3. Launch Phase

Launch of the MGS spacecraft will occur aboard a Delta 2/7925 launch vehicle from Space Launch Complex 17A (SLC-17A) at the Cape Canaveral Air Force Station (CCAFS). In the past, Delta rockets have primarily been used by NASA and commercial organizations to launch small to medium sized payloads to low Earth or geosynchronous orbits. The current Delta rocket design represents the latest in a long line of highly reliable and successful family of launch vehicles.

Mars Global Surveyor will represent only the second interplanetary launch for the Delta vehicle. The primary launch period for MGS spans a 22-day period from 4 November 1996 through 25 November 1996. Launch phase for this mission extends from the start of the launch countdown until separation of the MGS spacecraft from the Delta's third stage following the trans-Mars injection burn.

McDonnell Douglas Aerospace (MDA), Space Transportation Division builds the Delta 2. The launch vehicle for MGS is being procured as an option on the Medium Expendable Launch Vehicle Services (MELVS) contract, managed by NASA-OLS (Orbital Launch Services) at the Goddard Space Flight Center (GSFC).

3.1. Brief Launch Vehicle Description

Delta 2 launch vehicles consist of five major components that include the three main stages, nine solid rocket motors that attach to the first stage, and the payload fairing. The Delta measures 38.2 meters tall from the tip of the payload fairing to the bottom of the first stage's body. Including the MGS spacecraft (assumed to weigh 1,060 kg for launch vehicle planning purposes), the Delta will weigh approximately 231,447 kg at the time of launch.

3.1.1.First Stage

Stage one employs a Rocketdyne RS-27 main engine with a 12:1 expansion ratio. This engine is a single start, liquid bi-propellant rocket engine that will provide approximately 894,094 N of thrust at the time of lift-off. Its propellant load (96,000 kg) consists of RP-1 fuel (thermally stable kerosene) and liquid oxygen (LOX) for oxidizer. The RP-1 fuel tank and liquid oxygen tank on the first stage are separated by a center body section that houses control electronics, ordnance sequencing equipment, a telemetry system, and a rate gyro. In addition, stage one also employs two Rocketdyne vernier engines. They will provide roll control during the main engine burn, and attitude control between main engine cutoff (MECO) and second stage ignition.

Figure 3-1: Parts of the Delta 2 Launch Vehicle

3.1.2. Strap-on Solid Rocket Motors

A set of nine solid-propellant graphite epoxy motors (GEMs), each fueled with 12,000 kg of hydroxyl-terminated polybutadiene (HTPB) solid propellant, attach to the first stage to provide augmentation thrust. Each GEM will provide an average thrust of 439,796 N at lift-off. Six of the nine GEMs, the main engine, and the vernier engines will ignite at the time of lift-off. The remaining three GEMs will ignite 65.5 seconds into flight, shortly after the initial six burn-out. All nine will be jettisoned shortly after they burn-out.

3.1.3. Second Stage

The Delta second stage uses a re-startable, liquid bi-propellant Aerojet AJ10-118K engine that consumes a combination of Aerozine 50 fuel (a 50/50 mix of hydrazine and un-symmetric dimethyl hydrazine) and nitrogen tetroxide (N₂O₄) oxidizer. Since this propellant combination is hypergolic, no catalyst or igniter in the engine thrust chamber is required. In total, the second stage will burn approximately 6,000 kg of propellant at an average thrust of 43,370 N. A set of hydraulically activated engine gimbals will provide pitch and yaw control during powered flight, and a nitrogen cold gas jet system will provide the roll authority. In addition, the nitrogen jets will also provide attitude control for the coast phases.

3.1.4. Third Stage and Payload Fairing

A spin-stabilized third stage will provide most of the velocity required to boost the MGS spacecraft from low Earth orbit to a trans-Mars trajectory. This stage consists primarily of a Thiokol Star-48B solid motor. The engine on the third stage will provide an average thrust of 66,370 N and will burn about 2,000 kilograms of ammonium perchlorate propellant.

During launch and ascent through the lower atmosphere, a 2.9 meter diameter payload fairing will protect the MGS spacecraft and Delta third stage from aerodynamic forces. The fairing will be jettisoned from the launch vehicle at an altitude of approximately 129 kilometers, shortly after second stage ignition.

3.2. Pre-Launch Activity Overview

Delivery of the spacecraft to Cape Canaveral is currently scheduled for August 1996, after completion of system testing Lockheed-Martin's Waterton facility near Denver, Colorado. Arrival inspection and checkout will take place at the KSC Payload Hazardous Servicing Facility (PHSF), where a DSN end-to-end compatibility test via MIL-71 will be conducted. Technicians at KSC will then load the spacecraft with propellants and mate it to the Delta third stage. Following mating at the PHSF, initial interface verification testing will be performed.

Stacking of the Delta first and second stage will occur at Pad-17A in parallel with the activities at the PHSF. Roughly three weeks before launch, the coupled spacecraft-PAM combination will be moved to Pad-17A for mating with the bottom two stages of the Delta booster. Following final interface testing and close-out activities at the launch pad, the payload fairing will be installed, and Delta launch readiness will be verified.

3.3. Launch Strategy

3.3.1. Launch Period

The current mission baseline calls for an opening of the launch period on 4 November 1996 and closing 22 days later on 25 November 1996. This time period is bound on both ends by trans-Mars trajectory C₃ values too high for the Delta to achieve, given the mass of the spacecraft. The minimum C₃ corresponds to a launch day of 18 November 1996, about three-fourths of the way into the launch period.

During the duration of the launch period, the declination of the departure asymptote (DLA) varies from a minimum of 20.1° at the open, and 36.5° at the close. Geometrical constraints of the interplanetary injection problem dictate that the parking orbit prior to trans-Mars injection must lie at an inclination greater than or equal to the DLA. In order to satisfy this constraint and minimize the number of parking orbits in the launch vehicle targeting specification, the Delta will fly at only two different launch azimuths.

Use of two azimuths for launch results in a split-period launch strategy. For the first half of the launch period, called the early launch period, the Delta will launch at a 95° azimuth to reach a 28.7° park orbit. During the second half, called the late launch period, the Delta will launch at a 110° launch azimuth to reach a 36.5° park orbit. Although a booster performance penalty normally exists for reaching higher inclination orbits, they are in part offset by the lower C₃ requirements during the second half of the launch period. The dates for the early and late launch periods are 4 November to 15 November, and 16 November to 24 November, respectively.

In reality, rockets launched from the Cape must fly at an azimuth several more degrees south of East than 110° in order to reach an inclination of 36.5° directly. However, such a strategy is not possible due to range safety constraints. Instead, the Delta second stage will perform a "dog-leg" maneuver to reach the proper inclination.

3.3.2. PCS and Contingency Launch Period

Probability of commanded shut-down (PCS) for the Delta 2 second stage plays a key role in determining the duration of the MGS launch period. The reason is that the second burn of the second stage provides part of the energy required to place the spacecraft on the trans-Mars trajectory. PCS defines the probability that the second stage engine will complete its burn before the propellant supply depletes. By accepting a lower PCS value, it is possible to achieve a higher injected mass. Current project policy dictates a minimum PCS value of 95% in order to achieve the necessary C₃ for a given launch day.

The contingency launch period is defined to be the period of time after the end of the primary launch period which provides injection opportunities to Mars at the expense of certain launch vehicle and spacecraft performance parameters. A contingency launch period strategy using a lower than 95% PCS value, or 95% PCS but less total mission Delta-V capability has not yet been developed.

3.3.3. Launch Window

Delta 2 rockets do not possess a daily, variable launch azimuth capability, nor does the Delta first stage possess the capability to yaw-steer during ascent. Consequently, traditional launch windows (typically two hours in duration for the Mars Observer launch) are not possible. Instead, MGS must launch within a very small time period (about 1 second). These extremely short windows are referred to as "instantaneous launch windows."

3.3.4. Launch Opportunity

For interplanetary injections, two opportunities exist every day for a rocket to launch and inject its payload onto the proper Earth escape trajectory. The primary difference between the two, called the long and short coasts, is the length of time that the Delta must wait in its low Earth parking orbit before reaching the proper location to perform the trans-Mars injection burn. On any given day (relative to the time when the launch site on Earth rotates through the departure asymptote), the long coast launch opportunity always occurs first.

During the wait in low Earth orbit for the trans-Mars injection burn, the MGS spacecraft will rely on its batteries for power because the solar panels will have not yet been deployed. Preliminary analysis shows that on the long coast, the spacecraft will need to rely on battery power for up to 91 minutes. However, battery depth of discharge does not represent the limiting factor in the choice between the two launch opportunities.

The major constraint involves sun avoidance. An MGS flight rule specifies that science instruments (located on the +Z axis of the spacecraft) must always remain pointed at least 30° away from the Sun. On this mission, the long coast requires a dawn launch from the Cape. Because launch occurs generally in the eastward direction, the science instruments will be pointed forward in the booster's direction of flight, and the payload fairing will be jettisoned when the Delta has pitched over to a horizontal orientation, the science instruments will be pointed almost directly at the Sun at the time of payload fairing jettison. For this reason, the MGS mission will utilize the short coast launch opportunity as baseline.

3.4. Boost Profile and Injection

In general, the exact mission elapse times for key events depend on the orientation and location of the Earth departure asymptote (variable with each launch day), and the launch azimuth of the booster (either 95° or 110^o). However, in all cases, the event times for the first stage boost profile always remain constant.

During the launch phase, the booster will not provide the spacecraft with any power or telemetry capabilities. The spacecraft will launch with power for the computer, receiver, and attitude control sensors supplied from the batteries. Switch-over from launch-pad power to internal spacecraft power will occur at T- 4 minutes prior to launch.

3.4.1. Stage One

Lift-off will occur from SLC-17A at Cape Canaveral Air Force Station. This time will vary from as late as 13:40 to as early as 10:01 EST. At the time of lift-off, the main engines and six of the nine solid rocket motors will ignite. Approximately 63 seconds into flight, the solids will burn-out and be jettisoned. Then, the remaining three solids will ignite and burn for another 63 seconds before being ejected. Main engine cut-off will occur 260.7 seconds after lift-off a sub-orbital altitude of 116.51 km, 540.5 km down range from the launch site. The vernier engines will continue to burn for another six seconds, and first stage jettison will occur two seconds later at L+ 268.7 seconds.

3.4.2. Stage Two

At L+ 274.2 seconds, about 5.5 seconds after stage one jettison, stage two will ignite. About eight to 15 seconds after second stage ignition, the free molecular heating rate will have dropped to below 1135.0 W/m², allowing the Delta to jettison its payload fairing. In total, stage two will thrust for approximately five minutes to boost the spacecraft from its sub-orbit state at stage one jettison to a circular, low Earth parking orbit at an altitude of 185 km. Second stage cut-off (SECO1) will occur at L+ 580.8 seconds and an inclination of 28.7^o, or at L+ 575 seconds and an inclination of 36.5° for the early and late launch periods, respectively.

3.4.3. Trans-Mars Injection

After parking orbit insertion, the booster and spacecraft will coast for roughly 40 minutes to an hour until they reach the proper position to begin the two-burn trans-Mars injection sequence. First, stage two will re-start and thrust for roughly 120 seconds to raise the apogee of the parking orbit. After cut-off of the second stage engine (SECO2), the Delta will coast for 53 seconds before jettisoning the second stage. During that time, small rockets on the spin table (attached to the bottom of stage three) will fire to spin stage three and the MGS spacecraft to a rate of 60 r.p.m. for spin stabilization purposes.

Once 37 seconds have elapsed after stage two jettison (90 seconds after SECO2), stage three will ignite and burn for 87.14 seconds to complete the trans-Mars injection sequence. At the end of the burn, the MGS spacecraft will be on an Earth escape trajectory. For the baseline short coast option, trans-Mars injection almost always occurs in darkness, somewhere over the Indian ocean.

The combination of the second burn of the second stage and the third stage burn will provide the Delta-V needed for trans-Mars injection. During every day of the launch period, the third stage will impart the same amount of Delta-V to the spacecraft. The burn time and Delta-V of the second stage's second burn will vary depending on the specific C₃ requirements of the given launch day.

3.4.4. De-Spin and Spacecraft Separation

At the time of stage three burn-out, the Delta and MGS spacecraft will still be spinning at 60 r.p.m. This rotation must be nullified because the MGS spacecraft functions on three-axis stabilization. In order to de-spin the spacecraft, a yo-yo cable device on the third stage will deploy approximately 282 seconds after burn-out (369 seconds after ignition). The device works by transferring the angular momentum of the third stage and spacecraft to the cable similar to how a spinning figure skater slows her spin by extending her arms.

One second after yo-yo release, pyrotechnic devices will fire to sever the connection between the spacecraft and third stage. A set of four springs will then uncoil to impart a relative separation velocity of between 0.6 and 2.4 m/s between the third stage and spacecraft. The 283 second wait after burn-out for separation is designed to allow adequate time for residual thrust from the third stage to tail-off and ensure that the stage will not collide with the spacecraft after separation. During this waiting period, the spacecraft a set of thermal blankets located on the third stage will protect the spacecraft from thermal soakback.

The actual mission elapse time of separation depends on the length of time that the spacecraft spends in the low Earth park orbit before trans-Mars injection and will vary with each launch day. However, separation will always occur 370 seconds after third stage ignition. The choice of 370 seconds is in part driven by the standard cascaded event timers that McDonnell Douglas installs on the third stage.

3.4.5. Sequence of Major Launch Events

Event                         Mission Elapsed Time (11/3    Mission Elapsed Time (11/16   
                              to 11/15)                     to 11/24)                     

Lift-Off                      0 .000 seconds                0.000 seconds                 
Mach 1                        32.258                        32.258                        
Maximum Dynamic Pressure      49.458                        49.458                        
Solid Rocket Motor Jettison   67.000                        67.000                        
(6 of 9)                                                                                  
Solid Rocket Motor Jettison   131.500                       131.500                       
(3 of 9)                                                                                  
Stage 1 Main Engine Cut-Off   260.664                       260.664                       
Stage 1 Jettison              268.664                       268.664                       
Stage 2 Ignition              274.164                       274.164                       
Jettison Payload Fairing      282.000                       289.000                       
Stage 2 First Cut-Off         575.010                       580.832                       
Stage 2 Restart               Stage 2 cut-off, minus ~120   Stage 2 cut-off, minus ~120   
                              seconds                       seconds                       
Stage 2 Second Cut-Off        Stage 3 ignition, minus       Stage 3 ignition, minus       
                              90.000 seconds                90.000 seconds                
Stage 2 Jettison              Stage 3 ignition, minus       Stage 3 ignition, minus       
                              37.000 seconds                37.000 seconds                
Stage 3  Ignition             Stage 3 burn-out, minus       Stage 3 burn-out, minus       
                              87.14 seconds                 87.14 seconds                 
Stage 3 Burn-Out              44 to 50 minutes after        39 to 47 minutes after        
                              lift-off                      lift-off                      
Yo-yo deploy and De-spin      369 seconds after Stage 3     369 seconds after Stage 3     
                              ignition                      ignition                      
Spacecraft Separation from    370 seconds after Stage 3     370 seconds after Stage 3     
Stage 3                       ignition                      ignition

3.4.6. Booster Tracking

NASA-HQ requires monitoring of all launch vehicle powered flight and separation events. Because the injection of the MGS spacecraft will occur over either the Indian Ocean where ground station coverage is unavailable, the MGS mission will require use of the Airborne Range Instrumentation Aircraft (ARIA) to monitor and record telemetry from the Delta third stage. These aircraft must be available and on station in order for the MGS mission to launch. The Deep Space Network (DSN) does not track the launch vehicle.

3.5. Launch Targets

Launch vehicle targets represent the state that the Delta must deliver to the spacecraft in order to place MGS on the proper trans-Mars trajectory. These states are defined as osculating C₃, DLA, and RLA achieved at the target interface point, defined as 10 minutes after third stage ignition in agreement with McDonnell Douglas. The departure states have been biased to satisfy planetary quarantine requirements.

C₃	Departure energy or hyperbolic excess velocity squared
	(km²/s²)

DLA	Declination of the departure asymptote vector (degrees, EME2000)

RLA	Right ascension of the departure asymptote vector (degrees, EME2000)

TLO	Time of lift-off (hh:mm:ss, UTC)

TIP	Target interface point (hh:mm:ss, UTC)

The next table lists the launch vehicle target states for the first half, or early launch period. For these dates, the Delta 2 will fly at a 95.0° flight azimuth to a 28.7° inclination low Earth parking orbit.

Launch Date   Departure C₃  Departure     Departure     TLO           TIP           
                            DLA           RLA                                       
4 Nov 1996    10.6724       20.1043       173.8375      17:52:58.8    18:45:18.4    
5 Nov 1996    10.4107       20.6853       173.5075      17:41:11.1    18:33:49.9    
6 Nov 1996    10.1883       21.1564       173.3397      17:31:06.1    18:24:00.7    
7 Nov 1996    10.0009       21.6948       173.3564      17:20:45.8    18:14:00.2    
8 Nov 1996    9.8095        22.3253       173.2212      17:08:23.3    18:02:02.5    
9 Nov 1996    9.6470        22.9636       173.1850      16:55:54.2    17:50:00.2    
10 Nov 1996   9.4846        23.6794       172.7976      16:41:08.8    17:35:46.9    
11 Nov 1996   9.3517        24.3706       172.8790      16:26:26.7    17:21:38.4    
12 Nov 1996   9.2235        25.1477       172.6105      16:08:46.6    17:04:39.4    
13 Nov 1996   9.1151        25.9545       172.3173      15:48:51.1    16:45:32.3    
14 Nov 1996   9.0339        26.7008       172.1430      15:28:26.6    16:25:59.2    
15 Nov 1996   8.9643        27.5515       171.8147      15:00:58.5    15:59:43.4

The following table lists the launch vehicle target states for the second half, or late launch period. For these dates, the Delta 2 will fly at a 110.0° flight azimuth to a 36.5° inclination low Earth parking orbit.

Launch Date   Departure C₃  Departure     Departure     TLO           TIP           
                            DLA           RLA                                       
16 Nov 1996   8.9050        28.4748       171.5981      18:40:24.9    19:27:40.8    
17 Nov 1996   8.8778        29.3827       171.2561      18:25:29.7    19:13:18.7    
18 Nov 1996   8.8683        30.1808       171.0522      18:11:43.5    19:00:03.3    
19 Nov 1996   8.8737        30.9830       170.8524      17:57:16.7    18:46:09.9    
20 Nov 1996   8.9014        31.9596       170.5218      17:39:06.3    18:28:42.7    
21 Nov 1996   8.9317        32.7891       170.3349      17:12:13.2    18:12:30.2    
22 Nov 1996   8.9859        33.6270       170.1564      17:03:39.7    17:54:42.2    
23 Nov 1996   9.0679        34.6905       169.8594      16:37:17.5    17:29:28.2    
24 Nov 1996   9.1465        35.5644       169.6988      16:09:27.2    17:02:50.7    
25 Nov 1996   9.2366        36.4385       169.5452      15:21:29.5    16:17:16.1

The times and target states listed in these two tables are preliminary values pending further refinement with McDonnell Douglas.