sign in sign up
<<
Home , Aerodynamics, Aerodynamic force, Gliding flight, Hypersonic speed, Orbiter (simulator), Space Shuttle Enterprise

The shuttle was uniquely winged so it could reenter Earth’s atmosphere and fly to assigned nominal or abort strips. and The allowed the to glide and bank like an Dynamics during much of the return flight phase. This versatility, however, did not come without cost. The combined ascent and re-entry capabilities required a major government investment in new design, development, Introduction verification facilities, and analytical tools. The aerodynamic and Aldo Bordano flight control engineering disciplines needed new aerodynamic and Aeroscience Challenges Gerald LeBeau aerothermodynamic physical and analytical models. The shuttle required Pieter Buning new adaptive guidance and flight control techniques during ascent and Peter Gnoffo re-entry. Engineers developed and verified complex analysis simulations Paul Romere that could predict flight environments and vehicle interactions. Reynaldo Gomez Forrest Lumpkin The shuttle design architectures were unprecedented and a significant Fred Martin challenge to government laboratories, academic centers, and the Benjamin Kirk industry. These new technologies, facilities, and tools would Steve Brown Darby Vicker also become a necessary foundation for all post-shuttle spacecraft Ascent Flight Design developments. The following section describes a US legacy unmatched Aldo Bordano in capability and its contribution to future endeavors. Lee Bryant Richard Ulrich Richard Rohan Re-entry Flight Design Michael Tigges Richard Rohan Transition Charles Campbell Thomas Horvath

226 Engineering Innovations Aeroscience Challenges

One of the first challenges in the development of the was its aerodynamic design, which had to satisfy the conflicting requirements of a spacecraft-like re-entry into the Earth’s atmosphere where blunt objects have certain advantages, but it needed wings that would allow it to achieve an -like runway landing. It was to be the first winged vehicle to fly through the regime, providing the first real test of experimental and theoretical technology for high-speed flight. No design precedents existed to help establish necessary requirements. The decision that the first flight would carry a crew further complicated the Early conceptual designs for the Orbiter looked much like a traditional airplane with a fairly sharp challenge. Other than approach and nose, straight wings, and common horizontal and vertical stabilizers, as shown in this artist’s rendering. As a result of subsequent aerodynamic and aerothermodynamic testing and analysis, NASA made the landing testing conducted at Dryden nose more spherical to reduce heating and used a double delta planform due to the severe heating Flight Research Center, , encountered by straight wings and the horizontal stabilizer. in 1977, there would be no progressive “envelope” expansion as is typically moments that, when coupled with flight with a crew on board. This done for winged aircraft. Nor would such as and dictated that the aerodynamic test there be successful uncrewed launch , determine how a spacecraft program had to be extremely thorough. demonstrations as had been done for will fly. Aerothermodynamics focuses Further complicating this goal was the all spacecraft preceding the shuttle. on heating to the spacecraft’s fact that much of the expected flight Ultimately, engineers responsible for during flight. This information is used regime involved breaking new ground, characterizing the aeroscience in the design of the Protection and thus very little experimental data environments for the shuttle would System that shields the underlying were available for the early Space find out if their collective predictions structure from excessive . Shuttle studies. The design of the shuttle employed were correct at the same as testing—an experimental state-of-the-art aerodynamic and the rest of the world: during the launch technique used to obtain associated aerothermodynamic prediction and subsequent landing of Space data—forces air past a scaled model techniques of the day and subsequently Transportation System (STS) -1 (1 981 ). and measures data of interest, such as expanded them into previously local , total forces, or heating Aeroscience encompasses the uncharted territory. engineering specialties of aerodynamics rates. Accomplishing the testing and aerothermodynamics. For the The historical precedent of flight testing necessary to cover the full shuttle shuttle, each specialty was primarily is that it is not possible to “validate”— flight profile required the cooperation associated with analysis of flight or prove—that aerodynamic predictions of most of the major wind through the Earth’s atmosphere. are correct until vehicle performance in North America. The Space Shuttle is measured at actual flight conditions. effort was the largest such program Aerodynamics involves the study In the case of the shuttle, preflight ever undertaken by the United States. of local pressures generated over predictions needed to be accurate It involved a traditional phased the vehicle while in flight and the enough to establish sufficient approach in the programmatic design resultant integrated forces and confidence to conduct the first orbital evolution of the shuttle configuration.

Engineering Innovations 227 The shuttle started on the launch pad Aerodynamic loads decreased to Main (SSMEs). The plume composed of four primary aerodynamic fairly low levels as the shuttle flow fields blocked and diverted air elements: the Orbiter; External Tank; accelerated past about Mach 5 and the moving around the spacecraft, thus and two Solid Boosters (SRBs). atmospheric decreased with influencing pressures on the aft It built speed as it rose through the altitude, thus the aerodynamic testing and altering the vehicle’s atmosphere. Aeronautical and for the ascent configuration was aerodynamic characteristics. aerospace engineers often relate to focused on the subsonic through high Unfortunately, testing speed in terms of —the supersonic regimes. with gas plumes was significantly ratio of the speed of an object relative Other aspects of the shuttle design more expensive and time consuming to the speed of in the gas through further complicated the task for than “standard” aerodynamic testing. which the object is flying. Anything engineers. Aerodynamic interference Thus, the approach implemented was traveling at less than Mach 1 is said to existed between the shuttle’s four to use the best available testing be subsonic and greater than Mach 1 is elements and altered the resultant techniques to completely characterize said to be supersonic. The flow regime loads and aerodynamics on the basic “power-off” (i.e., no plumes) between about Mach 0.8 and Mach 1.2 neighboring elements. Also, since database. “Power-on” (i.e., with is referred to as being . various shuttle elements were designed plumes) effects were then measured to separate at different points in the from a limited number of exhaust trajectory, engineers had to consider the plume tests and added to the power-off various relative positions of the measurements for the final database. elements during separation. Yet another The re-entry side of the design also complication was the effect of plumes posed unique analysis challenges. generated by SRBs and Space Shuttle During ascent, the spacecraft continued

This photo shows clouds enveloping portions of the vehicle (STS-34 [1987]) during ascent. When the launch vehicle was in the transonic regime, shocks formed at various positions along the vehicle to recompress the flow, which greatly impacted the structural loads and aerodynamics. Such shocks, which abruptly transition the flow from supersonic to subsonic flow, were positioned at the trailing edge of the condensation “clouds” that could be seen enveloping portions of the vehicle during ascent. These clouds were created in localized areas of the flow where the pressure and conditions caused the ambient While it may be intuitive to include the major geometric elements of the launch vehicle (Orbiter, moisture to condense. External Tank, and two Solid Rocket Boosters) in aerodynamic testing, it was also important to include the plumes eminating from the three main engines on the Orbiter as well as the boosters. The tests were conducted in the 4.9-m (16-ft) Transonic Wind Tunnel at the US Air Arnold Engineering and Development Center, Tennessee.

228 Engineering Innovations Initial Flight Experience Traditionally, a flight test program was used to validate and make any necessary updates to the preflight aerodynamic database. While flight test programs use an incremental expansion of the flight envelope to demonstrate the capabilities of an aircraft, this was not possible with the shuttle. Once launched, without initiation of an abort, the shuttle was committed to flight through ascent, orbital operations, re-entry, and landing. NASA placed a heavy emphasis on comparison of the predicted vehicle performance to the observed flight performance during the first few shuttle missions, and those results showed good agreement over a majority of flight regimes. Two prominent areas, however, were Every effort was made to accurately predict a vehicle’s aerodynamic characteristics using wind tunnel deficient: predictions of the launch testing. Engineers also had to be aware of anything that could adversely affect the results. This image is of the NASA Ames Research Center 2.4 x 2.1 m (8 x 7 ft) Unitary Wind Tunnel, California. vehicle’s ascent performance, and the “trim” attitude of the Orbiter during to accelerate past the aerodynamically the effect on flow-field scaling; the the early phase of re-entry. relevant portion of the ascent trajectory. support structure used to hold the On STS -1 , the trajectory was steeper During re-entry, this speed was carried aerodynamic model in the wind tunnel than expected, resulting in an SRB deep into the atmosphere until there test section, which can affect the altitude about 3 km was sufficient atmospheric density to on the model itself; and any influence (1 .9 miles) higher than predicted. measurably dissipate the related kinetic of the wind tunnel walls. To protect Postflight analysis revealed differences energy. Therefore, the aerodynamics of against any inaccuracies in the database, between preflight aerodynamic the Orbiter were critical to the design each aerodynamic coefficient was predictions and actual aerodynamics of the vehicle from speeds as high as additionally characterized by an observed by the shuttle elements due Mach 25 down through the supersonic associated uncertainty. Great care had to higher-than-predicted pressures and subsonic regimes to landing, with to be taken to not make the uncertainties on the shuttle’s aft region. It was the higher Mach numbers being too large due to the adverse effect an subsequently determined that wind characterized by complex physical gas uncertainty would have on the design tunnel predictions were somewhat dynamics that greatly influenced the of the flight control system and the inaccurate because SRB and SSME aerodynamics and heating on the ultimate performance of the spacecraft. plumes were not adequately modeled. vehicle compared to lower supersonic In the end, given the 20,000 hours of This issue also called into question Mach numbers. wind tunnel test time consumed during the structural assessment of the wing, Challenges associated with wind tunnel the early design efforts and the 80,000 given the dependence on the preflight testing limited direct applicability to the hours required during the final phases, prediction of aerodynamic loads. actual flight environment that engineers a total of 100,000 hours of wind tunnel After additional testing and cross were interested in simulating, such as: testing was conducted for aerodynamic, checking with flight data, NASA was subscale modeling of the vehicle aerothermodynamic, and structural able to verify the structural assessment. necessary to fit in the wind tunnel and dynamic testing to characterize the various shuttle system elements.

Engineering Innovations 229 Advances in Computational Aerosciences The use of computational was eventually developed as a complementary means of obtaining aeroscience information. Engineers used computers to calculate flow-field properties around the shuttle vehicle for a given flight condition. This included pressure, shear stress, or heating on the vehicle surface, as well as density, velocity, temperature, and pressure of the air away from the vehicle. This was accomplished by numerically solving a complex set of nonlinear partial differential equations that described the motion of the fluid and satisfied a fundamental The Space Shuttle Enterprise was used to conduct approach and landing testing (1977) at the Dryden requirement for , Flight Research Center, California. In the five free flights, the crew separated the spacecraft , and energy everywhere from the and maneuvered to a landing. These flights verified the Orbiter’s in the flow field. pilot-guided approach and landing capability and verified the Orbiter’s subsonic airworthiness in preparation for the first crewed orbital flight. Given its relative lack of sophistication and maturity, coupled with the modest Another discrepancy occurred during During re-entry, the Orbiter compressed computational power afforded by the early re-entry phase of STS -1 . the air of the atmosphere as it smashed computers in the 1970s, computational Nominally, the Orbiter was designed into the atmosphere at hypersonic fluid dynamics played almost no role in to reenter in an attitude with the nose speed, causing the temperature of the the development of the Space Shuttle of the vehicle inclined 40 degrees to air to heat up thermodynamically. aerodynamic database. In the following the oncoming air. In aeronautical The temperature rise was so extreme decades, bolstered by exponential terms, this is a 40-degree angle of that it broke the chemical bonds that increases in computer capabilities and attack. To aerodynamically control hold air molecules together, continuing research, computational fluid this attitude, the Orbiter had movable fundamentally altering how the flow dynamics took on a more prominent control surfaces on the trailing edge around the Orbiter compressed and role. As with any tool, demonstrated of its wings and a large “body .” expanded. These high-temperature gas validation of results with closely related To maintain the desired angle of dynamic effects influenced the pressure experimental or flight data was an attack, the Orbiter could adjust the distribution on the aft portion of the essential step prior to its use. position of the body flap up out of , thus affecting its nominal The most accurate approach for the flow or down into the flow, trim condition. The extent to which this using wind tunnel data to validate accordingly. During STS -1 , the body effect affected the Orbiter had not been computational fluid dynamics flap deflection was twice the amount observed before; thus, it was not predictions was to directly model the than had been predicted would be replicated in the wind tunnel testing wind tunnel as closely as possible, required and was uncomfortably close used during the design phase. NASA computationally. After results were to the body flap’s deployment limit researchers developed an experimental validated at wind tunnel conditions, of 22.5 degrees. NASA determined technique to simulate this experience the computational fluid dynamics tool that the cause was “real gas effects”— using a special test gas that mimicked could be run at the flight conditions a phenomenon rooted in the behavior of high-temperature air at and used directly, or the difference high-temperature gas dynamics. the lower temperatures achieved during between the computed flight and wind tunnel testing.

230 Engineering Innovations wind tunnel predictions could be added to the baseline experimental wind tunnel measured result. Because different flight regimes have unique modeling challenges, NASA developed separate computational fluid dynamics tools that were tuned to specific flight regimes. This allowed the computational algorithms employed to be optimized for each regime. Although not available during the preflight design of the Space Shuttle, several state-of-the-art computational tools were created that contributed significantly to the subsequent success of the shuttle, providing better understanding of control surface effectiveness, aerodynamic interference effects, and damage assessment. The examples of OVERFLOW and Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) software packages were both based on traditional computational fluid dynamics methods while the digital to analog converter (DAC) software employed special-purpose algorithms that allowed it to simulate rarefied, low-density flows. The OVERFLOW computational fluid dynamics tool was optimized for lower This image depicts the geometric detail included in this high-fidelity modeling capability, as well Mach number subsonic, transonic, as some representative results produced by the OVERFLOW tool. The OVERFLOW computational fluid and supersonic flows. It was thus dynamics tool was optimized for lower Mach number subsonic, transonic, and supersonic flows. The surface pressure is conveyed by a progressive color scale that corresponds to the pressure magnitude. most applicable for ascent and late A similar color scale with a different range is used to display Mach number in the flow field. re-entry simulations. Additionally, its OVERFLOW provided extremely accurate predictions for the launch vehicle aerodynamic environments. underlying methodology was based on Color contouring depicts the nominal heating distribution on the Orbiter, where hotter colors represent an innovative and extremely flexible higher values and cooler colors represent lower values. approach for discretization of the domain around the vehicle. This was the effect of design changes to the information it provided was used to especially beneficial for analysis of a shuttle’s aerodynamic performance. predict trajectories of potential debris complex like the shuttle. Some of these directly impacted shuttle sources. OVERFLOW became a key operations, including all of the changes tool for commercial and military The development of this computational made to the tank after the Columbia transport analyses and was heavily fluid dynamics tool allowed engineers accident in 2003 to help minimize the used by industry as well as other to effectively model the requisite debris. Additionally, OVERFLOW NASA programs. geometric detail of the launch vehicle, solutions became a key element in the The LAURA package was another as well as the plumes. OVERFLOW program’s risk assessment for ascent traditional computational fluid was subsequently used to investigate debris, as the detailed flow-field

Engineering Innovations 231 dynamics code, but designed specifically to predict hypersonic flows associated with re-entry . It incorporated physical models that account for chemical reactions that take place in air at the extremely high temperatures produced as a spacecraft reenters an atmosphere, as well as the temporal speed at which these reactions take place. This was essential, as the “resident” time a fluid element was in the vicinity of the Orbiter was extremely short given that the vehicle traveled more than 20 times the and the chemical reactions taking place in the surrounding fluid occurred at a finite rate. LAURA underwent extensive validation Special computational fluid dynamics programs appropriately model the complex chemically reacting necessary to accurately predict a spacecraft’s aerodynamic characteristics and the through comparisons to a wide body aerothermodynamic heating it will experience. Heating information was needed to determine the of experimental and flight data, and it appropriate materials and thickness of the Thermal Protection System that insulated the underlying was also used to investigate, reproduce, structure of the vehicle from hot gases encountered during re-entry into Earth’s atmosphere. and answer questions associated with Color contouring depicts the nominal heating distribution on the Orbiter, where hotter colors represent the Orbiter body flap trim anomaly. higher values and cooler colors represent lower values. LAURA was used extensively during the post-Columbia accident investigation activities and played a prominent role in supporting subsequent shuttle operations. This included assessing damaged or repaired Orbiter Thermal Protection System elements, as well as providing detailed flow field characteristics. These characteristics were assessed to protect against dangerous early transitioning of the flow along the heat shield of the Orbiter from smooth to turbulent conditions, and thus

NASA used the Direct Simulation to simulate low-density flows, such as those created by maneuvering thrusters during orbital rendezvous and docking of the shuttle to the . While the method made use of a distincly different modeling technique to make its predictions, it produced the same detailed information about the Plume Source flow field as would a traditional computational Boundaries fluid dynamics technique.

232 Engineering Innovations greatly elevated heating that would have Ascent Flight Design and provide acceptable first stage endangered the vehicle and crew. performance. This was achieved by flying a precise and While traditional computational fluid NASA’s challenge was to put wings sideslip profile and by throttling the dynamics tools proved extremely on a vehicle and have that vehicle main engines to limit useful, their applicability was limited survive the atmospheric heating that to five-times-gravity loads. The Solid to denser portions of the atmosphere. occurred during re-entry into Earth’s Rocket Boosters (SRBs) had a built-in NASA recognized the need to also be atmosphere. The addition of wings throttle design that also minimized the able to perform accurate analysis of resulted in a much-enhanced vehicle maximum dynamic pressure the low-density flows. Subsequently, the with a -to- ratio that allowed vehicle would encounter and still agency invested in the development of many abort options and a greater achieve orbital insertion. a state-of-the-art computer program that cross-range capability, affording more would be applicable to low-density return-to-Earth opportunities. This During the first stage of ascent, the rarefied flows. This program was based Orbiter capability did, however, create vehicle angle of attack and dynamic on the Direct Simulation Monte Carlo a unique ascent flight design challenge. pressure produced a lift force from (DSMC) method—which is a The launch configuration was no the wings and produced vehicle simulation of a gas at the molecular longer a smooth profiled rocket. structural loading. First stage guidance level that tracks molecules though The vehicle during ascent required and control algorithms ensured that physical space and their subsequent new and complex aerodynamic and the angle of attack and sideslip did deterministic collisions with a surface structural load relief capabilities. not vary significantly and resulted in and representative collisions with other flying through a desired keyhole. molecules. The resulting software, The Space Shuttle ascent flight design The keyhole was defined by the named the DSMC Analysis Code, was optimized payload to orbit while product of dynamic pressure and angle used extensively in support of shuttle operating in a constrained environment. of attack. The product of dynamic missions to the Russian space station The Orbiter trajectory needed to pressure and sideslip maintained the and the International Space Station, restrict wing and tail structural loading desired loading on the vehicle tail. as well as during maximum dynamic pressure servicing missions. It also played a critical role in the analysis of the Mars Varying Throttle to Meet Dynamic Pressure Constraints During Ascent Global Surveyor (1 996) and the Mars Odyssey (2001) missions.

800 40 105%

) t

o Throttle o F 700 35 Dynamic Pressure Constraints Leveraging the Space e r

a 100% u

Shuttle Experience q S

/ 600 30 Points at which the s

d actual dynamic pressure Never before in the history of flight had n meets the constraints u 95% ) t o n 500 25 T P such a complex vehicle and challenging ( h e

ĺ r

c s o r l p t e a t U flight regime been characterized. l

e c P ( e

s l

D

400 20 t 90% a t e o l o As a result of this challenge, NASA p w t r t o h n l i o T

ĺ r developed new and improved K

h , 300 15 T e understanding of the associated physics, r 85% u s s

and subsequently techniques and tools e 200 10 r P to more accurately simulate them. The c 80% i aeroscience techniques and technologies m 100 5 a n that successfully supported the Space y D 0 0 75% Shuttle are useful for exporation of our 0 30 60 90120 150 . Mission Elapsed Time (Seconds)

During ascent, the shuttle’s main engines were throttled down due to dynamic pressure constraints. The goal was to get as close as possible to the constraints to maximize performance.

Engineering Innovations 233 Ascent Abort Flying Through a Keyhole During ascent, a first stage Orbiter main engine out required the shuttle Shuttle

k to return to the launch site. The

c Structural Limit a

t Optimum t on-board guidance adjusted the pitch A

Without Loads Negative f Load o Angle of profile to achieve SRB staging e l 0 Dispersions Mach Number

Attack g

n conditions while satisfying structural tor c A

e y V cit × elo and heating constraints. For a side V e Nominal Steering r u

s Keyhole to AAvoidvoid Loads Orbiter main engine out, the vehicle s e r Load was rolled several degrees so that P Dispersions c i the normal m a n y Shuttle canceled the side force induced by the D Structural Limit remaining good side engine. Also, vehicle sideslip was maintained near Load dispersions, which are mostly due to atmospheric and thrust variations, added further zero to satisfy structural constraints. constraints to the shuttle’s flight. To avoid the various load dispersions at certain Mach numbers, the shuttle had to deviate from its optimum angle of attack. After the SRBs were safely separated, second stage guidance commanded a fixed pitch attitude around 70 degrees Because day-of-launch aloft Ascent flight design was also to minimize vehicle heating and burn significantly altered vehicle angle constrained to dispose the External the fuel no longer required. This was of attack and sideslip during ascent, Tank (ET) in safe waters—either the called the fuel dissipation phase and balloon measurements were taken Indian Ocean or the Pacific Ocean— lasted until approximately 2% of the near liftoff and in proximity of or in a location where tank debris fuel remained. At this point, guidance the launch site. Based on these wind was not an issue. commanded the vehicle to turn measurements, Orbiter guidance around and fly back to the launch site After main engine cutoff and ET parameters were biased and updated using the powered explicit guidance separation, the remaining main engine via telemetry. algorithm. As the vehicle returned, fuel and oxidizer were dumped. This it was pitched down so the ET could Also during first stage, a roll event provided some additional be safely separated. Dynamic pressure maneuver was initiated after the performance capability. was also minimized so a safe re-entry vehicle cleared the tower. This roll After the shuttle became operational, could occur. maneuver was required to achieve additional ascent performance was the desired orbital inclination and During second stage ascent, a main added to provide safe orbit insertion put the vehicle in a heads-down engine failure usually required the for some heavy payloads. Many attitude during ascent. vehicle to abort to a transatlantic guidance and targeting algorithm landing site. An abort to a downrange Vehicle performance was maximized additions provided more payload landing site was preferred to a return to during second stage by a linear capability. For example, standard launch site to reduce complex trajectory steering law called powered explicit targets were replaced by direct targets, targeting and minimize the loads and guidance. This steering law guided the resulting in one Orbital Maneuvering heating environments, therefore vehicle to orbital insertion and provided System maneuver instead of two. increasing abort success. If a main abort capability to downrange abort This saved propellant and resulted in engine failure occurred late during sites or return to launch site. Ascent more payload to orbit. second stage, an abort to a safe orbit performance was maintained. If one The ascent flight design algorithms and was possible. Abort to orbit was main engine failed, an intact abort techniques that were generated for the preferred over an abort to a transatlantic could be achieved to a safe landing site. shuttle will be the foundation for ascent landing site. Once the shuttle was in Such aborts allow the Orbiter and crew flight of any new US launch vehicle. a safe orbit, the vehicle could perform to either fly at a lower-than-planned a near nominal re-entry and return to orbit or land. the planned US landing strip.

234 Engineering Innovations

Space Shuttle Ascent Abort Scenarios Normal Orbit

Abort To Orbit

Transatlantic Abort Abort Once Around

Return To Launch Site

Paci c Ocean Near Hawaii Dryden Flight Launch Research Site Europe Center, Eastern or Africa California Kennedy Space Western Center, Florida Atlantic Atlantic

Return To The shuttle had four types of intact aborts: Launch Site Return to Launch Site; Transatlantic Abort Landing; Abort to Orbit; and Abort Once Turnaround Around. The aborts are presented as they occurred in the mission timeline. The Fuel Depletion and preferred order of selecting aborts based Turnaround Prediction on performance and safety was: Abort to Flyback Orbit; Abort Once Around; Transatlantic Abort Landing; and Return to Launch Site. Return To Launch Site Selection

Solid Rocket Booster Separation Pitchdown

Main Engine Cutoff

External Tank Separation

Engineering Innovations 235 If more than one main engine failed Re-entry Flight Design initial re-entry at a speed of during ascent, a contingency abort 28,000 kph (1 7,400 mph), an was required. If a contingency altitude of 122 km (76 miles), and a The shuttle vehicle reentered the abort was called during first stage, distance of 7,600 km (4,722 miles) Earth’s atmosphere at over 28,000 km guidance would pitch the vehicle up from the runway until activation of per hour (kph) (1 7,400 mph)—about to loft the trajectory, thereby terminal area guidance (a distance nine times faster than the muzzle minimizing dynamic pressure and of about 90 km [56 miles] and 24 km speed of an M1 6 bullet. Designing allowing safe separation of the SRBs [1 5 miles] altitude from the runway). a guidance system that safely and ET. After these events, a pullout During this interval, a tremendous decelerated this rapidly moving maneuver would be performed to amount of kinetic energy was spacecraft to runway landing speeds bring the vehicle to a flight transferred into heat energy as the while respecting vehicle and crew so a crew bailout could occur. vehicle slowed down. This was all constraints was a daunting challenge, done while the crew experienced only Two engines out early during second one that the shuttle re-entry guidance about 1.5 times the acceleration of stage allowed the crew to attempt a accomplished. gravity (1 .5 g). As a comparison, landing along the US East Coast at The shuttle re-entry guidance 1g acceleration is what we feel while predefined landing strips. Two engines provided steering commands from sitting on a chair at sea level. out late in second stage allowed an abort to a transatlantic site or abort to

safe orbit, depending on the time of the Entry Guidance Drag Velocity Profile second failure.

In general, Mission Control used 40 12 )

2 2.2g ControlControl d vehicle telemetry and complex vehicle d n n Systemysttem LimitL o o SurfaceSurfacce c performance predictor algorithms to c e e TTemperaturTemperatureemperaturature Constant Dynamic S S 10 / / t assist the crew in choosing the best t LimitsLimitts Drag PressurePressure e e e e 30 (overheatingoverheaating Phase CConstraint F abort guidance targets and a safe F ( (

due to highh drag) EqEquilibriumquilibriumbrium 2 d landing site. The Abort Region d 8 GlidGlidede TransitionTransition n Phasese o “Undershoot”“Undershhoot” Phase Determinator was the primary ground c e S flight design tool that assisted Mission / s

r 20 6 e Control in making abort decisions. t e M

If communication with the ground , Equilibriumum n

o 4 Constant was lost, the crew would use on-board i Glide Boundaryndary t Heat Rate

a (lossoss of drag rreferenceefereence contrcontrolol computer data and cue cards to assist r 10 Phase e or skip out due too low drag) l (Quadratic(Quadratiatic e

in selecting the abort mode. c Velocity)Velocitcity)

c 2 “Overshoot”oot” A

g a r Pre-entryPree-eentrye y Summary D 0 0 9 8877 6 5 43214 3 2 1 0 The shuttle ascent and ascent flight 3030 2525 20 15 10 5 0 design were complex. NASA Relative VelocityVelocity,Velocityy,, KKilometers/Secondilometers/Second (Thousands of Feet/Second) developed and verified many innovative guidance algorithms to Shuttle re-entry guidance was segmented into several phases—each designed to satisfy accomplish mission objectives and unique constraints during flight. The narrow region of acceptable flight conditions was called maintain vehicle and crew safety. the “flight corridor.” The surface temperature constraints resided at the lower altitude and This legacy of flight techniques and high drag “undershoot” side of the flight corridor. In contrast, if the vehicle flew too close to the “overshoot” boundary, it would not have enough drag acceleration to reach the landing computer tools will prove invaluable site and could possibly skip back into orbit. As the vehicle penetrated deeper into the to all new spacecraft developments. atmosphere, the undershoot corridor was redefined by the vehicle control system and dynamic pressure constraints.

236 Engineering Innovations How did Space Shuttle Guidance Accomplish This Feat? First, it’s important to understand how the shuttle was controlled. Air molecules impacting the vehicle’s surface imparted a pressure or force over the vehicle’s surface. The shuttle used jets initially to control the attitude of the vehicle; however, as the dynamic pressure increased on entering denser atmosphere, the position of the body flap was used to control the angle of attack and the were Shuttle re-entry guidance generated bank angle and angle-of-attack commands. The body flap used to control bank. was used to control the angle of attack by balancing the aerodynamic forces and moments about the vehicle center of gravity. The bank angle controlled the direction of the lift vector about the Changing the angle of attack had an wind velocity vector at a fixed angle of attack. Drag, which was opposite to the wind-relative immediate effect on the drag velocity, slowed the vehicle down. Lift was normal to the drag vector and was used to change the acceleration of the vehicle, whereas rate at which the vehicle reentered the atmosphere. The total normal load force was the sum of changing the bank angle had a more the lift acceleration and drag acceleration and resulted in the force felt by the crew. gradual effect. It took time for the vehicle to decelerate into different portions of the atmosphere where ag density and speed affected drag. Dr ttle Overshoot Controlling the direction of the vehicle Li oo lift vector by banking the vehicle was T Boundary of n the primary control mechanism available io g e to achieve the desired landing target. R ridor The vehicle banked about the relative ht Cor Flig velocity vector using a combination of ntry E Undershoot aft yaw Reaction Control System jets Boundary and deflection. The lift vector moved with the vehicle as it banked n of Too about the wind vector. The angle of egio R h Drag attack was maintained constant during Muc these maneuvers by the balanced aerodynamic forces at a given body flap trim position. The vehicle banked around this wind vector, keeping the blunt side of the shield facing against the flow of the atmosphere. Banking The Entry Flight Corridor defined the atmospheric re-entry angles required for safe re-entry about the wind vector until the lift flight. Before any successful re-entry from low-Earth orbit could occur, the shuttle needed to fire pointing down accelerated the vehicle engines to place the vehicle on a trajectory that intercepted the atmosphere. This deorbit into the atmosphere. Over time, this maneuver had to be executed precisely. With too steep of a re-entry, the guidance could not compute steering commands that would stop the vehicle from overheating. With too shallow of a increased drag caused the vehicle to re-entry, the guidance could not adequately control the trajectory or, for very shallow trajectories, decelerate quickly. Banking about the even stop the vehicle from skipping back out into space. The area between these two extremes wind vector until the lift vector pointed was called the Entry Flight Corridor. up accelerated the vehicle out of the

Engineering Innovations 237 atmosphere. Over time, this decreased the drag acceleration and caused the vehicle to decelerate gradually. Boundary Layer Transition Control of the vehicle lift-and-drag acceleration by bank angle and Accurate characterization of the aerothermodynamic heating experienced by a angle-of-attack modulation were the spacecraft as it enters an atmosphere is of critical importance to the design of a Thermal two primary control parameters used Protection System. More intense heating typically requires a thicker Thermal Protection to fly the desired range and cross range System, which increases a vehicle’s . During the early phase of entry, the flow during re-entry. These concepts had near the surface of the spacecraft—referred to as the boundary layer—has a smooth to be clearly grasped before it was laminar profile. Later in the trajectory, instabilities develop in the boundary layer that possible to understand the operation of the guidance algorithm. cause it to transition to a turbulent condition that can increase the heating to the spacecraft by up to a factor of 4 over the laminar state. Subsequently, a Boundary Layer Within each guidance phase, it was Transition Flight Experiment was conceived and implemented on Space Shuttle possible to use simple equations to analytically compute how much range Discovery’s later . This experiment employed a fixed-height protuberance (speed was flown. As long as the shuttle bump) on the underside of the wing to perturb and destabilize the boundary layer. trajectory stayed “close” to reference NASA used instrumentation to measure both the elevated heating on the protuberance profiles, the guidance algorithm as well as the downstream effect so that the progression of the transition could be could analytically predict how far the captured. The experiment provided foundational flight data that will be essential for the vehicle would fly. validation of future ground-based testing techniques or computational predictions of By piecing together all of the guidance this flow phenomenon, thus helping improve the design of all future spacecraft. segments, the total range flown from the current vehicle position all the way to the last guidance phase could be predicted and compared to the actual range required to reach the target. Any difference between the analytically computed range and the required range would trigger an adjustment in the drag-velocity/energy references to remove that range error. The analytic reference profiles were computed every guidance step (1.92 seconds) during flight. In this manner, any range error caused by variations in the environment, navigated state, aerodynamics, or mass properties was sensed and A NASA team—via a US Navy aircraft—captured high-resolution, calibrated infrared imagery of ’s lower surface in addition to compensated for with adjustments to discrete instrumentation on the wing, downstream, and on the Boundary Layer the real-time computed drag-velocity Transition Flight Experiment protuberance. In the image, the red regions or drag-energy reference profiles. represent higher surface temperatures. In fact, the entire shuttle re-entry guidance system could be described as a set of interlocked drag-velocity or Constant Heat-rate Phase up outside of the vehicle. That blast drag-energy pieces that would fly the furnace was due to the high-velocity The guidance phase was required to required range to target and maintain impact of the vehicle with the air in protect the structure and interior from the constraints of flight. the atmosphere. the blast furnace of building

238 Engineering Innovations

The Thermal Protection System surface was designed to withstand extremely Typical Angle-of-Attack Profile high temperatures before the 50 temperature limits of the material were y t t i i v Higher Drag and Heat Rate a exceeded. Even after a successful m i r

45 L Reduced Range and CrossCrosss Range G

t l - landing, structural damage from heating i f o r m o i t - r L could make the vehicle un-reuseable; n

e 40 l o t s s o C n r e e therefore, it was essential that the t e e e e n n i r r r C

o g g g surface remain within those limits. g r

35 d C e e r a a D D

To accomplish this, different parts of M

, , w ) ) r c i a a o the vehicle were covered with different 30 t F h h a t p p l l types of protective material, depending S A A ( ( on local heating. 25 k k c c Lower DragD a a t t IncreasedIncreased Rangee and t The objective of the re-entry guidance t 20 Static Margin Control Limit A A Crossross RangeRa

f design during this phase was to ensure f o o

e that the heat-rate constraints of the e 15 l l g g Forward Center-of-Gravity Control Limit n Thermal Protection System were not n A compromised. That is why the constant A 10 heat-rate phase used quadratic 5 Aft Center-of-Gravity drag-velocity segments. A vehicle Control Limit following a drag acceleration profile that 0 was quadratic in velocity experienced a 0 551010 15 20 25 constant rate of heating on the Thermal Mach Number Protection System. Because the shuttle tile system was designed to radiate heat, The shuttle guidance was forced to balance conflicting trades to minimize the weight, cost, the quadratic profiles in shuttle guidance and complexity of the required subsystems, maximize re-entry performance (range and were designed to provide an equilibrium cross-range capability), and maintain constraint margins. An ideal example was the selection of a constant angle-of-attack (Alpha) profile with a linear-velocity ramp transition. It was heating environment where the amount known that a high heat-rate trajectory would minimize the tile thickness required to protect of heat transferred by the tiles and to the the substructure. An initially high Alpha trim (40 degrees) was therefore selected to reduce substructure was balanced by the Thermal Protection System mass and quickly dissipate energy. The 40-degree profile helped amount of heat radiated. This meant that shape the forward center-of-gravity control boundaries and define the hypersonic static there was a temperature at which the margin control limits provided by the body flap and ailerons. A linear ramp in the Alpha profile was then inserted to increase the lift-to-drag and cross-range capability and improve the radiant heat flux away from the surface static and dynamic stability of the vehicle. matched the rate of atmospheric heating. Once the vehicle Thermal Protection System reached this equilibrium constraint boundaries, the vehicle is dissipated. The vehicle would, as a temperature, there would no longer be a substructure was maintained at a result, fly a shorter range. If the vehicle net heat flow into the vehicle. safe temperature. The Thermal was too far away from the landing site, Protection System would be the combined velocity and reference The existence of a temperature limit undamaged and reusable, and the drag profiles were automatically on the Thermal Protection System crew would be comfortable. shifted downward, causing a reduction material implied the existence of a in the rate at which energy was maximum heat rate the vehicle could During flight, if the vehicle was too dissipated. The vehicle would, as a withstand. As long as guidance close to the landing site target, the result, fly a longer range. commanded the vehicle to achieve a velocity and reference drag profiles quadratic velocity reference that was were automatically shifted upward, at or below the surface temperature causing an increase in the rate energy

Engineering Innovations 239 A

Lateral Deadband Azimuth Error

Landing Site

80 s e e r

g 40 e D

, Azimuth e e l l 0 g g EErrorrror n n A A

k k -40 n n a a B B -80 30

20 R R s o o e l l e l l r

R 10 R Shuttle g e e e v v D e e Banking

, r r r r r s s a a o o r r l 0 l r r E E

h h t t u u m m -10 Deadband is Widened i

z After First Roll Reversal The Space Shuttle removed azimuth errors A during flight by periodically executing roll -20 reversals. These changes in the sign (plus or minus) of the vehicle bank command would

-30 shift the lift acceleration vector to the opposite side of the current orbit direction 7 665432105 4 3 2 1 0 2424 2200 16 12 8804 0 and slowly rotate the direction of travel back Relative VelocityVelocity, Kilometers/Second (Thousands of Feet/Second) toward the desired target.

Equilibrium Glide Phase balance gravitational and centrifugal communication and tracking loss due forces on the vehicle. During this to plasma shield interference) and also As the speed of the shuttle dropped phase, only the reference drag profile in change runway landing direction due below about 6,200 m/s (20,500 ft/s), the equilibrium glide phase was to landing wind changes. the constant heat-rate phase ended and modified to correct range errors. All the equilibrium glide phase began. This future phases were left at their nominal was an intermediate phase between setting. This ranging approach was Constant Drag Phase high heating and the rapidly increasing designed into the shuttle re-entry The constant drag phase began and deceleration that occurred as the guidance to reserve ranging capability. the equilibrium glide phase ended vehicle penetrated deeper into the This enabled the vehicle to when either the desired constant drag atmosphere. This phase determined the accommodate large navigation errors acceleration target of 10 m/s 2 (33 ft/s 2) drag-velocity reference required to post ionization blackout (ground

240 Engineering Innovations occurred or the transition phase trajectory-range errors and issued a previous guidance phases, and a velocity of about 3,200 m/s command to begin reducing the angle concise closed-form solution for the (10,500 ft/s) was achieved. of attack. This pitch-down maneuver range flown at higher flight-path angles prepared the vehicle for transonic and was obtained. At the end of transition During the constant drag phase, the subsonic flight. During the transition phase, the vehicle was about 90 km drag-velocity reference was computed phase, the angle of attack was reduced (56 miles) from the runway, flying to maintain constant drag acceleration and the vehicle transitioned from flying at an altitude of 24 km (1 5 miles) and on the vehicle. This constrained the on the “back side” to the “front side” a speed of 750 m/s (2,460 ft/s). accelerations on the vehicle structure of the lift-to-drag (lift acceleration and crew. It also constrained maximum divided by drag acceleration) vs. load accelerations for crew members angle-of-attack curve. A vehicle flying Summary confined to a sitting position during on the back side (at a higher angle of re-entry with normal accelerations At this point, the “unique” phase of attack) was in an aerodynamic posture directed along their spine. For the re-entry required to direct the shuttle where increasing the angle of attack shuttle, the normal force constraint was from low-Earth orbit was complete. decreased the lift-to-drag. In this set at 2.5 g maximum; however, typical Although other phases of guidance orientation, the drag on the vehicle normal force operational design was were initiated following the transition was maximized and the vehicle set at 1.5 g. The form of the phase, these flight regimes were dissipated a great deal of energy, which drag-velocity reference during this well understood and the guidance was highly desirable in the early phase was particularly simple since the formulation was tailored directly for phases of re-entry flight. A vehicle drag accelerations were held constant. airplane flight. flying on the front side of the Operationally, shuttle guidance lift-to- (or at a lower angle continued to command a high of attack) was in an aerodynamic 40-degree angle of attack during this posture where increasing the angle of phase while the velocity was rapidly attack increased the lift-to-drag. In this reduced and kinetic energy was rapidly front-side orientation, the drag was removed from the vehicle. Guidance reduced and the vehicle sliced through commanded higher drag levels to the air more efficiently. Most remove extra energy from the vehicle fly on the front side of the lift-to-drag and to attain a target site that was curve, and it was during the transition closer than the nominal prediction. phase that shuttle guidance began Guidance commanded lower drag commanding the vehicle to a flying levels to reduce the rate energy orientation that mimicked the flight removed from the vehicle and to attain characteristics of an airplane. a target site that was farther away than the nominal prediction. It was also during the transition phase that the flight-path angle became significantly steeper. This happened Transition Phase naturally as the vehicle began to dig When the velocity dropped below deeper into the atmosphere. A steeper approximately 3,200 m/s (1 0,500 ft/s), angle was what influenced the the transition phase of guidance was formulation of the shuttle guidance to entered and the constant drag phase switch from velocity to energy as the was terminated. It was during this independent variable in the reference phase that the guidance system finally drag formulation. The linear began to modulate the energy-vs.- drag-energy reference acceleration drag reference to remove final did not use a shallow flight-path angle approximation as was done in the

Engineering Innovations 241