave boundary-layer interactions with computational modeling of the same. The particular interaction chosen was the impingement of a planar shock on the boundary layer on a von Karman ogive forebody with a finess ratio of 2.3 and cylindrical aft section with diameter of 5 cm. The shock was generated by a 10 beveled wedge mounted above the ogive-cylinder to provide a shock wave impacting on the cylindrical part of the body. Schlieren photography and oil flow visualization provided details of the flowfield around the model. Time-averaged surface pressure ports and a force balance were used to measure the interaction of the shock wave with the ogive-cylinder body. In general the computational results compared well with the experimental results though several discrepancies were observed with each of the measurements. The results confirmed the existence of a shock induced horseshoe vortex around the cylindrical body, and found that the primary measurable lateral force on the ogive-cylinder body was Cy = −0.081 in the direction normal to the wedge shock generator. This normal force was primarily due to the sharp pressure rise induced near the initial point of impact of the oblique shock wave upon the receiver body.