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NASA NTRS1999

Static Performance of a Fixed-Geometry Exhaust Nozzle Incorporating Porous Cavities for Shock-Boundary Layer Interaction Control

An investigation was conducted in the model preparation area of the Langley 16-Foot Transonic Tunnel to determine the internal performance of a fixed-geometry exhaust nozzle incorporating porous cavities for shock-boundary layer interaction control. Testing was conducted at static conditions using a sub-scale nozzle model with one baseline and 27 porous configurations. For the porous configurations, the effects of percent open porosity, hole diameter, and cavity depth were determined. All tests were conducted with no external flow at nozzle pressure ratios from 1.25 to approximately 9.50. Results indicate that baseline nozzle performance was dominated by unstable, shock-induced, boundary-layer separation at over-expanded conditions. Porous configurations were capable of controlling off-design separation in the nozzle by either alleviating separation or encouraging stable separation of the exhaust flow. The ability of the porous nozzle concept to alternately alleviate separation or encourage stable separation of exhaust flow through shock-boundary layer interaction control offers tremendous off-design performance benefits for fixed-geometry nozzle installations. In addition, the ability to encourage separation on one divergent flap while alleviating it on the other makes it possible to generate thrust vectoring using a fixed-geometry nozzle.

Asbury, Scott C.↗

NASA NTRS1994

Preparations for flight research to evaluate actuated forebody strakes on the F-18 high-alpha research vehicle

As part of the NASA High-Angle-of-Attack Technology Program (HATP), flight tests are currently being conducted with a multi-axis thrust vectoring system applied to the NASA F-18 High Alpha Research Vehicle (HARV). A follow-on series of flight tests with the NASA F-18 HARV will be focusing on the application of actuated forebody strake controls. These controls are designed to provide increased levels of yaw control at high angles of attack where conventional aerodynamic controls become ineffective. The series of flight tests are collectively referred to as the Actuated Nose Strakes for Enhanced Rolling (ANSER) Flight Experiment. The development of actuated forebody strake controls for the F-18 HARV is discussed and a summary of the ground tests conducted in support of the flight experiment is provided. A summary of the preparations for the flight tests is also provided.

Murri, Daniel G.↗

NASA NTRS1973

Satellite control mercury ion thruster

Program status is reviewed for development of an 8-cm structurally integrated ion thruster for application to attitude control and stationkeeping of synchronous satellites. The system is designed to achieve a thrust level of 1 and 2 mlb at a beam voltage of 1200 V, a propellant-utilization efficiency of 85%, and electrical efficiency of 75%. Thrust vectoring will be implemented by a dished-gird roll misalignment system. Design features are incorporated which ensure thruster life of 20,000 hours, with 10,000 thruster-start cycles and 100,000 beam-deflection cycles.

Hyman, J., Jr.↗

NASA NTRS1974

V/STOL flying qualities research using the X-22A

The X-22A has four ducted propellers and four engines. The engines are connected to a common system of rotating shafts which distribute propulsive power to the four propellers. Changes in the direction of the thrust vector are accomplished by rotating the ducts, which are interconnected so that all rotate through the same angle. Thrust magnitude is determined by a collective pitch lever, very similar to a helicopter. There are four variable stability system (VSS) controllers: thrust, pitch, roll, and yaw, and three artificial feel servos for the evaluation pilot cockpit controls, each employing electrohydraulic servos. Longitudinal flying qualities for STOL landing approach, and lateral-directional flying qualities and control power requirements for STOL landing approach are discussed. Attention is given to the data acquisition and processing system.

Smith, R. E.↗

NASA NTRS1989

A closed-form trim solution yielding minimum trim drag for airplanes with multiple longitudinal-control effectors

Airplane designs are currently being proposed with a multitude of lifting and control devices. Because of the redundancy in ways to generate moments and forces, there are a variety of strategies for trimming each airplane. A linear optimum trim solution (LOTS) is derived using a Lagrange formulation. LOTS enables the rapid calculation of the longitudinal load distribution resulting in the minimum trim drag in level, steady-state flight for airplanes with a mixture of three or more aerodynamic surfaces and propulsive control effectors. Comparisons of the trim drags obtained using LOTS, a direct constrained optimization method, and several ad hoc methods are presented for vortex-lattice representations of a three-surface airplane and two-surface airplane with thrust vectoring. These comparisons show that LOTS accurately predicts the results obtained from the nonlinear optimization and that the optimum methods result in trim drag reductions of up to 80 percent compared to the ad hoc methods.

Goodrich, Kenneth H.↗

NASA NTRS2011

Fuel-Efficient Descent and Landing Guidance Logic for a Safe Lunar Touchdown

The landing of a crewed lunar lander on the surface of the Moon will be the climax of any Moon mission. At touchdown, the landing mechanism must absorb the load imparted on the lander due to the vertical component of the lander's touchdown velocity. Also, a large horizontal velocity must be avoided because it could cause the lander to tip over, risking the life of the crew. To be conservative, the worst-case lander's touchdown velocity is always assumed in designing the landing mechanism, making it very heavy. Fuel-optimal guidance algorithms for soft planetary landing have been studied extensively. In most of these studies, the lander is constrained to touchdown with zero velocity. With bounds imposed on the magnitude of the engine thrust, the optimal control solutions typically have a "bang-bang" thrust profile: the thrust magnitude "bangs" instantaneously between its maximum and minimum magnitudes. But the descent engine might not be able to throttle between its extremes instantaneously. There is also a concern about the acceptability of "bang-bang" control to the crew. In our study, the optimal control of a lander is formulated with a cost function that penalizes both the touchdown velocity and the fuel cost of the descent engine. In this formulation, there is not a requirement to achieve a zero touchdown velocity. Only a touchdown velocity that is consistent with the capability of the landing gear design is required. Also, since the nominal throttle level for the terminal descent sub-phase is well below the peak engine thrust, no bound on the engine thrust is used in our formulated problem. Instead of bangbang type solution, the optimal thrust generated is a continuous function of time. With this formulation, we can easily derive analytical expressions for the optimal thrust vector, touchdown velocity components, and other system variables. These expressions provide insights into the "physics" of the optimal landing and terminal descent maneuver. These insights could help engineers to achieve a better "balance" between the conflicting needs of achieving a safe touchdown velocity, a low-weight landing mechanism, low engine fuel cost, and other design goals. In comparing the computed optimal control results with the preflight landing trajectory design of the Apollo-11 mission, we noted interesting similarities between the two missions.

guidance logic↗

NASA NTRS1975

Effect of attitude constraints on solar-electric geocentric transfers

The present work assesses the increase in flight time and fuel consumption due to introducing attitude constraints on both the thrust vector and the plane of the solar cell arrays on geocentrically orbiting spacecraft. A modified version of the SECKSPOT computer program calculates nearly time-optimal trajectories for the constrained case of zero pitch and roll. Unconstrained cases are generated with the SECKSPOT code. It is concluded that with a pitch constraint but without a roll constraint, power would not be a function of thrust direction, and so the time-optimal thruster direction would be along the projection of the primer vector in the plane normal to the radius vector. The roll constraint would cause power to become a function of thrust angle and sun angle. For certain sun angles the locus of the ratio of power to maximum power is concave and thus there may be jumps in the control angle. Comparisons are made for a SERT-C type mission between constrained and unconstrained cases in an inverse square gravity field.

Sackett, L. L.↗

NASA NTRS1981

V/STOL status from the engine technology viewpoint

Directions for designs and powerplant configurations for V/STOL aircraft are explored. The Harrier is the only V/STOL aircraft in service, having performed for over 12 yr. Rolls-Royce has studied engines with powered lift separation from forward propulsion, and engines with composite lift/propulsion modes, resulting in concentration on vectored thrust using rotating or deflected nozzles. The necessity for higher combat performance creates a need for higher specific thrust for supersonic flight (Mach 2), which is achievable with plenum chamber burning. The core flows are exhausted through separate nozzle systems for independent control. Intake in a V/STOL follows a short path and results in nonuniform flow, and testing is described to account for the effects of nonuniform static pressures and temperatures.

Lewis, G. M.↗

NASA NTRS1989

NASA/USRA advanced space design program: The laser powered interorbital vehicle

A preliminary design is presented for a low-thrust Laser Powered Interorbital Vehicle (LPIV) intended for cargo transportation between an earth space station and a lunar base. The LPIV receives its power from two iodide laser stations, one orbiting the earth and the other located on the surface of the moon. The selected mission utilizes a spiral trajectory, characteristic of a low-thrust spacecraft, requiring 8 days for a lunar rendezvous and an additional 9 days for return. The ship's configuration consists primarily of an optical train, two hydrogen plasma engines, a 37.1 m box beam truss, a payload module, and fuel tanks. The total mass of the vehicle fully loaded is 63300 kg. A single plasma, regeneratively cooled engine design is incorporated into the two 500 N engines. These are connected to the spacecraft by turntables which allow the vehicle to thrust tangentially to the flight path. Proper collection and transmission of the laser beam to the thrust chambers is provided through the optical train. This system consists of the 23 m diameter primary mirror, a convex parabolic secondary mirror, a beam splitter and two concave parabolic tertiary mirrors. The payload bay is capable of carrying 18000 kg of cargo. The module is located opposite the primary mirror on the main truss. Fuel tanks carrying a maximum of 35000 kg of liquid hydrogen are fastened to tracks which allow the tanks to be moved perpendicular to the main truss. This capability is required to prevent the center of mass from moving out of the thrust vector line. The laser beam is located and tracked by means of an acquisition, pointing and tracking system which can be locked onto the space-based laser station. Correct orientation of the spacecraft with the laser beam is maintained by control moment gyros and reaction control rockets. Additionally an aerobrake configuration was designed to provide the option of using the atmospheric drag in place of propulsion for a return trajectory.

Source record↗

NASA NTRS1989

Laser powered interorbital vehicle

A preliminary design of a low-thrust Laser Powered Interorbital Vehicle (LPIV) intended for cargo transportation between an Earth space station and a lunar base is presented. The selected mission utilizes a spiral trajectory, characteristic of a low-thrust spacecraft, requiring eight days for a lunar rendezvous and an additional nine days for return. The ship's configuration consists primarily of an optical train, two hydrogen plasma engines, a 37.1 m box-beam truss, a payload module, and propellant tanks. The total mass of the vehicle, fully loaded, is 63,300 kg. A single plasma, regeneratively cooled engine design is incorporated into the two 500 N engines. These are connected to the spacecraft by turntables that allow the vehicle to thrust tangential to the flight path. Proper collection and transmission of the laser beam to the thrust chambers is provided through the optical train. This system consists of a 23-m-diameter primary mirror, a convex parabolic secondary mirror, a beam splitter, and two concave parabolic tertiary mirrors. The payload bay is capable of carrying 18,000 kg of cargo and is located opposite the primary mirror on the main truss. Fuel tanks carrying a maximum of 35,000 kg of liquid hydrogen are fastened to tracks that allow the tanks to be moved perpendicular to the main truss. This capability is required to prevent the center of mass from moving out of the thrust vector line. The laser beam is located and tracked by means of an acquisition, pointing, and tracking system that can be locked onto the space-based laser station. Correct orientation of the spacecraft with the laser beam is maintained by control moment gyros and reaction control rockets. In addition, an aerobrake configuration was designed to provide the option of using the atmospheric drag in place of propulsion for a return trajectory.

Clarke, M. T.↗

NASA NTRS1999

Control of a Thrust Alignment Table for Modeling the Coning Dynamics of a Spinning Spacecraft with a Follower Force

This document presents a system controlling the motion of a spherical air bearing used in the modeling of spacecraft dynamics and controls in a laboratory environment. The system is part of the Spinning Rocket Simulator (SRS), used to simulate the coning of spacecraft during a thrusting stage. The reaction force at the spherical air bearing supporting the spacecraft model must coincide with the thrust axis of the model for proper simulation. Therefore, the bearing is translated in a circular path to introduce a centrifugal force. This horizontal force along with the gravitational reaction force at the bearing combines to simulate the direction of the spacecraft's thrust force. The control system receives attitude information from the spacecraft model via a laser beam embedded in the model that impinges on a photosensitive array. The non-linear system is controlled using high-speed lookup tables and digital techniques. A vector-controlled motor and a stepper motor are given the necessary signals to accurately control the turntable and platform supporting the air bearing. Preliminary performance data is presented. Mechanical elements of the table and platform are described in detail. A wireless (RF) data path for all devices on the spacecraft model to an off-table command computer is also described.

Halsmer, Dominic↗

NASA NTRS2002

Automated Escape Guidance Algorithms for An Escape Vehicle

An escape vehicle was designed to provide an emergency evacuation for crew members living on a space station. For maximum escape capability, the escape vehicle needs to have the ability to safely evacuate a station in a contingency scenario such as an uncontrolled (e.g., tumbling) station. This emergency escape sequence will typically be divided into three events: The fust separation event (SEP1), the navigation reconstruction event, and the second separation event (SEP2). SEP1 is responsible for taking the spacecraft from its docking port to a distance greater than the maximum radius of the rotating station. The navigation reconstruction event takes place prior to the SEP2 event and establishes the orbital state to within the tolerance limits necessary for SEP2. The SEP2 event calculates and performs an avoidance burn to prevent station recontact during the next several orbits. This paper presents the tools and results for the whole separation sequence with an emphasis on the two separation events. The fust challenge includes collision avoidance during the escape sequence while the station is in an uncontrolled rotational state, with rotation rates of up to 2 degrees per second. The task of avoiding a collision may require the use of the Vehicle's de-orbit propulsion system for maximum thrust and minimum dwell time within the vicinity of the station vicinity. The thrust of the propulsion system is in a single direction, and can be controlled only by the attitude of the spacecraft. Escape algorithms based on a look-up table or analytical guidance can be implemented since the rotation rate and the angular momentum vector can be sensed onboard and a-priori knowledge of the position and relative orientation are available. In addition, crew intervention has been provided for in the event of unforeseen obstacles in the escape path. The purpose of the SEP2 burn is to avoid re-contact with the station over an extended period of time. Performing this maneuver properly requires knowledge of the orbital state, which is obtained during the navigation state reconstruction event. Since the direction of the delta-v of the SEPI maneuver is a random variable with respect to the Local Vertical Local Horizontal (LVLH) coordinate system, calculating the required SEP2 burn is a challenge. This problem was solved using a neural network as a model-free function approximation technique.

Flanary, Ronald↗

NASA NTRS2016

Introducing a Degree of Freedom in the Family of Generalized Logarithmic Spirals

The versatility of the family of generalized logarithmic spirals is improved by introducing a degree of freedom in the solution. The low-thrust acceleration profile now includes a control term that affects both the magnitude and the direction of the thrust. Exact and fully analytic solutions to the trajectory, the velocity, the time of flight, etc. are made available. Two integrals of motion are preserved. The first one is a generalization of the equation of the energy and depends on the values of the control parameter. The second one relates to the equation of the angular momentum. The problem of finding spiral transfers between two arbitrary state vectors reduces to solving one algebraic equation with one unknown. The degree of freedom allows fixing the time of flight of the transfer. If the time of flight is fixed, then there are two equations with two unknowns. No other iterative procedures are required. Coast arcs can be introduced in the solution naturally. An explicit expression for the maximum acceleration reached along the transfer is provided. Thanks to the symmetry properties of the solution a simple algorithm for generating periodic orbits is presented. An arbitrary number of intermediate nodes can be introduced to improve the flexibility of the solution when facing optimization problems. An example of a low-thrust gravity-assist Earth-Mars-Ceres trajectory shows that the solution is comparable to that obtained with other preliminary design techniques.

Roa, Javier↗

NASA NTRS1985

Powered-lift technology on the threshold

A design configuration study and comparative performance evaluation is presented for current and planned V/STOL military aircraft and their powerplant and control systems, with a view to the effect of the operational capabilities uniquely furnished by such aircraft for battlefield tactics. Attention is given to V/STOL naval aircraft and the carrier or ship's deck conditions under which they must be able to operate. Thrust-vectoring and tilt rotor configurations are the most prominent among successful experimental and operational aircraft of this type; cargo aircraft have most often incorporated short/rough airfield operation capabilities rather than full VTOL. A major development thrust has been mounted for a supersonic short takeoff/vertical landing aircraft.

Deckert, W. H.↗

NASA NTRS1983

Thrust-induced effects on subsonic longitudinal aerodynamic characteristics of a vectored-engine-over-wing configuration

An investigation was conducted in the Langley 4 by 7 Meter Tunnel of the thrust induced effects on the longitudinal aerodynamic characteristics of a vectored-engine-over-wing fighter aircraft. The investigation was conducted at Mach numbers from 0.14 to 0.17 over an angle-of-attack range from -2 deg to 26 deg. The major model variables were the spanwise blowing nozzle sweep angle and main nozzle vector angle along with trailing edge, flap deflections. The overall thrust coefficient (main and spanwise nozzles) was varied from 0 (jet off) to 2.0. The results indicate that the thrust-induced effects from the main nozzle alone were small and mainly due to boundary-layer control affecting a small area behind the nozzle. When the spanwise blowing nozzles were included, the induced effects were larger than the main nozzle alone and were due to both boundary layer control and induced circulation lift. No leading edge vortex effects were evident.

Quinto, P. F.↗

NASA NTRS1985

A parametric evaluation of supersonic STOVL

This paper describes the results of a study to evaluate parametric variations to a single engine short-takeoff vertical-landing fighter/attack aircraft design. The variables considered involved thrust vectoring, thrust degradation, maximum lift, and other changes to determine the impact on short-takeoff performance, but subject to a vertical-landing capability. The results indicate that there are certain parameters that have a significant effect on short-field performance. Also, the optimal control strategies for transitions from a short-takeoff to forward flight and from forward flight to hover are determined. The results have applicability beyond the configuration evaluated.

Kidwell, G. H.↗

NASA NTRS2009

Estimating Thruster Impulses From IMU and Doppler Data

A computer program implements a thrust impulse measurement (TIM) filter, which processes data on changes in velocity and attitude of a spacecraft to estimate the small impulsive forces and torques exerted by the thrusters of the spacecraft reaction control system (RCS). The velocity-change data are obtained from line-of-sight-velocity data from Doppler measurements made from the Earth. The attitude-change data are the telemetered from an inertial measurement unit (IMU) aboard the spacecraft. The TIM filter estimates the threeaxis thrust vector for each RCS thruster, thereby enabling reduction of cumulative navigation error attributable to inaccurate prediction of thrust vectors. The filter has been augmented with a simple mathematical model to compensate for large temperature fluctuations in the spacecraft thruster catalyst bed in order to estimate thrust more accurately at deadbanding cold-firing levels. Also, rigorous consider-covariance estimation is applied in the TIM to account for the expected uncertainty in the moment of inertia and the location of the center of gravity of the spacecraft. The TIM filter was built with, and depends upon, a sigma-point consider-filter algorithm implemented in a Python-language computer program.

Lisano, Michael E.↗

NASA NTRS1999

An Investigation of Terminal Guidance and Control Techniques for a Robotic Mars Lander

Two separate Mars lander touchdown scenarios arc considered and compared o a baseline study with the goal of minimizing the landed distance to a specified location on the Mars surface. This study considers a set of points from parachute handoff to touchdown on the surface. The first scenario examines the effect of thrust vectoring while the parachute is deployed and includes an algorithm for determining targeting initial guesses. The second considers a reverse gravity turn lo a hover condition 500 meters above!meterstl e surface and then uses lateral thrusting to minimize the range to target. The effects of both scenarios on fuel usage targeting and targeting success are discussed.

Birge, Brian K.↗